Abstract: A method of producing a metallic component includes: providing a body made of a first alloy; providing a preform comprising a metallic powder made of a second alloy and formed in the shape of an extension of the body; and heating the preform with microwave energy to sinter the metallic powder together and to bond the preform to the body. The method may be used to make new components as well as to repair or modify existing components.

Abstract: The present invention relates to a stringer for an aircraft wing and a method of reinforcing such a stringer in which the stringer is formed with a substantially lambda shaped section.

Abstract: An aircraft component, such as a wing, has perforations through an outer wall for boundary layer suction. In the space between the outer wall and an inner wall partition walls form pressure channels and suction channels that are adjacent to each other and alternate, which channels communicate with the perforations. For example, alternating channels may be formed by a corrugated structure having trapezoidal corrugations providing a larger area for the suction channels than for the pressure channels. The pressure channels may be coupled to a hot-air reservoir by a control device, lines and valves, and the suction channels may be coupled to a vacuum reservoir, unless a short-circuit valve is used to cross connect the lines.

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: An aircraft wing (104) comprises a rear spar (108), a first skin (110), a second skin (112) overhanging and converging to an opening (114) of width (A), a support structure (200) comprising a bracket (202) engaging the spar (108) and skins (110, 112) and a rib (216) engaging the bracket (202) via a rib attachment portion (218) with a dimension (C) less than dimension (A).

Abstract: A rib post (10) has a hollow substantially triangular cross-section comprising composite material. The rib post (10) can form a structural member to join two adjacently placed components together, for example a spar and internal rib in a wing of an aircraft. The rib post (10) can be formed from first (12), second (14) and third (16) walls. The first wall (12) can abut a surface of one component and the second wall (14) can abut a surface of a second component. The first and second walls (12, 14) can be attached to the respective components, such as a rib and spar, to join them together. The third wall (16) joins together the first and second walls (12, 14) to form the hollow triangular section.

Abstract: The structural strength of an aircraft wing is continuously tailored along its length to closely match continuously varying imposed wing loads. The wing comprises upper and lower skins each formed from panels joined together. Each of the skin panels includes a stiffening web sandwiched between and joined to the inner and outer facesheets. The web has a repeating pattern and at least one dimension that continuously varies along the length of the panel.

Abstract: A capped winglet comprising a helical portion having a smooth vertical curvature of 180°±approximately 20° from the plane of the unbounded end of a wing, and a generally planar cap contiguous with the helical portion of the winglet at the end of its curvature away from the wing, the cap having a free inboard end and a chord tapered toward the free end.

Abstract: An airplane wing has a lower wing cover, an upper wing cover and a wing spar. A first end portion of the wing spar is welded to the lower wing cover and/or a second end portion of the wing spar is welded to the upper wing cover.

Abstract: A rib of an airplane wing is prevented from deforming due to a machining load when machining a rivet hole in parts where the rib and a skin are joined. With regard to a rib that provides a connection between first and second spars of the airplane wing, the rib is formed with mounting flanges secured to a skin via a rivet, a stringer through hole for a stringer to be positioned through, and first and second cutouts facing parts where the first and second spars and the skin are connected. Since reinforcing beads are formed along the mounting flanges and reinforcing flanges are formed so as to extend in the span direction from the edges of the stringer through hole and the first and second cutouts and be connected integrally to the mounting flanges, the stiffness of the rib toward a load in the vertical direction when drilling a rivet hole can be enhanced.

Abstract: A composite shear tie for connecting a rib in a wing of an aircraft to a skin panel. The shear tie includes a web section having a first edge, a second edge, and a third edge. The web section is formed from a composite material. The first edge is parallel to the second edge, and the web section is configured to be connected to the rib in the wing of the aircraft. The shear tie also has a first free flange extending from the first edge, wherein the first free flange is around perpendicular to the web section and wherein the first free flange is formed from the composite material. A second free flange is present in the shear tie in which the second free flange extends from the second edge, wherein the second free flange is around perpendicular to the web section and wherein the second free flange is formed from the composite material.

