Abstract: The invention relates to a skin panel (11) of an aircraft (1), comprising a laminate of at least one first metal sheet (40) and preferably first fiber-reinforced polymer layers (41) connected thereto. The skin panel (11) is connected to stiffening elements (4) made of a laminate of second metal sheets (40) and second fiber-reinforced polymer layers (41) connected thereto, provided that the second fiber-reinforced polymer layer comprises fibers having a modulus of elasticity in tension of greater than 110 GPa. The invention also relates to an aircraft or spacecraft provided with such skin sheets.

Abstract: Methods of manufacturing structural panels for use in manufacturing aircraft fuselages and other structures are disclosed herein. In one embodiment, a structural panel configured in accordance with the invention includes a skin and at least first and second stiffeners. The first stiffener can have a first flange portion mated to the skin and a first raised portion projecting away from the skin. The second stiffener can have a second flange portion mated to the skin and a second raised portion projecting away from the skin. At least one of the first flange portion of the first stiffener and the second flange portion of the second stiffener can extend toward the other of the first flange portion and the second flange portion to form an at least approximately continuous support surface to which a frame can be attached.

Abstract: Quasi-isotropic chopped prepreg is used to make parts found in aerospace vehicles. Exemplary aerospace parts that are made using quasi-isotropic chopped prepreg include aircraft window frames, wing fairing supports, flange supports, frame gussets, rudder actuator brackets, shear ties, seat pedestals, cargo floor flange supports, storage bin fittings, antenna supports, torque tube pans, handle boxes, side guide fittings, wing box covers and intercostals.

Abstract: The present application is directed to systems for controlling the direction of thermal expansion of a composite part during thermal processing. The systems may comprise at least one guiding mechanism capable of being attached to a surface of a tool for fabricating composites and at least one moveable composite insert capable of being attached to a composite part. The moveable composite insert has an opening for receiving the at least one guiding mechanism. The opening is capable of controlling movement of the at least one moveable composite insert in a desired direction during thermal expansion of the composite part, when the guiding mechanism is received by the opening. Methods for controlling the direction of thermal expansion of a composite part during thermal processing are also disclosed.

Abstract: The shape of a laminated composite structure is optimized using finite element analysis. The optimized shape minimizes the weight of the structure while satisfying minimum strength and stiffness requirements.

Abstract: A multifunctional member with a first active member (30), which is adapted to contract if exposed to a temperature above a first transition temperature range. A second active member (40), which is adapted to contract if exposed to a temperature above a second transition temperature range. A core member (20), which is adaptive for deformation. The first and second active members (30, 40) are attached on opposite or different sides of the core member (20). A heat source operatively connected to the first and second active members (30, 40) to expose them to transition temperatures. The first active member (30) contracts while above the first transition temperature range causing the second active member (40) to expand, wherein the second active member (40) is below the second transition temperature range.

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 mould (2) designed in such a way that, once the geometry of the element (1) is known, the mould (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: An integrated composite structure with a graded coefficient of thermal expansion (CTE) is formed by selecting a plurality of layers of materials with a graded CTE and using joining approaches such as welding, brazing, or solid state bonding to produce a CTE-graded layered composite or near net shape. The integrated composite billet or near net shape is then processed to produce a first surface for attachment of a first structural member having a first CTE and to produce a second surface of for attachment of a second structural member having a second CTE.

Abstract: A process for incorporating light emitting materials into aircraft interiors is described. The method includes selecting at least one light emitting material, incorporating the selected at least one light emitting material into one or more of a laminate coating and an extruded component, and installing the one or more of a laminate coating and extruded component into an aircraft cabin configuration.

Abstract: An airframe includes a mass barrier layer mounted to a multitude of frame members and adjacent to each of a multitude of foam portions to provide a contiguous layer to seal the foam portions within the respective multitude of frame voids.

Abstract: Composite panels used in aircraft are formed by a sandwich including a lightweight honeycomb core reinforced by a solid plank in medial regions of the panel. The plank is bonded to the honeycomb and forms an integral part of the core where a strut is attached to the panel. The plank distributes point loads imposed on the panel by the strut.

Abstract: A method for monitoring the integrity of a permeable structure that is disposed in an environment containing a fluid at ambient pressure is provided. At least one cavity is formed in or on the permeable structure. A source of first fluid is provided at a first pressure greater than the ambient pressure. The cavity is coupled to the source through a high-fluid-flow impedance to establish a flow of the first fluid through the permeable structure via the cavity. A rate of flow of the first fluid through the permeable structure is allowed to stabilize to a steady-state rate. A change in the steady-state flow rate of the first fluid through the permeable structure is monitored.

