Abstract: A method for the integration of continuous J shaped spars in extended panels of composite material in which J shaped spars are each arranged on a rectilinear support having at least one right triangle shaped portion, and the spars carried by the respective rectilinear supports are subsequently arranged on a stratified structure carried by a mould. A vacuum bag is created which covers the stratified structure, the rectilinear supports, and the J shaped spars. Holes are created in the vacuum bag at respective connection points made along each rectilinear support. A supporting structure provided with hooking and adjusting elements with the connection points adjusts the position of each hooking and adjusting element with respect to the supporting structure in order to position the rectilinear support in a direction parallel to the vertical so that the position of the end portion of the J shaped spar, with respect to the mould, corresponds to that foreseen by the design.

Abstract: A method for manufacturing a panel for an aircraft is provided. The method includes arranging, on a curing surface, a skin of pre-preg composite material, arranging a plurality of pre-formed, uncured composite stringers, each having an open cross section with a web and a base transverse to each other so that a first base portion protrudes toward a first side and a second base portion protrudes toward a second side, arranging the stringers on the skin with the base in contact with the skin, arranging, on each stringer, a respective first mandrel to completely cover the first base portion and a respective second mandrel to completely cover the second base portion, and having the skin and the stringers undergo a co-curing process in autoclave with vacuum bag.

Abstract: A multilayer radar-absorbing laminate includes three juxtaposed blocks. A first electrically conductive block is arranged toward the inside of the aircraft in use. A second electromagnetic intermediate absorber block has a layer of electrically non-conductive fiber sheets is permeated by graphene-based nanoplatelets to achieve a periodic and electromagnetically subresonant layer, the conductive layers containing graphene nanoplatelets alternating with non-conductive layers. A third block of electrically non-conductive material is arranged towards the outside and forms part of the outer surface of the aircraft. The second block is produced by depositing on the fiber sheets a suspension of graphene nanoplatelets in a polymeric mixture, with controlled penetration of the graphene nanoplatelets into the fiber sheets. A plurality of dry fiber sheets sprayed with the suspension of graphene nanoplatelets is superimposed.

Abstract: A method for manufacturing a wing box for aircraft comprises the steps of arranging, on a curing surface, a first panel of composite material, alternately arranging, on the first panel, along a transverse direction, a rib of non-polymerized composite material and a tool comprising a central part, a bottom part and a top part, wherein the central part of each tool is interposed between said bottom part and the top part and may be extracted in a transverse direction, arranging a second panel of composite material by putting said second panel in contact with the flanges of each rib, pulling out the central part of each tool along the transverse direction and removing the top part and the bottom part of each tool, and subjecting the first panel, the second panel, and each rib to a curing process in autoclave with vacuum bag.

Abstract: A method for manufacturing a wing-box for aircraft comprises: arranging, on a curing surface, a first panel of composite material comprising a skin a plurality of longitudinal stringers, and arranging, on each stringer, a caul plate; arranging, on the first panel, pluralities of support inserts so that each support insert rests on a respective caul plate, and on the skin of the first panel, and arranging, on the first panel, a plurality of ribs of non-polymerized composite material, each rib comprising a plate, a first pair of flanges and a second pair of flanges arranged at opposed ends of the plate, and having openings on an edge of the plate, by placing the respective first pair of flanges on the first panel and the respective plurality of openings on the plurality of support inserts, and having the first panel and the plurality of ribs undergo a curing process in autoclave with vacuum bag.

Abstract: A multilayer radar-absorbing laminate includes three juxtaposed blocks. A first electrically conductive block is arranged toward the inside of the aircraft in use. A second electromagnetic intermediate absorber block has a layer of electrically non-conductive fiber sheets is permeated by graphene-based nanoplatelets to achieve a periodic and electromagnetically subresonant layer, the conductive layers containing graphene nanoplatelets alternating with non-conductive layers. A third block of electrically non-conductive material is arranged towards the outside and forms part of the outer surface of the aircraft. The second block is produced by depositing on the fiber sheets a suspension of graphene nanoplatelets in a polymeric mixture, with controlled penetration of the graphene nanoplatelets into the fiber sheets. A plurality of dry fiber sheets sprayed with the suspension of graphene nanoplatelets is superimposed.