Abstract: A pivoting panel for an aircraft and to a composite support piece for this kind panel. The support piece (17) is arranged as a part of the frame structure of the panel (3). The support piece has an extension portion (23) to which an actuator connection fitting (7) of an actuator (4) can be fastened. The extension portion (23) distributes the forces to the structures of the panel (3).

Abstract: A structural component such as a wing skin for an aircraft is formed generally of a first composite fiber and matrix material exhibiting a first tensile modulus in respect of tensile stress and a first tolerance of strain resulting from such stress. One or more openings are provided in the component and the material of a portion of the component bounding the opening(s) is a second composite fiber and matrix material. The second material exhibits a second tensile modulus in respect of tensile stress that is smaller than the first tensile modulus of the first material and a second tolerance of strain resulting from such stress that is greater than the first tolerance of strain of the first material.

Abstract: An integrated aerodynamic panel—e.g., for a trailing edge or leading edge of an aircraft aerodynamic surface—includes a first panel region defining inner and outer mold lines, and a second panel region contiguous with and extending from the first panel region in a tapering fashion. A splice plate region extends from the second panel region and includes an edge band region configured to accept a fastener. A filler region (e.g., a SYNCORE or fiberglass structure) adjacent the second panel region has an exposed surface substantially flush with the outer mold line.

Abstract: A leading edge structure is described for wing structures and empennage, comprising an outer shell suitable to define a front portion of a airfoil. The outer shell is formed by a bent plate of fiber-reinforced thermoplastic resin composite material. The structure further comprises an inner shell having a convex profile that is oriented in the same direction as the profile of outer shell, which is formed by a bent plate of fiber-reinforced thermoplastic resin composite material and is bonded to the outer shell at the longitudinal edges thereof, and at least one reinforcing element, transversally extending such as to connect the outer shell and inner shell to each other , which is formed by at least one piece of fiber-reinforced resin composite material and is fixed at opposite ends to the outer shell and inner shell, respectively.

Abstract: A metallic-composite joint is formed by inserting a metallic member into a slot of a pi-shaped composite preform. The preform is formed from woven carbon fiber in a binder of resin and may or may not be cured prior to assembly. An inert compliant layer is located between the legs of the preform and the metallic member. The resin binder or an adhesive is used to bond the compliant layer to the preform. The compliant layer has a coefficient of thermal expansion that more closely matches that of the preform. The properties of the compliant layer also avoid galvanic corrosion between the carbon in the preform and the metallic member.

Abstract: An airborne mobile platform may have a wing having a length and a chord wise dimension, and a plurality of elongated structural components extending span-wise along the length of the wing. The elongated structural components may each have a localized hinge area. A device may be used for manipulating the elongated structural component to position the localized hinge area to selectively change a hinge line of the wing in response to an airflow over the wing.

Abstract: A device for reinforcement of hollow structures such as aircraft boxes, composed of a front panel, an upper panel, a lower panel, and a rear panel, the device comprising a network of reinforcement rods with the ends attached to at least some of the panels comprising the structure. The device comprises at least a first rod and a second rod, both inclined, the first end of the first rod being attached to the internal side of the front panel, the second end of the first rod being attached to the internal side of the upper panel, the second end of the second rod being attached to the internal side of the lower panel, the device comprising in addition a third rod connecting said second ends of said first and second rods, the first, second and third rods thus comprising a reinforcement triangle.

Abstract: A structure protected from premature buckling for resisting multiple bending forces resulting from design loads in which the resulting structural configuration is determined by the force components produced by the design bending stresses. The structure is a compound curved, relatively thin sheet of metal or a composite having a longitudinally cambered shape that is initially arched opposite to the bend that occurs when the structure is loaded and is likewise curved laterally through a substantial portion of its length. The structure is particularly suited for the construction of fixed wing airplanes and airplane fuselages using the same principles of construction. The wing structure is characterized by a concave undersurface like that of a bird's wing which results in a greatly reduced stall speed which cannot stall out at slow speeds.