Abstract: The invention relates to a covering (22) of a fuselage (1) for aircraft comprising propeller engines (3) comprising an opening (12) which is surrounded by a structural frame formed by the ribs (5) and the stringers (6) of the aircraft around the area of maximum acoustic impact of the propeller engines (3), said opening (12) further comprising a non-structural panel (7) joined to the previous structural frame by means of a rather non-rigid joint which considerably attenuates the vibrations through such joint, thus reducing the indirect acoustic noise and the indirect structural noise generated by the propeller engines (3) inside the passenger cabin of the aircraft.

Abstract: A method and apparatus for a composite stringer. A method is used for manufacturing a hollow composite stringer. Foam is formed with a mandrel installed into the foam. A composite material and the foam is laid up onto a tool in a form of a stringer. The composite material and the foam in the form of the stringer is cured to form a cured stringer. The mandrel from the foam is removed to form the hollow composite stringer.

Abstract: Leading edge structures for high performance aircraft and the direct manufacture thereof are disclosed. Direct manufacturing technology is used to deposit leading edge structures directly from a digital model to form near-net shape products. This technique permits a wider range of materials to be utilized, such as casting alloys capable of the highest temperature usage that are not available in billet or sheet form. Refractory metals or high temperature ceramics also may be used to form the entire lead edge structure, or just a replaceable tip of the leading edge structure.

Abstract: An exemplary morphable ceramic composite structure includes a flexible ceramic composite skin and a truss structure attached to the skin. The truss structure can morph shape of the skin from a first shape to a second shape that is different than the first shape. The flexible ceramic composite skin may include a single-layer of three-dimensional woven fabric fibers and a ceramic matrix composite. The truss structure may include at least one actuatable element or an actuator may move a portion of the truss structure from a first position to a second position. A cooling component may be disposed in thermal communication with the skin. The cooling component may include thermal insulation or a cooling system that circulates cooling fluid in thermal communication with the skin. The morphable ceramic composite structure may be incorporated into any of an air inlet, combustor, exhaust nozzle, or control surfaces of a hypersonic aircraft.

Abstract: A method and apparatus for attaching parts. A composite male fastening component is placed through at least a first part and a second part. The composite male fastener component preferably has a threaded portion. A composite female fastener component is positioned adjacent to and surrounding the threaded portion of the male fastener component. A portion of the composite female fastener component is caused to flow around and form into the threaded portion of the composite male fastener component. The portion of the composite female fastener component that flowed around and formed into the threaded portion of the composite male fastener component is re-solidified and re-consolidated such that the composite female fastener component is securely attached to the composite male fastener component thereby joining the mating parts.

Abstract: Laminate structures and methods for forming same are disclosed. In one embodiment, a laminate structure includes a metal-polymer composite lamina. The metal-polymer composite lamina has a first face and a second face spaced apart, and extends to a terminal edge. The lamina includes a ply of fiber-reinforced polymer that extends between the first face and the second face and has an interior edge. The interior edge defines at least one cutout. A ply of metal foil extends between the first face and the second face substantially from the interior edge filling the at least one cutout.

Abstract: Components made from solid fiber inorganic-organic-polymer matrix composites and methods of making the components and composites are provided. In one embodiment, and by way of example only, the method includes the steps of impregnating a fiber preform with a liquid resin comprising at least one compound selected from the group consisting of silsesquioxane and polysilazane, and heating the resin-impregnated preform at a temperature below the resin pyrolysis temperature to form the solid fiber inorganic-organic-polymer matrix composite. In other embodiments, a component, such as a mechanical or an aircraft component is provided that includes a solid fiber inorganic-organic-polymer matrix composite. The solid fiber inorganic-organic-polymer matrix composite may include a fiber preform impregnated with a resin comprising at least one silsesquioxane or a fiber preform impregnated with a resin comprising at least one polysilazane.

Abstract: Systems and methods for discontinuously applying an insulating primer to a carbon fiber reinforced plastic (CFRP) component are disclosed. In one embodiment, a method for mitigating electrical surface discharges from a CFRP component includes first applying an insulating primer to a metallic component. Next, an insulating primer is applied discontinuously to the CFRP component adjacent the metallic component. The discontinuous application of the insulating primer forms a primed portions and unprimed portions. The unprimed portions are configured to enable electrostatic dissipation.

