Abstract: A laminate of metal layers and plastic bonding layers situated between the metal layers includes two external metal layers extending substantially continuously and at least one internal metal layer. At least one of the internal metal layers has at least one opening, and at the position of the opening the other metal layers and plastic bonding layers are bonded together to the form of a packet of lower thickness. The relatively thick parts of the laminate can be used for the transfer of forces.

Abstract: The present invention relates to an aircraft load frame (1) made of a composite material, characterized in that it comprises two side elements (2, 3) and a base element (7), the side elements (2, 3) being joined at the inner part of the frame (1) by means of the base element (7) and each of the side elements (2, 3) comprising a leg (4) joining the frame (1) to the aircraft fuselage skin, a web (5) and a lower flange (6) joining the web (5) and the base element (7).

Abstract: An aircraft forward pressure bulkhead and an aircraft crown panel are described herein. The forward pressure bulkhead includes a malleable and deformable dome that is configured to “catch” foreign objects, such as birds. The dome is intentionally designed to deform in response to a foreign object strike that imparts at least a threshold amount of impact energy to the bulkhead. The dome is free of rigid stiffeners and non-deformable reinforcement members that would otherwise hinder the flexible characteristic of the dome. Practical embodiments of the bulkhead utilize fewer parts, are less heavy, and are less expensive than traditional bulkheads that utilize rigid stiffeners. The crown panel has similar compliant characteristics. The crown panel is designed to simultaneously deform and deflect in response to foreign object impacts such as birds and hail. The crown panel is formed from a monolithic one-piece material, such as aluminum, and various reinforcing features are integrally formed in the material.

Abstract: A skid rail for aircrafts comprising (i) a base member comprising a long board fixed to a bottom surface of an aircraft fuselage, and bosses integrally projecting from an outer surface of the long board with longitudinal intervals, (ii) cores bonded to the long board between the bosses to form a rail-shaped, integral structure, and (iii) a skin covering the bosses and the cores, the base member and the skin being formed by a first fiber-reinforced plastic having fabrics of glass fibers and/or carbon fibers, and the cores being formed by a second fiber-reinforced plastic having fabrics of fibers having higher wear resistance than that of fibers in the first fiber-reinforced plastic.

Abstract: A surface seal for sealing a riveted and/or screwed joint between a first and a second component of an aircraft. The surface seal is arranged in the region of joining surfaces of the first and the second component. The surface seal is formed with a pressure-resistant plastic film that is provided with a pressure-sensitive adhesive layer on both sides. The surface seal 1 may furthermore comprise elements for detecting cracks. A method for producing a riveted and/or screwed joint between a first and a second component of an aircraft is also provided, with the joint being sealed with the surface seal.

Abstract: An aircraft leading edge panel having an outer aerodynamic surface; and an inner surface carrying the elastomeric impact protection membrane. The membrane comprises a woven or knitted fabric which is impregnated with an elastomeric material. The membrane and the panel may be bonded by an adhesive or co-cured to bond the membrane to the panel. The membrane provides impact protection to the panel by de-bonding from the face of the panel and absorbing at least part of the energy of an object impacting the panel.

Abstract: One embodiment relates to an article and a method for producing an article including a plurality of substrates, and an adhesive bonded between at least two of the plurality of substrates. The adhesive can include a polycarbonate copolymer that includes reacted resorcinol, siloxane, and bisphenol-A. Another embodiment relates to an article having a first polyimide substrate, a second polyimide substrate, and an adhesive bonded between the first substrate and the second substrate. The article can have a 2 minute integrated heat release rate of less than or equal to 65 kilowatt-minutes per square meter (kW-min/m2) and a peak heat release rate of less than 65 kilowatts per square meter (kW/m2) as measured using the method of FAR F25.4, in accordance with Federal Aviation Regulation FAR 25.853(d).

Abstract: An impact resistant composite sandwich structure includes least one laminate face sheet joined to a stiffening element. The stiffening element includes a plurality of fibers and a resin for binding the fibers. The resin has a strain-to-failure greater than approximately 6%.

