343results about "Spars/stringers" patented technology

An assembly for an aircraft includes a wing assembly, a plurality of underwing heat exchangers, and a mounting assembly. The wing assembly includes a wing body. The plurality of underwing heat exchangers are arranged along the wing body in a longitudinal direction. The plurality of underwing heat exchangers includes a first underwing heat exchanger and a second underwing heat exchanger. The first underwing heat exchanger is longitudinally adjacent the second underwing heat exchanger. The mounting assembly includes a first heat exchanger bracket, a second heat exchanger bracket, and a first support bracket. The first heat exchanger bracket is fixedly mounted to the first underwing heat exchanger. The second heat exchanger bracket is fixedly mounted to the second underwing heat exchanger. The first heat exchanger bracket and the second heat exchanger bracket are pivotably mounted to the first support bracket. The first support bracket is fixedly mounted to the wing body.

The invention discloses a deformable wing with the characteristics of large unfolding-folding ratio and high bearing capacity. The deformable wing comprises a bearing structure, a driving mechanism, a wing surface dimensional structure and an accessory structure, the bearing structure comprises a plurality of groups of telescopic mechanisms, wing ribs and corner boxes; the accessory structure comprises a mounting plate, and the mounting plate is connected with the machine body; each group of telescopic mechanisms comprises a plurality of stages of telescopic rods which are arranged along the chord length direction and are fixed on the mounting plate; the wing ribs are arranged in the wingspan direction and are connected with the multi-stage telescopic rods through the angle boxes to form a wing framework; the driving mechanism penetrates through the hollow parts of the wing ribs, and two ends of the driving mechanism are respectively connected with the mounting plate and the wing rib on the outermost side of the wing; the wing skeleton is externally coated with the wing surface dimensional structure to form a wing shape.

Aerodynamic lifting structures, such as aircraft wings, having embedded engines and associated methods and systems are disclosed herein. A wing assemblies configured in accordance with embodiments of the present technology can include, for example, an upper wing portion, a lower wing portion, and a plurality of independent ducts positioned between the upper wing portion and the lower wing portion. Each duct can extend between a corresponding inlet positioned toward a leading portion of the wing assembly and a corresponding outlet positioned toward a trailing portion of the wing assembly. The wing assembly can further include a plurality of fans and a plurality of electric motors operably coupled to the plurality of fans. The fans and electric motors are positioned in the corresponding individual ducts and the fan is rotatable to propel fluid received in the inlet through the duct to create lift.

Disclosed here are unmanned aerial vehicle embodiments including some embodiments having a fuselage, tail, and wings including example embodiments with an adaptable payload section, alternatively or additionally, modular flight surfaces including tail, wings and motor, alternatively or additionally the vehicle configured for short landings with reversible thrust, alternatively or additionally, the unmanned aerial vehicle configured with direct connection to moveable flight control surfaces.

An electric aircraft according to an embodiment of the present invention comprises: a fuselage equipped with a power means, a front spar and a rear spar extending from the fuselage to an end of a wing, and a plurality of ribs extending from the rear spar to the front spar and coupled to the front spar and the rear spar, in which a plurality of solid state batteries are mounted in a plurality of individual spaces partitioned by the front spar, the rear spar, and the plurality of ribs, respectively, and the front spar and the rear spar are used as members for serial connection of the plurality of solid state batteries, and the plurality of ribs are used as members for parallel connection of the plurality of solid state batteries.

A device to control movement of fluid that is emitted from an aircraft engine that is mounted to a wing of an aircraft. The device includes a support member positioned on an exterior of the aircraft and having an elongated shape with a first end connected to the wing and a second end connected to the fuselage. The support member has a first sectional size. A deflector shield is mounted to the support member and positioned on the exterior of the aircraft. The deflector shield has a larger sectional size than the support member. The deflector shield has an outer edge that extends around the support member and is positioned radially outward from the support member.