Abstract: 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.

Abstract: A hybrid composite-metal component is provided. The component includes an elongate inner metal piece, an outer metal piece disposed about at least a portion of the inner metal piece, and composite material disposed between the inner metal piece and the outer metal piece. The component may further include at least one of a seal and at least one fastener joining the inner metal piece and outer metal piece. Both the inner metal piece and the outer metal piece may include at least one tapered end. The tapered ends of both the inner metal piece and the outer metal piece each may include a double taper.

Abstract: Hybrid composite panel systems and methods are disclosed. In one embodiment, an assembly includes a primary section, a matrix member engaged with the primary section, and a secondary section engaged with the matrix member opposite the primary section. The primary section includes a plurality of first composite layers reinforced with a first reinforcing material, and the secondary section includes a plurality of second composite layers reinforced with a second reinforcing material. The primary and secondary sections are configured to bear an operating load at least partially transversely to the first and second composite layers, and are asymetrically configured such that the primary section bears a majority of the applied operating load.

Abstract: 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.

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: An acoustic absorption blanket mounting system includes a blanket mounting strip (BMS) attached to a fastener strip mounted to an airframe member. The BMS is generally of a shape to removably capture a sound proofing blanket between opposed pair of wings. At least one wing of the pair of wings includes an acoustic gasket strip sewn to the BMS base material layer. Opposed blanket hook and loop strips correspond to wing hook and loop strips such that a soundproofing blanket is retained between the wings such that the acoustic gasket strip minimizes acoustic leakage heretofore experienced in soundproofing blanket attachments.

Abstract: Disclosed is a floor panel, in particular for covering floors of cargo holds of aircraft. The floor panel is formed by a plurality of milled profiles, which are friction stir welded in each case along their longitudinal edges, with preference continuously to form longitudinal (butt) seams. The milled profiles each form a beam “of equal stress”, with an upper flange and a lower flange which are spaced apart from one another along their extent. The milled profiles have in each case a double-T-shaped cross-sectional profile, the upper flanges and the lower flanges being respectively connected by webs. With preference, the milled profiles are milled out in one piece from solid material of a sufficiently tough aluminium alloy.

Abstract: Sealing system for the gap (2) existing between the fuselage (3) and the elevator (4) of the orientable horizontal stabiliser (5) of an aircraft which comprises a main body (1) formed by an upper surface (18) and a lower surface (19) and at least one longitudinal vertical element (22) arranged between the two surfaces (18,19) which grants rigidity to the main body (1), and a plurality of first elastic sealing profiles (9) between the first surface (11) of the main body (1) 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 of the gap (2) takes place and an aerodynamic continuity is produced between 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 horiz

Abstract: Device for coupling a coupling pivot for a trimmable horizontal stabilizer (1b) to the tail fuselage (1a) of an aircraft (1), with an inner plate (10), an outer plate (11) with first fastening holes (10a, 11a) and first passage holes (10c) in the inner plate respectively facing second fastening holes (10b, 11b) and second passage holes (11c) in the outer plate (11) for the passage of a coupling pivot (9) of the horizontal stabilizer (1b); and with an upper lug (12) and a lower lug (13) respectively located above and below a cut-out (2) for the horizontal stabilizer (1b). The lugs (12, 13) are integral with separate reinforced zones (14, 15) of the covering (1f) in respective neighboring areas of the cut-out (2), in such a way that the plates (10, 11) between them clasp the respective reinforced zones (14, 15) in which the lugs (12, 13) are provided.

Abstract: The present invention provides methods for manufacturing an article having a wetting-resistant surface. The method includes providing a substrate. The method further includes disposing a coating mixture on a surface of the substrate, wherein the coating mixture comprises a braze material and a texture-providing material. The method further includes heating the braze material to bond the texture-providing material to the surface of the substrate to form the article having the wetting-resistant surface.

Abstract: A coating composition and a method for coating metallic substrates, such as aircraft components exposed to elevated temperatures and/or oxidative conditions. The coating composition includes an aqueous mixture having about 2 vol % to about 50 vol % organosilane, about 0.3 to about 25 vol % metal alkoxide, and about 0. 1 to about 30 vol % organic complexing agent. The coating composition may be deposited on a metallic substrate and cured to form a sol-gel coating on the surface of the substrate that is adherent to the substrate, and oxidation and discoloration resistant.