Abstract: A method for manufacturing a wing-box for aircraft comprises: arranging, on a curing surface, a first panel of composite material comprising a skin a plurality of longitudinal stringers, and arranging, on each stringer, a caul plate; arranging, on the first panel, pluralities of support inserts so that each support insert rests on a respective caul plate, and on the skin of the first panel, and arranging, on the first panel, a plurality of ribs of non-polymerized composite material, each rib comprising a plate, a first pair of flanges and a second pair of flanges arranged at opposed ends of the plate, and having openings on an edge of the plate, by placing the respective first pair of flanges on the first panel and the respective plurality of openings on the plurality of support inserts, and having the first panel and the plurality of ribs undergo a curing process in autoclave with vacuum bag.

Abstract: A method for manufacturing a wing box for aircraft comprises the steps of arranging, on a curing surface, a first panel of composite material, alternately arranging, on the first panel, along a transverse direction, a rib of non-polymerized composite material and a tool comprising a central part, a bottom part and a top part, wherein the central part of each tool is interposed between said bottom part and the top part and may be extracted in a transverse direction, arranging a second panel of composite material by putting said second panel in contact with the flanges of each rib, pulling out the central part of each tool along the transverse direction and removing the top part and the bottom part of each tool, and subjecting the first panel, the second panel, and each rib to a curing process in autoclave with vacuum bag.

Abstract: Methods for manufacturing elongated structural elements are provided. Such methods provide composite material having optimal properties such as weight and strength which can be produced at much lower costs compared to conventional methods. Composite materials including such elements are also provided. In addition, commercial products incorporating such structural elements are provided.

Abstract: Methods for manufacturing elongated structural elements are provided. Such methods provide composite material having optimal properties such as weight and strength which can be produced at much lower costs compared to conventional methods. Composite materials including such elements are also provided. In addition, commercial products incorporating such structural elements are provided.

Abstract: Methods of making single piece multi-spar boxes made of composite material for aircraft are provided. Such methods include providing outer spars of the boxes which are made with a C-shaped cross-section such that the spars, during curing, are molded on the inside onto rigid, outer, lateral forming tools. As a result it is possible to obtain very precise outer profiles and positions of the spars.

Abstract: Method for manufacturing stiffened panels comprising the following steps: forming a first and a second wet skin, each comprising a plurality of plies of preimpregnated composite material; forming a plurality of pairs of wet stringers, comprising a plurality of plies of preimpregnated material, wherein forming each pair of wet stringers comprises a) forming an assembly having a central web extending between opposite end flanges, and b) cutting the assembly along a median plane so as to form a pair of wet stringers, placing the pairs of wet stringers on the first wet skin by means of removable supporting members; turning the second wet skin over and placing it on the pairs of stringers, so as to obtain a pair of wet panels arranged above each other; and subjecting the pair of wet panels to a same curing cycle.

Abstract: Method for manufacturing stiffened panels, comprising the following steps: forming a first and a second wet skin, each comprising a plurality of plies of preimpregnated composite material; forming a plurality of pairs of wet stringers, comprising a plurality of plies of preimpregnated material, wherein forming each pair of wet stringers comprises a) forming an assembly having a central web extending between opposite end flanges, and b) cutting the assembly along a median plane so as to form a pair of wet stringers, placing the pairs of wet stringers on the first wet skin by means of removable supporting members; turning the second wet skin over and placing it on the pairs of stringers, so as to obtain a pair of wet panels arranged above each other; and subjecting the pair of wet panels to a same curing cycle.