Abstract: A composite structure comprising: a panel formed from two or more plies of composite material, the panel having a surface formed with a step where the thickness of the panel changes. Two or more stiffeners are attached to the surface of the panel. One of the stiffeners engages at least part of the step so as to transmit load forces in the panel.

Abstract: A part of composite material having a wedge between two zones, the second zone being shorter than the first zone, whose structure includes from its outer surface to its inner surface: a first section formed from at least two continuous sheets extending parallel to its outer surface, the gradient of the wedge being between 20% and 50%; a wedge in the shape of a triangular prism with its larger surface dimensioned in such a way that it forms a wedge having a gradient of less than 20%; a second section formed by a plurality of continuous sheets extending parallel to the surface bounded by the first section with the wedge placed upon it. The invention also relates to a process for manufacture of the part.

Abstract: A method of forming an exterior surface protective structure (12) for an aircraft (10) includes uniting a loaded surfacer (52) having a carrier (51) to a hybrid prepreg substrate (32). The prepreg substrate (32) includes a carbon fabric (44) with an integral conductive component (48) having conductivity with in a metal conductivity range and is united to a base substrate (30). The loaded surfacer (52) and the prepreg substrate (32) are cured, which includes interfacially adhering matter between the loaded surfacer (52) and the prepreg substrate (32). A protective fabric system (12) for an exterior (14) of an aircraft (10) includes the base substrate (30). The hybrid prepreg substrate (32) is coupled to the base substrate (30). The loaded surfacer (52) with the carrier (51) is interfacially adhered to the prepreg substrate (32).

Abstract: An integrated floor for an aircraft fuselage includes a composite panel forming a floor surface and composite beams bonded to the floor panel. The beams extend longitudinally within the fuselage and support the floor.

Abstract: Impact resistant cores and methods of manufacturing such impact resistant cores are provided. Additionally, components made using the impact resistant cores are provided. A particular impact resistant core includes a composite structure of at least one non-woven sheet including poly p-phenylene-2,6-benzobisoxazole fiber and a binding resin.

Abstract: A structural, energy absorbing composite element for aircraft structures includes a metallic or nonmetallic skin shell with a porous filler material insert positioned within the skin shell. The porous filler material is a material having ligaments that collapse resulting in a densification of the porous filler material in response to impact loading or a compression force sufficient to cause failure of the combined assembly.

Abstract: A monomolecular carbon-based film can be placed on an aircraft part, such as the leading edge designed to directly impinge against air during flight, ascent or descent, in order to form a smooth surface having increased lubricity and reduced air friction. The aircraft part may be in the form of a helicopter rotor, wing, propeller, fin, aileron, nose cone, and the like. The monomolecular carbon-based film can be deposited on the aircraft part, for example, using a reactor that includes a bed of silica and through which emissions from a diesel engine are passed. The monomolecular carbon-based film decreases air friction and increased lift of a modified aircraft that includes an aircraft part treated with the film. It also provides a structured shock absorber.

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 rod for supporting components in a fuselage cell structure, in particular produced with fibre-reinforced composite materials, of an aircraft, in particular for supporting at least one floor frame on at least one annular former and/or at least one further component of the fuselage cell structure. In accordance with the invention the rod is formed, at least in portions, with a metal foam, in particular for the absorption of crash loads. The metal foam can transform, irrespective of its inherent weight, the high mechanical loads produced in the event of an accident into work of plastic deformation at the microcavities within the metal foam. A first variant of the rod is almost completely formed by a metal foam, whereas a second embodiment comprises merely a core made of the metal foam with a relatively small diameter, which core is surrounded coaxially by a support casing and an outer sleeve.

Abstract: A strut (18, 50) includes a fiber-reinforced, elongate composite body (36, 54) and has first and second ends and a metal and/or tubular support (20, 52) at least partially embedded in the composite body (36, 54), where the composite body (36, 54) has first (40, 60) and second ends (42, 62) forming first and second load-bearing end portions of the strut (18, 50). Also a method forms such a strut using a mold (70).