Abstract: An aircraft structure comprising: a rear spar; and a composite cover attached to the rear spar and having an overhanging portion extending to its rear. The overhanging portion of the cover comprises one or more ramps along which the thickness of the cover reduces. A second cover is attached to the rear spar and overhangs to its rear. A hinge rib is attached to one or both of the covers, and an aerodynamic control element such as an aileron, rudder or elevator is pivotally mounted to the hinge rib.

Abstract: There is provided a method of manufacturing a rib for an aircraft wing, the method comprising providing a billet of material for forming the rib, machining away material from all sides of the billet to form a rib for use in an aircraft, the rib including a web that extends between first and second flanged portions, each flanged portion having a flange that facilitates fixing of the rib to another component of the aircraft, wherein the material that forms the web extends diagonally across the billet, the web having a first face on a first side and a second face on a second side opposite to the first side of the web, and the first flanged portion extends more to the first side of the web than to the second side and the second flanged portion extends more to the second side of the web than to the first side.

Abstract: An aircraft wing box comprises a pair of half ribs (18, 20) joined to form a rib extending from an upper wing skin (12) to a lower wing skin (10). Each half rib (18) comprises a base (19) and a plurality of projections (4) that extend away from the base (19), the projections (4) defining a series of troughs (2a) and peaks (4a). Respective projections of one half rib (18) overlap with respective projections of the other half rib (20) thereby defining overlapping portions (3), the half ribs (18, 20) being joined in the regions of the overlapping portions (3). There may be projections (4) shaped such that the overlapping portion (3b) defined by the two projections is separated from the neutral axis N of the rib (18, 20).

Abstract: Apparatuses and methods for joining composite members and other structural members in aircraft wings and other structures. An aircraft wing box structure configured in accordance with one embodiment of the invention includes a first composite member having a first surface portion positioned at an angel relative to a second surface portion of a second composite member. The wing box structure of this embodiment further includes at least one metallic joining member having an upstanding leg portion extending from a base portion. The base portion of the joining member is bonded to the first surface portion of the first composite member with a first portion of adhesive, and the upstanding leg portion of the metallic joining member is bonded to the second surface portion of the second composite member with a second portion of adhesive.

Abstract: A wing cover panel assembly, wing cover and method of forming thereof are disclosed in which an attachment surface of the wing cover panel is provided with a locating channel into which a wing structural element such as a stringer or spar is captured so as to fix the wing structural element to the wing cover panel.

Abstract: There is provided a method of manufacturing a rib for an aircraft wing, the method comprising providing a billet of material for forming the rib, machining away material from all sides of the billet to form a rib for use in an aircraft, the rib including a web that extends between first and second flanged portions, each flanged portion having a flange that facilitates fixing of the rib to another component of the aircraft, wherein the material that forms the web extends diagonally across the billet, the web having a first face on a first side and a second face on a second side opposite to the first side of the web, and the first flanged portion extends more to the first side of the web than to the second side and the second flanged portion extends more to the second side of the web than to the first side.

Abstract: The invention is directed a trapezoidal panel pin joint for allowing deflection of an aircraft between a fuselage section and a wing section, wherein the pin joint comprises a vertical pin portion, a lug portion, and a clevis portion; wherein at least one vertical flexible tee member is positioned below the pin joint; and wherein the pin joint is coupled to at least two horizontal flexible tee members, such that the pin joint and vertical and horizontal flexible tee members, in combination, release two rotational degrees of freedom.

Abstract: A split-cycle aircraft engine includes a crankshaft rotatable about a crankshaft axis. A power piston is slidably received within a power cylinder and is operatively connected to the crankshaft such that the power piston reciprocates through an expansion stroke and an exhaust stroke during a single rotation of the crankshaft. A compression piston is slidably received within a compression cylinder and is operatively connected to the crankshaft such that the compression piston reciprocates through an intake stroke and a compression stroke during a single rotation of the crankshaft. A gas crossover passage operatively interconnects the compression cylinder and the power cylinder. An air reservoir is operatively connected to the gas crossover passage by a reservoir passage. The air reservoir is selectively operable to receive and deliver compressed air. The engine is mounted to an aircraft and the air reservoir is disposed within the aircraft.