Abstract: A tool for forming a C-section beam made of carbon fiber includes a straight elongated mandrel with a rectangular or trapezoidal cross-section. The mandrel has three consecutive outer faces, with a planar face configured to form a surface of the web of the C-section beam, and two opposing lateral faces, configured to form two respective surfaces of the flanges of the C-section beam. The mandrel has a distributed perforation on at least one of the three outer faces and an inlet/outlet connection to connect the mandrel to an external source of vacuum and/or compressed air. Passages are formed inside the mandrel to establish a fluid communication between the distributed perforation and the inlet/outlet connection. Through the distributed perforation, a vacuum is applied to the interface between the mandrel and the C-section beam to retain the C-section beam on the mandrel while lifted and turned by 180° for coupling to a similar C-section beam to assemble an H-beam.

Abstract: Methods of making single piece multi-spar boxes made of composite material for aircraft are provided. Such methods include providing outer spars of the boxes which are made with a C-shaped cross-section such that the spars, during curing, are moulded on the inside onto rigid, outer, lateral forming tools. As a result it is possible to obtain very precise outer profiles and positions of the spars.

Abstract: Methods for manufacturing elongated structural elements are provided. Such methods provide composite material having optimal properties such as weight and strength which can be produced at much lower costs compared to conventional methods. Composite materials including such elements are also provided. In addition, commercial products incorporating such structural elements are provided.

Abstract: A straight threaded rod has three reference elements spaced apart from each other. A shoulder and a knob having a fourth reference element are screwed onto the rod. A spring is mounted on the rod between the knob and the shoulder. The rod is inserted through two axially aligned holes in two abutting pieces, with the shoulder bearing on a first surface of one of the two pieces. By adjusting the position of the shoulder and the knob along the rod it is possible to apply a calibrated clamping force to a pair of abutting pieces, regardless of the thickness of the pieces.

Abstract: A tool for forming a C-section beam made of carbon fibre includes a straight elongated mandrel with a rectangular or trapezoidal cross-section. The mandrel has three consecutive outer faces, with a planar face configured to form a surface of the web of the C-section beam, and two opposing lateral faces, configured to form two respective surfaces of the flanges of the C-section beam. The mandrel has a distributed perforation on at least one of the three outer faces and an inlet/outlet connection to connect the mandrel to an external source of vacuum and/or compressed air. Passages are formed inside the mandrel to establish a fluid communication between the distributed perforation and the inlet/outlet connection. Through the distributed perforation, a vacuum is applied to the interface between the mandrel and the C-section beam to retain the C-section beam on the mandrel while lifted and turned by 180° for coupling to a similar C-section beam to assemble an H-beam.

Abstract: A beam of fiber-reinforced curable thermosetting composite material is preliminarily assembled in an uncured condition. The beam is placed between two forming tools, namely a lower and an upper tool, shaped in such a way as to give the lower and upper flanges a predetermined shape, at least one of the flanges having a non-rectilinear profile viewed in a longitudinal vertical plane. Two longitudinally elongate rigid inserts are placed between the flanges at the sides of the web and are covered in respective airtight tubular bags. The spaces between the upper tool and the lower tool on the two opposite longitudinal sides are sealed. The open opposite ends of the tubular bags are also sealed in such a manner that the pressure applied in an autoclave during the step of curing causes the tubular bags to swell and presses them against the two opposite faces of the web and against the flanges of the beam. On completion of curing, the upper forming tool is removed and the inserts are extracted sideways.

Abstract: A method of manufacturing a curved structural element made of composite material has a complex, open cross-section the structural element is formed by the assembly of components made of composite material. A plurality of flat, fresh components are prepared, each component being constituted by a flat layer of curved fibres previously infused with resin, combined with a pre-impregnated layer with straight fibres. Hot-forming of the flat, raw components is then performed. The hot-formed, fresh components are then assembled in accordance with a predetermined sequence on a polymerization tool. The tool includes a shaped, inflatable device astride which a C-shaped raw component is deposited. Finally, the fresh components assembled on the tool are subjected to a polymerization cycle in an autoclave. The inflatable device is inflated by the pressure produced in the autoclave to apply a predetermined pressure from the inside of the C-shaped fresh component.