Abstract: Aircraft empennage structures have impact resistance satisfying transport aircraft certification requirements (for example, FAR part 25 and/or other international certification requirements). The structures may be specifically embodied in leading edge structures that are preferably in the form of one-piece components formed from fiber-reinforced composite materials (e.g., reinforcement fibers such as glass fibers, aramid fibers and/or carbon fibers embedded in a polymeric matrix such as an epoxy resin). The leading edge structures will most preferably include an arcuate skin and longitudinally extending internal reinforcement which is generally Y-shaped and includes a planar rib element and a pair of planar divergent reinforcement arms. The rib element may be positioned so as to be coplanar with a longitudinal bisecting plane of the leading edge structure.

Abstract: An article such as a fan blade of a turbofan engine comprises a core made up of components at least some of which comprise packs of rods embedded in a resin matrix material. The rods extend in the span-wise direction of the blade to resist centrifugal forces imposed on the blade during operation. The core is encased in a skin formed from preforms which may comprise fabric reinforcements.

Abstract: A laminate is provided for forming an aesthetically pleasing fire worthy exterior surface on an object, for example a decorative interior sidewall used in a commercial aircraft. A first film layer forms an exterior surface of the sidewall, and an ink layer is disposed adjacent to the first film layer to provide color and/or images in an aesthetically pleasing manor, and to protect the subsequent layer from damaging UV light exposure. An embossing resin layer is disposed adjacent to the ink layer to improve the fire worthiness of the laminate by being situated adjacent to the exterior surface of the laminate. A second film layer is disposed adjacent to the embossing resin layer and a surface of the sidewall. The location of the embossing resin layer enhances the fire worthiness of the laminate by operating to absorb thermal energy released by the sidewall in the event the sidewall is exposed to fire or to a high temperature sufficient to otherwise cause higher thermal energy release of a conventional laminate.

Abstract: A composite structure is made by placing a stiffener preform and a composite structure reinforcement on a tool. The preform and the reinforcement are vacuum bagged processed and co-infused with a thermoset resin to form a stiffened, unitized structure.

Abstract: A composite panel having a first skin forming the outer wall of the panel, known as the lower face (1), and a second skin forming part of the inner wall of the panel. The second skin is formed hollow such as to produce an inner reinforcing frame (2) in the form of a half-box structure (3) which, together with the first skin, forms box parts. The aforementioned frame is equipped with a rim (4a, 4b) for fixing to the first skin, a face known as the upper face (5) and connecting side walls (6, 6a, 6b, 6c, 6d, 6e, 7) between the rim (4a, 4b) and the upper face (5) the first and second skins forming a monolithic self-stiffened panel. The half-box frame (2) has perforated sections (8) which are defined by hollow beam sections forming the half-box frame (2).

Abstract: Aircraft systems having baffle seals with low coefficients of friction and methods of assembling aircraft systems are disclosed herein. In one embodiment, an aircraft system includes: (a) an aircraft baffle; (b) an aircraft cowl separated from the aircraft baffle by a gap, the aircraft cowl having a contact surface; and (c) a flexible aircraft baffle seal extending from the aircraft baffle to the aircraft cowl to seal the gap, the aircraft baffle seal having a contact side for contacting the contact surface of the aircraft cowl, the contact side having a kinetic coefficient of friction that is not more than 0.4, the aircraft baffle seal comprising an elastomer sheet and a laminate, the laminate being the contact side.

Abstract: A nose cone formed from a shape memory alloy (SMA) having a recoverable strain of at least 2% collapses about the dome for storage, deploys at launch to protect the sensor dome and reduce drag during atmospheric flight and is shed to allow sensing for terminal maneuvers. The SMA is shape-set at elevated temperatures in its Austenite phase with a memorized shape having a radius of curvature greater than that of the sensor dome to reduce aerodynamic drag. The temperature is reduced and the SMA collapsed to conform to the curvature of the sensor dome within the recoverable strain for storage. A first mechanism is configured to return the collapsed SMA to its memorized shape at launch or prior to going supersonic. In one embodiment, the SMA is stored below its Martensite finish temperature in a temperature-induced Martensite phase in which case the mechanism heats the SMA above the Austenite finish temperature to return the material to its memorized shape.