Abstract: A wing load alleviation structure for use on an aircraft, comprising a front spar, wherein the front spar includes a plurality of alternating rigid spar structures and inflatable spar structures; and a rear spar, wherein the rear spar includes a plurality of alternating rigid spar structures and pivot joints, such that when a load is applied to the front and rear spars, deflection of each of the front and rear spars continues at a first rate until a critical load is reached, and then as the loading increases, deflection of each of the front and rear spars continues at a second rate.

Abstract: A method of locally reinforcing an element made of composite material having a longest dimension, such as a composite beam, which is intended to bear multidirectional forces. To do that, at least one metal reinforcing piece is secured to the composite element, at a point on the composite element that is intended to bear vertical forces. The same may be done at each of those points on the composite element that is intended to bear vertical forces.

Abstract: The present invention concerns a sort of split return wing. The split return wing contains main wing body, which connects with the fuselage main support frame, and the tilting-able segments of main wing-spar are installed on the fuselage main support frame actively. Two ends of the tilting-able segments of main wing-spar connect with the fixed segments of main wing-spar actively. Two split return wing sector wing-pieces are installed on the right and the left tilting-able segment of main wing-spar separately. There is one group tilting-able dynamic device in front of every split return sector wing-piece. The split return sector wing-pieces can be in the perforation spaces of the main wing body, and their shapes are same. And two big wingspan outer-wings are set on the right and left of the main wing body separately. The Split return wing will decrease unbalanced moment of aircraft to the minimum when the aircraft is taking off & landing vertically.

Abstract: A method of making a shim comprises the steps of providing a first structural component, such as a rib foot of an aircraft wing. The rib foot is then scanned or mapped to generate the surface profile of the rib foot. The shim is then formed to include a feature on its surface which complements the surface profile of the assembly surface.

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 method and apparatus for manufacturing and installing a fairing on an aircraft. In one advantageous embodiment, a composite fairing is installed on the aircraft. A composite sheet having a form of the composite fairing is formed. A plurality of cured composite tiles is created from the composite sheet, wherein spaces are present between the plurality of composite tiles. The plurality of composite tiles is cured to form the composite fairing. The composite fairing is attached to a surface of the aircraft. The spaces between the plurality of composite tiles are filled with a flexible filler.

Abstract: The invention concerns an assembling device including a splice bar (1) and a backing plate (2) for securing at least first (3) and second (4) structural elements received between the splice bar (1) and the backing plate (2). The invention is characterized in that the backing plate (2) includes a number of fixing points (M, N) less than the number of fixing points (P, Q) of the splice bar so that the first (3) and second (4) structural elements are fixed between the splice bar (1) and the backing plate (2) by a number of fixing points (M+N) less than the number of fixing points (P+Q) of the structural elements on the slice bar (1), the assembly being thus a non-symmetrical assembly, at least one of the structural elements being an element made of composite materials.

Abstract: An airfoil for a micro air vehicle that includes components enabling the airfoil to adjust the angle of attack (AOA) of the airfoil in response to wind gusts, thereby enabling the airfoil to provide smooth flight. The airfoil may include a first compliant region positioned between an inboard section and a first outboard section and may include a second compliant region between a second outboard section and the inboard section. The compliant regions enable the first and second outboard sections to bend about a leading edge section and move relative to an inboard section. This action creates smoother flight due to numerous aerodynamic advantages such as a change in the angle of attack and improved wind gust rejection due to adaptive washout as a result of the airfoil flexing, twisting and decambering.