A wing-fuselage truss joint assembly for an aircraft includes joining members secured to the aircraft's wing and fuselage structures. The joining members each include forward and aft wing joints and forward and aft fuselage joints connected by an upright member. The forward joining members are pivotally connected using interconnecting members and the aft joining members are connected using aft truss members. The forward-right and aft-right joining members and the forward-left and aft-left joining members are connected by truss link members and diagonal truss link members. The wing-fuselage truss joint assembly is configured to reduce forces experienced by the wing from being transferred to the aircraft fuselage.

The invention provides a central wing box and an electric tilting helicopter composite material wing, and belongs to the technical field of helicopter structural design, the central wing box adopts a double-beam type structure, the central wing box comprises a force-bearing covering cap, a front beam, a rear beam, a lower skin and wing ribs, and the front beam, the rear beam and the lower skin jointly form a battery installation room. The battery mounting room is divided into a plurality of battery cabins through a plurality of wing ribs, each battery cabin is provided with a plurality of battery mounting bases, and the force-bearing opening cover is positioned at the upper part of the battery mounting room; the front beam, the rear beam, the first central wing rib and the second central wing rib are connected with the fuselage. According to the invention, the integrated design of the wings is realized, the purpose of weight reduction is realized, the problem of difficult assembly and adjustment of the aircraft body and the wings is solved, the problem of large-opening load transfer caused by battery installation is solved, and the maintainability is greatly improved by adopting the bolt connection measures of the covering caps with uniform specifications.

An aircraft wing assembly includes a wing box including an upper surface, a lower surface, and a leading edge structure connecting the upper surface to the lower surface and secured to the upper surface at a connection point. The aircraft wing assembly includes a drive mechanism secured to the wing box and a high lift device mounted to the drive mechanism. The drive mechanism is configured to move the high lift device between a retracted position and an extended position in use, and wherein the high lift device is configured to overlap the connection point when in the retracted position.

A wing for a mobility apparatus include battery modules a plurality of ribs that are spaced apart from one another in a longitudinal direction of the wing, the plurality of ribs respectively defining a plurality of first through-holes through which the battery modules pass, a plurality of first flange portions that are respectively disposed at the plurality of ribs along edges of the first through-holes, the plurality of first flange portions defining battery seating surfaces that support the battery modules, and a spar that supports the plurality of ribs.

Truss-braced wing aircraft engine mounting and associated systems are disclosed. An example aircraft includes a fuselage including a central structural section, a wing extending from the fuselage, a truss extending from the central structural section and coupled to the wing, a pylon coupled to the central structural section, and an engine coupled to the pylon, wherein the pylon is positioned at an angle substantially 45 degrees from a horizontal plane and a vertical plane.

A lightweight structure for a vehicle or aircraft includes a longitudinal member with a base bridge, having a first collective conductor and a transversal member with a central bridge and transversal bridge with a first connection conductor extending on a first surface of the transversal bridge and on a second surface of the transversal bridge oriented opposite the first surface, and a second connection conductor extending separately from the first connection conductor. The transversal member is connected to the base bridge at the first end section, so the first connection conductor contacts the first collective conductor of the base bridge. The lightweight structure includes a carbon fiber structural battery (CFSB) connected with the central bridge of the transversal member, a first collector of the CFSB electrically connected to the first or second connection conductor and a second collector of the CFSB electrically connected to the other connection conductor.

An aircraft wing apparatus includes a plurality of wing-skin sections, an extendable spar assembly, and a controller. The plurality of wing-skin sections includes a first wing-skin section and a second wing-skin section. The extendable spar assembly includes a first spar member coupled with the first wing-skin section and a second spar member coupled with the second wing-skin section. The controller is constructed and arranged to move the second spar member to predefined positions relative to the first spar member to place the plurality of wing-skin sections into predefined configurations. The predefined configurations including a stowed configuration, a first deployed configuration, and a second deployed position that is different from the first deployed position.