Abstract: A vehicle includes a housing having a wall with a central port and radial port formed at a lower portion of the housing, a blower disposed within the housing to move air through the central port, a distributor coupled to the blower to distribute air inside the housing, a heat-absorbing body disposed within the housing to absorb thermal energy, and at least one arm coupled to the housing.

Abstract: A layered material for use in transparent, flame resistant components couples a thin, fire resistant outer polymeric film to a thicker, transparent inner polymeric material. The resultant transparent layered material meets FAA flammability requirements for OSU heat release, has excellent solvent resistance and cleanability, is scratch resistant, and is transparent enough for use in interior applications in the aerospace industry. The layered material is formed via a co-extrusion or co-lamination process.

Abstract: A multilayer structural element, such as a skin panel for an aircraft fuselage, comprises a plurality of metal webs and a plurality of fiber webs arranged as layers in a sandwich structure, where the metal webs of neighbouring layers form stepped connecting regions. A material strip is provided on an outer surface of the sandwich structure for at least partially covering an outlying connecting region, and the material strip enables the placement of rivets at uniform spacing for fastening a structure to the structural element.

Abstract: The present invention provides a stiffening element for stiffening a step-like substructure, in particular in an aircraft or spacecraft, having a foot section which is composed of a fiber composite material and is shaped in the form of a step such that it essentially fills the substructure. A method is also provided for production of a stiffening element for stiffening of a step-like substructure in an aircraft or spacecraft, according to which a profile part with a foot section which is composed of a carbon-fiber-reinforced plastic is provided, at least one layer of a glass-fiber-reinforced plastic is laminated onto the foot section, and the glass-fiber-reinforced plastic is milled corresponding to a step-like shape which essentially fills the substructure.

Abstract: Laminar flow surfaces with selected roughness distributions, and associated methods are disclosed. A representative method for designing an airfoil includes selecting a parameter that includes a flow behavior distribution and/or a surface shape for an airflow surface. Based at least in part on the selected parameter, the method can include (a) selecting a target roughness value and determining a chordwise location forward of which surface roughness is at or below the target roughness value and/or (b) selecting a target chordwise location and determining a roughness value for a region forward of the chordwise location, with the surface roughness at or below the roughness value. In particular embodiments, a percentage of a local chord length of the airfoil over which the roughness is below a target value decreases in a spanwise direction. In another embodiment, the roughness at a particular spanwise location can increase over at least three values, continuously, in a step manner, or otherwise.

Abstract: The invention relates to a material for absorbing noise in aircraft, comprising at least one barrier layer and at least one absorptive layer. The barrier layer is provided with an elastomeric material having a minimum specific density of 1 and a mass per unit area of 1.3 kg/m2 or less while the absorptive layer is equipped with a porous and/or open-cell material having a weight per volume of 4 to 50 kg/m3, preferably 10 to 40 kg/m3, more preferably 15 to 35 kg/m3 even more preferably 20 to 30 kg/m3. The mass per unit area of the sound-absorbing material amounts to 2 kg/m2 or less.

Abstract: This invention relates to improving the tolerance of composite laminate structures to damage and/or delamination, resulting in weakening of the structure and/or to reducing the possibility of such damage and/or delamination. In particular, but not exclusively the invention relates to protecting the end surfaces of composite laminate structures present on aircraft from impact damage. The invention provides an aircraft component comprising a composite laminate structure, the composite laminate structure including an end surface at which a multiplicity of the layers of the laminate structure terminate, wherein the aircraft component further comprises a deformable strip mounted to protect the end surface of the composite laminate structure from direct impacts.