Abstract: The present invention is a BLISK airfoil heat treating apparatus and method for heat treating the leading and/or trailing edge section(s) of a BLISK airfoil using the BLISK airfoil heat treating apparatus. The apparatus comprises a pair of hingedly connected heat treating shells, each shell having a cavity for receiving an airfoil edge section requiring heat treatment. A resistive heating element is positioned with the shells to heat the cavities.

Abstract: The present invention is directed toward a laminate composite structure and a method for fabricating a laminate composite structure, wherein the laminate composite structure comprises a solid laminate and a composite sandwich structure. The solid laminate includes a bend having composite pins inserted there through. The composite sandwich structure is comprised of a core material sandwiched between top and bottom composite face sheets.

Abstract: A body part including a surface having a fluid flow thereover. The body part also includes flocking coupled to at least a portion of the surface that adjusts an aerodynamic characteristic relative to the surface that is devoid of the flocking.

Abstract: Provided is a composite joint comprising a pair of composite panels each defining a composite panel thickness and having an inner face sheet and an outer face sheet sandwiching a core panel therebetween. The core panel defines a constant thickness portion and a ramp portion along which the core panel tapers to a reduced thickness portion. The outer face sheet is thickened in the area of the reduced thickness portion. The core panel includes a plurality of cells extending between the inner and outer face sheets. The cells within the reduced thickness portions are filled with a filler material. The cells within the ramp portions and constant thickness portions are filled with aerogel. A primary splice plate extends across the reduced thickness portions. Mechanical fasteners extend through the outer face sheet and into the primary and secondary splice plates. A themal barrier compound covers the primary splice plate.

Abstract: A helicopter with one or more components formed from a self extinguishing composite primary structure material to provide increased passenger safety. The self extinguishing composite primary structure material has a core formed from a fire retardant material, which core has a first surface and a second surface, and at least one ply of a structural graphite prepreg material bonded to each of the first surface and the second surface.

Abstract: An aircraft includes a partition designed to separate a cargo part from a cockpit or a passenger compartment. The partition includes a corrugated structure, defining several corrugations with an approximately omega section extending along the direction of the height of the partition. The partition also includes one or several stiffener elements.

Abstract: A metallic layer material includes at least two metal sheets. Each sheet has a thickness of less than 1.5 mm. Between the sheets, a plastic layer is disposed, which is bonded to the metal sheets. The plastic layer comprises a plastic material and basalt fibers, and the basalt fibers have a modulus of elasticity exceeding 50 GPa. The basalt fibers may constitute 35-75% of the volume of the plastic layer. and may be comprised of 35-55 weight % of SiO2, 10-25 weight % of Al2O3, and 3-10 weight % of MgO. In aerospace applications and other applications requiring good specific strength, stiffness and damage tolerance, the metal sheets may be made of a metal having a tensile strength exceeding 0.20 GPa.

Abstract: A reinforced structural member having a tubular member and a support structure. The tubular member has a sidewall that defines a hollow interior portion. The support structure is formed from a composite material and bonded to the sidewall in at least two predetermined locations. The support structure is positioned within the tubular member and configured to transmit force between the predetermined locations so as to prevent the tubular member from being crushed in response to the application of a force of a predetermined magnitude to the tubular member. A method for forming a reinforced structural member is also provided.

Abstract: Laminate structures and methods for forming same are disclosed. In one embodiment, a laminate structure includes a metal-polymer composite lamina. The metal-polymer composite lamina has a first face and a second face spaced apart, and extends to a terminal edge. The lamina includes a ply of fiber-reinforced polymer that extends between the first face and the second face and has an interior edge. The interior edge defines at least one cutout. A ply of metal foil extends between the first face and the second face substantially from the interior edge filling the at least one cutout.

Abstract: An insulation package to reduce the transmission of heat and/or noise between the outer skin and the cabin lining of an aircraft composed of an insulating material enclosed by an encapsulating membrane. The encapsulating material has a vapor diffusion resistance (sd value) at least partially dependent on the ambient humidity, and for a gradient of relative humidity across the encapsulating membrane of 3% to 50% it has a vapor diffusion resistance (sd value) of at least 0.8 m and for a gradient of relative humidity across the encapsulating membrane of 50% to 93% it has a maximum vapor diffusion resistance (sd value) of 0.3 m.