The invention discloses a built-in main control actuator connecting structure of an aircraft wing box, and relates to the technical field of aircraft structure design, an actuator is located in the aircraft wing box structure, an actuating rod of the actuator penetrates through an actuator rod avoiding hole of the aircraft wing box structure, and the actuating rod stretches out and draws back to cooperate with the actuator to rotate up and down in the wing box to achieve deflection of a control surface. The invention is suitable for small airfoils and small and medium-sized airfoils with low airfoil height, solves the problems that the small airfoils of the small airfoils are small, the trailing edge space is limited, and a control surface actuating system cannot be mounted, and avoids the problem that the aerodynamic efficiency is influenced as an actuating structure is arranged outside the airfoils; the structure of the actuator connector is designed and arranged on the front main bearing beam of the wing box, the load transfer continuity of the wall plate structure is not broken, meanwhile, the actuating load can be directly diffused to the skin through the support connector, and the load transfer efficiency is higher. By separating a daily inspection and maintenance passage from a disassembly and assembly passage, the opening size of the airfoil wall plate is reduced, the load transfer capacity of the lower airfoil skin is guaranteed, and structural weight reduction is facilitated.

Apparatus and methods for ultrasonic inspection of elongated composite members in a single scan pass using pulse echo phased arrays (97, 98, 99) operating in a bubbler method. The system concept is fully automated by integrating an inspection NDI probe assembly (32, 34, 36) to a robot (100) and using the robot to move the inspection NDI probe assembly along the part (i.e., outside of an inspection tank); and by integrating tooling fixtures that move out of the way as the inspection NDI probe assembly travels along the length of the part during the inspection. In addition, the system allows for generally elongated composite members having lengthwise variation in shape, curvature and dimensions.

A method for producing a torsion box for a structure of an airplane. The method includes providing a first component made of a fiber composite material, the first component has a first planar base having a first inner side and a first outer side, first stiffening elements on the first inner side forming a composite with the first base. A second component is provided of a fiber composite material and has a second planar base having a second inner side and a second outer side. Second stiffening elements are on the second inner side and form a composite with the second base. The method includes superimposing the first component and the second component such that the first stiffening elements lie, at least in some areas, on the second inner side and the second stiffening elements lie, at least in some areas, on the first inner side. The methods includes connecting the first stiffening elements to the second base and connecting the second stiffening elements to the first base.

A composite laminated layup panel structure suitable for wing combined connection, comprising a main box section skin (100) and a plurality of stiffeners (200) which are integrally cured and formed. The main box section skin is sequentially provided with recesses (300) at a front edge and a rear edge, forming a front recessed edge (110) for overlapping with a leading edge skin (400) and a rear recessed edge (120) for overlapping with a trailing edge skin. A front spar stiffener (600) configured to be connected to a front spar web (500) is fixedly arranged at the bottom of a transition region between the front recessed edge (110) and a main box section skin main body. A rear spar stiffener (700) configured to be connected to a rear spar web is fixedly provided at the bottom of a transition region between the rear recessed edge (120) and the main box section skin main body. The panel structure simplifies connection forms of panels and related structures, reducing the number of parts, improving assembly efficiency, and also facilitating disassembly and replacement maintenance during aircraft use and maintenance.

An aircraft component including a structural foam having a predetermined shape, an outer flange located near a first end of the structural foam, an inner flange located near a second end of the structural foam, a web skin encapsulating the inner flange and at least a portion of the structural foam wherein the structural foam is affixed to the web skin, wherein the aircraft component comprises a portion of a blended wing body aircraft (BWB), the BWB having no demarcation between a wing and a main body of the BWB.

The utility model belongs to the field of aircraft structure design, and particularly relates to a high-rigidity shear angle sheet structure which comprises a turned-over edge, a web plate, vertical ribs and edge strips. The flange, the vertical ribs and the edge strip define a whole ring structure, the web plate is integrally connected to one end of the whole ring structure, the flange is located on the outer side of the whole ring structure, the two sets of vertical ribs are symmetrically arranged on the two sides of the flange, and the edge strip is correspondingly arranged on the inner side of the whole ring structure; the outer walls of the vertical ribs are connected with the stringer, and the vertical ribs are arranged between the adjacent vertical ribs on the two sides of the stringer. The turnup, the web plate, the vertical rib, the edge strip and the supporting column are designed to be equal in height at the notch of the stringer, so that a thin neck structure is avoided. When the load is large, the phenomenon of stress concentration on the shearing angle piece is avoided. The rigidity of the wing rib at the position of the shearing angle piece is improved by changing the structural form among the traditional wing rib shearing angle piece, the edge strip and the supporting column, and then the structural strength requirement is met.