Abstract: A pre-fabricated article for a composite aircraft component includes a dielectric layer and a pre-fabricated metalized pattern on the dielectric layer. The metalized pattern is fabricated on the dielectric layer prior to integration with the composite aircraft component. The metalized pattern diverts EME current away from EME protection areas of the aircraft component.

Abstract: A composite heat shield includes a ceramic composite heat shield body comprising a pair of generally spaced-apart heat shield side panels, a heat shield bottom panel extending between the heat shield side panels and having a heat exposure surface, a heat shield top panel extending between the heat shield side panels in spaced-apart relationship with respect to the heat shield bottom panel and having at least one heat shield surface and at least one heat shield cavity defined between the heat shield bottom panel and the heat shield top panel.

Abstract: The invention relates to a method of joining a thermoplastic material (10) to a fibre composite material (11), the thermoplastic material (10) being welded to the fibre composite material (11), wherein the fibre composite material is a fibre-rein-=forced matrix material or comprises a fibre-reinforced matrix material and wherein the matrix material is a thermoset or comprises a thermoset. Preferably, the thermoplastic material (10) and the fibre composite material (11) are welded together by motion welding and/or by induction welding. In addition, it is preferred that, for the preferred induction welding, an electrically conductive material is arranged in a joint area between the thermoplastic material and the fibre composite material and induction heat is generated in the electrically conductive material which is sufficient to cause the thermoplastic material to start to melt and to heat the fibre composite material.

Abstract: Process for manufacturing an assembly made of a composite material and wing and stabilizing element obtained by this process. The junction zone for a part (P) made of a composite material comprising two parts (A1, A2) in which the principal directions intersect at a non-singular angle, is made in a single operation by continuous lay-up. The 0° directions of the lay-up coordinate systems (R1, R2) of the parts (A1, A2) are oriented such that these directions make an angle of close to 180°??°, where ±?° represents the other directions of the lay-up coordinate systems (R1, R2). Thus, the layers of fibers oriented at 0° in one of the coordinate systems correspond approximately to a lay-up direction of ±?° in the other coordinate system.

Abstract: An apparatus may comprise a first skin panel having a first surface, a second skin panel having a second surface, a first flange located at an end of the first skin panel, a second flange located at an end of the second skin panel, and a strip having a third surface. The first skin panel may be located adjacent to the second skin panel such that the first flange and the second flange form a channel. The strip may be bonded in the channel. Fluid flow over the third surface of the strip, the first surface of the first skin panel, and the second surface of the second skin panel may have a desired state.

Abstract: A structure for an aerospace vehicle includes a composite beam chord clamped between first and second metal plates. The beam chord is clamped at a force that precludes or reduces beam chord delamination under axial loading during operation of the vehicle.

Abstract: Laminar flow surfaces with selected roughness distributions, and associated methods are disclosed. A representative method for designing an airfoil includes selecting a parameter that includes a flow behavior distribution and/or a surface shape for an airflow surface. Based at least in part on the selected parameter, the method can include (a) selecting a target roughness value and determining a chordwise location forward of which surface roughness is at or below the target roughness value and/or (b) selecting a target chordwise location and determining a roughness value for a region forward of the chordwise location, with the surface roughness at or below the roughness value. In particular embodiments, a percentage of a local chord length of the airfoil over which the roughness is below a target value decreases in a spanwise direction. In another embodiment, the roughness at a particular spanwise location can increase over at least three values, continuously, in a step manner, or otherwise.

Abstract: A method of manufacturing an aircraft having adjacent spaced apart surface panels using a hollow filler within the gap between the spaced apart surface panels.

Abstract: Embodiments of integral composite panels and joints for composite structures are described In one implementation, an integrated panel spanning substantially the entire wingspan of an aircraft, includes at least a center portion and a pair of outwardly projecting wing portions. The portions may include a skin formed from successive layers or plies of composite material which overlap and offset at the joint between respective sections creating a pad-up area to carry loads between the portions. In a particular implementation, the skin is laid over one or more structural stringers which are transitioned into the joints between sections such as by tapering of the thickness and/or stiffness of the stringer.