Abstract: A solar powered aerial vehicle includes an elongated airframe incorporating lifting and control surfaces, a mechanism for propelling the airframe through the air such that lift developed by the lifting surface is equal to or greater than the weight of the aerial vehicle, a planar solar sail coupled to the airframe and having at least one surface adapted to collect solar energy during the day and to power the propelling mechanism with a first portion of the energy collected, and an apparatus such as a fuel cell/electrolyzer for storing a second portion of the solar energy collected by the solar sail during the day and for powering the propelling mechanism with the second portion of energy during the night. The vehicle is capable of continuous operation at northern latitudes and during the winter months for extended periods without landing or refueling.

Abstract: An intercostal for an aircraft for dissipating a load from a first frame onto a second frame and/or a skin of the aircraft, the intercostal having a framework.

Abstract: A wing for use on a supersonic aircraft that includes an inboard section a central section of the wing outboard of the inboard portion, and an outboard section. The outboard section can be a winglet oriented anhedrally relative to a lateral axis of the supersonic aircraft. Leading edge segments on the inboard section, central section and outboard winglet may have mounted thereon leading-edge flaps. These flaps are adjusted by a control system operable to reposition the leading-edge flaps in order to improve the aerodynamic performance of the supersonic aircraft. This winglet promotes sonic boom minimization. Further, the wing tip anhedral allows greater inboard dihedral. This effectively pushes lift aft for sonic boom and control purposes while minimizing the movement of control surfaces.

Abstract: Al—Mg—Ag wrought products and methods of making such products useful in aircraft applications. The Al—Mg—Ag wrought products have improved strength when compared to traditional AA5XXX alloys. The alloys may comprise from about 3.5 to about 10 weight percent Mg, from about 0.05 to about 0.5 weight percent Ag, from about 0.01 to about 1.0 weight percent Mn, from about 0.01 to about 0.15 weight percent Zr, and the remainder Al and incidental impurities. In addition, from about 0.05 to about 0.4 weight percent Sc may be added to further improve the strength characteristics.

Abstract: Rib structure for a torsion box of an aircraft wing or a horizontal stabiliser. The structure comprises a center element which extends between the front spar and the rear spar of the torsion box, a series of vertical stiffening elements arranged between an upper skin and a lower skin of the torsion box. The vertical stiffening element consists of a first lateral wing which extends in the direction of said front spar and a second lateral wing which extends in the direction of said rear spar. These lateral extensions emerge from respective opposite sides of the vertical body and are attached side to side to the center element of the rib.

Abstract: In a method for manufacturing a structural element intended for aeronautical construction, at least a first and second metal block are made available, the limit of elasticity under compression of the first metal block being greater than that of the second metal block. The first metal block is machined in such a manner as to obtain a first machined monolithic part which has a first web portion and at least one stringer element whose height is such that a stringer portion extends beyond the first web portion. There is prepared, by shaping the second metal block, at least one second part having at least a second web portion capable of co-operating with the first web portion to form the web. The first monolithic part and the second part are assembled by placing the first and second web portions end-to-end over their entire common length.

Abstract: Methods and structures for a composite truss structure are provided. The structure includes an upper chord member, a lower chord member, and a plurality of web members extending therebetween. Each of the upper chord member, the lower chord member, and the plurality of web members are formed of a continuous composite fiber positioned in each of the upper chord member, the lower chord member, and each of the plurality of web members wherein each of the upper chord member, the lower chord member, and the plurality of web members includes a predetermined number of passes of the continuous composite fiber corresponding to a predetermined load. The composite truss structure also includes at least a first gusseting plate coupled to the upper chord member, the lower chord member, and the plurality of web members.

Abstract: Methods and structures for a composite truss structure are provided. The structure includes a web formed of a plurality of sheets of composite material, each sheet including a first face and an opposing second face and each face including a length and a width. Each of the plurality of sheets are coupled to at least one other of the plurality of sheets face to face such that the length and width of each face substantially match the length and width of a face of an adjacent sheet. The plurality of sheets are formed to include an upper chord member, a lower chord member, and a plurality of web members extending therebetween. The structure also includes at least a first flange plate coupled to the web proximate an outer periphery of the web.