Abstract: A Glare type composite laminated sandwich panel comprising (i) aluminum (ii) glass fibre with adhesive (iii) aluminum is disclosed. According to the present invention, at least one of the aluminum sheets is preferably made of an aluminum non-heat treatable alloy type Al—Mg with a magnesium content of between 4 and 6%. These sandwich panels have mechanical characteristics similar to sandwich panels comprising sheets with a heat treatable alloy. Panels of the present invention further generally have better corrosion resistance and are less expensive than those made with heat treatable alloys such as 2xxx and 7xxx type alloys.

Abstract: A composite material-stiffened panel has a skin obtained by molding a fiber-reinforced resin composite material into a flat skin, stiffeners arranged in rows on one surface of the skin, and a fiber-reinforced resin-composite material stitched on the skin covering the stiffeners. The composite material-stiffened panel can be manufactured as follows: A first fabric material is placed on a tool having a panel-shaped surface. Stiffeners are placed on the first fabric materials. A second fabric material is placed on the stiffeners to cover at least some of the stiffeners. The second fabric materials are reformed to match the shape of the stiffeners. The reformed second fabric material is stitched on the first fabric material along edges of the stiffeners. All the materials are covered with a bagging film for vacuum. Resin is infiltrated into the fabric materials by a RTM or a RFI method. The infiltrated resin is heated to be hardened.

Abstract: A structural panel comprised of a composite preform having an electrically conductive stitching. The stitching forms an electrically conductive grid-like network on the structural panel to better dissipate electrical energy received in a lightning strike throughout at least a portion of the thickness of the composite preform. Advantageously, the electrically conductive stitching is applied to the composite preform during the primary manufacturing process, rather than in a subsequent manufacturing step, to thereby help reduce the overall manufacturing cost of the structural panel. In one alternative embodiment a non-conductive polymer stitching is also applied throughout the entire thickness of the composite preform.

Abstract: Methods and compositions for fabricating composite parts including at least one structural material and at least one protective material that are integrally bonded without the use of secondary bonding operations. One or more of the materials forming the layers of the composite parts may be a ceramic composition with or without porosity and one or more of the materials may be a polymer composition. Methods including co-injection processes also are provided for fabricating multi-layered structures in which each layer serves a desired function while still being integrated into the overall structure.

Abstract: The cockpit door has a hinge on one side and a latch device on the other side that can be operated only from the cockpit side. To the inner side of the door panel is mounted a reinforcement member 100 having a body formed by laminating layers of aromatic polyamide fiber sheets and thermoplastic adhesive and subjecting the same to hot pressing to form an integral member. Mounting portions 150 and 160 are formed to the rim portion of the reinforcement member body 110, and rivet inserting holes 180 are machined to the portions. The mounting portions are bent by 90 degrees, and fixed to a frame using rivets.

Abstract: Composite tanks are desirable for aerospace applications since they are much lighter than comparable metal tanks. Various polymeric compositions satisfy the requirements to make a Lox compatible tank.

Abstract: A method and system for assembly and attachment of high temperature materials used as thermal protection systems (TPS) for applications such as re-entry or hypersonic vehicles are disclosed. The attachment system comprises a fastener that passes through an aperture of and a fitting attached to a ceramic matrix composite (CMC) TPS panel. The flexure fitting allows thermal growth in the panel while providing a structurally stiff interface to support the panel through structural and acoustic loading. The flexure fitting is detachably attached to the substructure, or fuel tank of a vehicle. The fastened connection is accessed through an aperture via an access cover plate conformal to the surface of the CMC TPS panel. It is by way of the access cover that the fastened connection is externally serviced.

Abstract: A preform is joined to converging surfaces of a structure, such as an airfoil, to form a structural edge. The preform has a body having an acute-angle edge at one end and connecting areas extending from the opposing end for connecting the preform to converging planar elements of a wing or other similar structure. In several embodiments, the acute angle of the preform generally matches the angle relative to converging ends of the planar elements, the preform forming the outer end of smooth, continuous outer surfaces after assembly. The connecting areas may be connected to interior or exterior surfaces of the elements, depending on the orientation of the preform and elements. In another embodiment, the entire preform is placed between the outer ends of the elements for aligning and supporting the elements at the desired location and angle relative to each other.

Abstract: An aircraft composite assembly 10 is provided, including a first composite member 12, at least one first flange joint 26 comprising a first plurality of plies 24 cured to the first composite member 12, a second composite member 30, and at least one second flange joint 38 comprising a second plurality of plies 36 cured to the second composite member 30. The at least one second flange joint 38 is bonded to the at least one first flange joint 26 such that the first composite member 12 is affixed to the second composite member 30.