Abstract: A composite beam chord is between first and second reinforcement plates. The beam chord includes a first ply of reinforcing fibers with a fiber orientation of +? degrees with respect to a longitudinal axis of the beam chord and a second ply of reinforcing fibers with a fiber orientation of ?? degrees with respect to the longitudinal axis. The angle ? is between 2 and 12 degrees for suppression or delay of ply splitting.

Abstract: Systems and methods to distribute power in a system having a composite structure are provided. A particular aircraft includes a fuselage having a composite skin. The composite skin includes a first resin and a plurality of fibers. The fuselage also includes an expanded conductive patch bonded to an interior surface of the composite skin by a second resin. The expanded conductive patch is in electrical contact with the plurality of fibers. The fuselage also includes a connector in electrical contact with the expanded conductive patch.

Abstract: The present invention provides a profiled part, for use, in particular, within an aircraft or spacecraft, the profiled part comprising at least one hollow profiled part portion and a T-shaped and/or L-shaped profiled part portion.

Abstract: An aluminum alloy comprises aluminum, magnesium, scandium, and an enhancing system. The magnesium is from about 0.5 percent to about 10.0 percent by weight based on the aluminum alloy. The scandium is from about 0.05 percent to about 10.0 percent by weight based on the aluminum alloy. The enhancing system is from about 0.05 percent to about 1.5 percent by weight based on the aluminum alloy.

Abstract: Methods and systems for manufacturing composite components, including asymmetric composite components, are disclosed. In one embodiment, a method of forming a composite component includes providing a forming tool having a base portion and at least one retaining portion; forming an uncured composite component on the forming tool; curing the composite component; performing a manufacturing operation on the cured composite component engaged with the forming tool; and after performing the manufacturing operation on the cured composite component, removing the cured composite component from the forming tool. In further embodiments, the forming tool may have a plurality of reference apertures disposed therethrough for performing drilling operations on the cured composite component.

Abstract: A structural, energy absorbing composite element for aircraft structures includes a metallic or nonmetallic skin shell with a porous filler material insert positioned within the skin shell. The porous filler material is a material having ligaments that collapse resulting in a densification of the porous filler material in response to impact loading or a compression force sufficient to cause failure of the combined assembly.

Abstract: A monomolecular carbon-based film can be placed on an aircraft part, such as the leading edge designed to directly impinge against air during flight, ascent or descent, in order to form a smooth surface having increased lubricity and reduced air friction. The aircraft part may be in the form of a helicopter rotor, wing, propeller, fin, aileron, nose cone, and the like. The monomolecular carbon-based film can be deposited on the aircraft part, for example, using a reactor that includes a bed of silica and through which emissions from a diesel engine are passed. The monomolecular carbon-based film decreases air friction and increased lift of a modified aircraft that includes an aircraft part treated with the film.

Abstract: Hybrid fiberglass composite structures and methods for forming the same are disclosed. In one embodiment, a hybrid fiberglass composite structure includes a first laminate substrate and a second opposing laminate substrate. At least one of the first laminate substrate and the second laminate substrate includes at least one polymer fiber layer and at least one glass fiber layer. A cellular core structure is interposed between the first laminate substrate and the second laminate substrate and fixedly coupled to the first laminate substrate and the second laminate substrate.

Abstract: An unmanned aerial vehicle selectively formed of high strength composite structural part portions and lightweight aerodynamic foam portions to provide a low-cost and lightweight UAV that comports with export, civil airspace, and safety regulations. To further to reduce an overall weight of the UAV, mechanical elements are designed to provide multiple functionalities. Structural elements may be manufactured in same or similar non-specialized processes, and non-structural elements manufactured in same or similar non-specialized processes, reducing overall manufacturing costs. Materials and bonding elements are selected to provide frangibility and yet maintain normal flight structural integrity.