Abstract: A metalization structure of aircraft panels and process for obtaining such structure including a panel made in composite material, the panel including a metal mesh or foil in its outer part, said panel further including a bore which in turn comprises a shank and an upper countersinking, such that the mentioned panel is fixed to another structural aircraft element by means of metal fixing elements. The structure includes a metallization layer arranged on the upper countersinking, extending in the adjacent area of the metal mesh or foil, such that it is ensured that electrical continuity exists between the upper countersinking and the metal mesh or foil during the time that the lightning bolt strikes the structure of the panel, further ensuring that most of the current discharged by the lightning bolt is conducted over the surface of the structure of the panel.

Abstract: The present invention relates to a leading edge (1) for aircraft made of composite material, characterized in that it comprises on its inner face a metallic-type reinforcement (2) firmly adhered to the mentioned inner face of the leading edge (1) arranged such that it confers the leading edge (1) with better capacity to adapt to deformability and greater absorption of energy due to impacts. The invention also relates to a method of manufacturing a leading edge (1) with a metallic reinforcement (2) for aircraft made of composite material.

Abstract: The invention relates to a curing tool (11) for curing a composite material part formed by bundles of fibers embedded in a resin matrix in an installation, such as an autoclave, allowing the curing of the part by means of applying high pressures and/or temperatures, comprising at least one mechanical vibrating actuator (1) acting on the curing tool (11) itself to facilitate resin distribution in the part during the curing process.

Abstract: 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.

Abstract: The invention relates to a metalization structure of aircraft panels and process for obtaining such structure, such that it comprise a structure of a panel (1) made in composite material, the mentioned panel (1) comprising a metal mesh or foil (2) in its outer part, said panel (1) further comprising a bore which in turn comprises a shank and an upper countersinking (8), such that the mentioned panel (1) is fixed to another structural aircraft element (3) by means of metal fixing elements (4), characterized in that said structure comprise a metallization layer (7) arranged on the upper countersinking (8), extending in the adjacent area of the metal mesh or foil (2), such that it is ensured that electrical continuity exists between the upper countersinking (8) and the metal mesh or foil (2) during the time that the lightning bolt strikes the structure of the panel (1), further ensuring that most of the current discharged by the lightning bolt is conducted over the surface of the structure of the mentioned panel

Abstract: The method comprises: a first stage in which the layers of the prepreg are spread over a mold having cavities corresponding to reliefs (3) of the complex geometry panel (1) to be obtained, the layers of prepreg presenting lines of discontinuity (12) in the vecinity of the reliefs (3); a second stage in which the stack (11) is done by application of a cycle of pressure and temperature; and a third stage of finishing of the panel to be obtained, comprising the curing of the prepreg.

Abstract: The method basically consists of mounting the spars (2, 2?) and the ribs (3) with reference to the skin (4) in such a manner that, once the spars (2, 2?) are positioned with their proper reference in an assembly jig (12), the ribs (3) are transversely positioned with the condition that they are adjusted making contact against the inner surface of the skin (4). The device comprises a template element (15) which is used for previously providing the skin with the theoretical form it must acquire once mounted, and to firmly maintain that theoretical form by including means for securing and tightening (16) the skin (4), as well as means (17, 18, 19, 20) for supporting and positioning the skin (4).

Abstract: Structural component of non-conductive composite material, particularly for aircraft (panel (1), frame (3), stringer (2), etc), that comprises in the surface a metal-type layer (4), such that the mentioned structural component together with the remaining metallized structural components can provide the aircraft with the sufficient conductive metal mass. According to a second aspect of the invention, a metallization method is proposed in the manufacture of a structural component, particularly for aircraft, made of composite material, which method allows the geometric configuration of the add-on by applying it in an extensive or limited manner to predetermined contours by means of using templates. This constructive feature allows creating integrated electric circuits in the airplane structure by means of creating independent tracks with different widths and thicknesses.

Abstract: The present invention relates to a leading edge (1) for aircraft made of composite material, characterized in that it comprises on its inner face a metallic-type reinforcement (2) firmly adhered to the mentioned inner face of the leading edge (1) arranged such that it confers the leading edge (1) with better capacity to adapt to deformability and greater absorption of energy due to impacts. The invention also relates to a method of manufacturing a leading edge (1) with a metallic reinforcement (2) for aircraft made of composite material.

Abstract: 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.

Abstract: The invention relates to a curing tool (11) for curing a composite material part formed by bundles of fibers embedded in a resin matrix in an installation, such as an autoclave, allowing the curing of the part by means of applying high pressures and/or temperatures, comprising at least one mechanical vibrating actuator (1) acting on the curing tool (11) itself to facilitate resin distribution in the part during the curing process.