Abstract: A rib element for aircraft and a composite flange for use in aircraft. The rib element (24) and the composite flange (14), manufactured from a composite material, comprise at least two elongated flanges (15, 16, 26) interconnected with one another in a corner (17) provided with a rounded inner corner (18) and a rounded outer corner (19). The corner is reinforced by at least one reinforcing component (20) which covers at least a part of the inner corner.

Abstract: Sealing device for the flap track for the flap-actuating shaft, in the belly fairing of an aircraft, with a top hollow body 1 and a bottom hollow body 7 which are flexibly compressible, each comprising at least one interior chamber 2, 8 with open sides 2a, 2b, a wall 3, 9 with respective external contact surfaces 3a which are in flexible contact with each other, the outer surfaces 4a, 10a of the front walls 4, 10 of the hollow bodies 1, 7 jointly forming an outer sealing surface 11 which covers the flap track 19, said hollow bodies 1,7 having jointly a vertical compression flexibility such that the contact surfaces 3a, 9a retract around the actuating shaft 20, the interior chambers 2, 8 extending between the front walls 4, 10 and rear walls 5, 11 of the hollow bodies 2, 7.

Abstract: Methods and systems for manufacturing composite aircraft wings and other structures are disclosed herein. A tool assembly for use in manufacturing composite laminates in accordance with one embodiment of the invention includes a tool plate carried by a movable support system. The tool plate includes a tool surface configured to support fiber-reinforced resin material and define an outer mold line (OML) of the fiber-reinforced resin material. The movable support system is configured to respond to signals from a controller to automatically change the shape of the tool surface and alter the OML of the finished part to suit the particular application. In one embodiment, the movable support system can include a plurality of telescoping actuators operably coupled to the tool plate.

Abstract: The invention relates to an aircraft wing panel, for example, a trailing edge upper panel 1 which is joined to the wing skin (20) by means of spigots (19) which project and engage with slots (13). Preferably, the slots are provided in lips (10, 11, 12) which extend from the leading edge (5) of the panel and the spigots (19) are fastened to and extend from the wing skin (20).

Abstract: Variation in the contours of first and second compliant surfaces is produced by a compliant frame having a first resiliently variable frame element (120) having a corresponding first outer surface (122) and a first inner surface (124), and a second resiliently variable frame element (130) having a corresponding second outer surface (132) and a second inner surface (134). The first and second outer surfaces (122, 132) communicate with respective ones of the first and second compliant surfaces. A linkage element (141-144) having a predetermined resilience characteristic is coupled at a first end thereof to the first inner surface (124) and at a second end thereof to the second inner surface (134). A frame coupler (151) couples the first resiliently variable frame element (120) to a support element (150).

Abstract: In a wing structure for an aircraft, a rib extends in a chord direction for connecting spars to each other with upper and lower stringer through holes through which stringers of skins pass being formed in upper and lower edges of the rib. A bead is formed on the rib so as to extend between the upper and lower stringer through holes. Upper and lower ends of the bead are formed into arcuate shapes so as to surround the upper and lower stringer through holes, respectively. Thus, it is possible to enhance the buckling strength of the rib against a shear load acting on the wing of the aircraft.

Abstract: A laminated composite material structure, in the form of a load-bearing rib of a wing of an aircraft, comprises an upper laminated portion, which is angled with respect to a middle laminated portion, and a curved corner portion being continuous with and interposed between the upper and middle portions. (The rib has a curve that connects the upper and middle portions, which are perpendicular or transverse to each other). The corner portion is in the form of a kinked portion or joggled, that is, the portion includes two regions of positive curvature (concave regions) and a region of negative curvature (convex region) interposed therebetween, whereby the formation or propagation of delaminations or cracks at the bend between the middle and upper portions, due to fuel pressure loading and low through-thickness strength, may be reduced.