Assembly comprising a wall and at least one beam provided with at least one attachment passing through the wall, central wing box of aircraft comprising said assembly. The invention relates to an assembly comprising: at least one wall (54) having an inner face (Fint) and an outer face (Fext) opposite the inner face (Fint), at least one beam (52) positioned against the outer face (Fext) of the wall (54) and connected to the latter by connecting elements (56), at least one attachment (62) projecting from the inner face (Fint) of the wall (54) integral with the beam (52) and forming with the latter a single and the same piece made in one piece, the wall (54) comprising, for each attachment (62), a through hole (66) to allow the attachment (62) to pass through the wall (54).The invention also relates to an aircraft wing center box comprising at least one such assembly and to an aircraft comprising said wing center box. Figure 11.

The application provides a swept-back three-beam thin wing composite wing structure, which comprises a wing main body and an aileron installed in a wing main body mounting groove, the wing main body comprises a wing skeleton and a wing outer skin structure outside the wing skeleton; the wing skeleton comprises a longitudinal member unit and a transverse member unit, the longitudinal member unit comprises a front beam, a middle beam and a rear beam, the front beam and the middle beam and the middle beam and the rear beam are arranged at acute angles, the transverse member unit comprises a plurality of wing ribs arranged between the front beam and the middle beam, between the middle beam and the rear beam and on one side of the front beam towards the front of the fuselage; the front beam, the middle beam and the rear beam are all provided with a spar connecting head for connecting with the fuselage. The wing provided by the application adopts a three-beam structure, each wing beam has a mouth-shaped cross section, so that each wing beam forms a multi-wall form, can bear the bending moment and the torque brought by a large wing effective load, and the wing structure is light in weight, high in strength and good in integrity.

A method for manufacturing a structural element (2) of a composite material is described, the structural element extending in a straight or curved longitudinal direction (D), having a Z-shaped cross section and comprising a central web (2a) and two flanges (2b, 2c) in its final configuration, the two flanges extending from opposite ends of the web (2a) in their respective opposite directions at a given final angle; the method comprises the steps of: arranging a plurality of layers (3) of the composite material on a forming portion (4) of a molding die; laminating the layers (3) onto the forming portion (4) such that the web (2a) is at least partially arranged in its final configuration, such that a first flange (2b) is positioned relative to the web in the final configuration of the web. (2a) is arranged at a final angle to form the final configuration of the first flange, and the second flange (2c) is arranged at an initial angle different from and greater than the final angle relative to the portion of the web (2a) arranged as the final configuration of the web, to form the initial configuration of the second flange; the movable part (6) of the molding die is moved from a rest position to a bent position, the movable part being movable relative to the fixed part (5) of the molding die; the second flange (2c) is shifted from the initial configuration to the final configuration, in which the second flange is at the final angle relative to the portion of the web (2a) arranged as the final configuration of the web; the shifting step (d) is performed by the shifting step (c).

A wing for an aircraft include battery modules a plurality of ribs that are spaced apart from one another in a longitudinal direction of the wing, the plurality of ribs respectively defining a plurality of first through-holes through which the battery modules pass, a plurality of first flange portions that are respectively disposed at the plurality of ribs along edges of the first through-holes, the plurality of first flange portions defining battery seating surfaces that support the battery modules, and a spar that supports the plurality of ribs.

The utility model relates to the technical field of gliders, in particular to a carbon beam composite material wing structure. The airfoil comprises an airfoil main beam and foam layers located at the two ends of the airfoil main beam, the airfoil is of a structure which is thick in the middle and thin at the two ends on the whole, the airfoil main beam and the foam layers are wrapped with a layer of airfoil skin, and the airfoil main beam and the airfoil handle are connected in an inserted mode and fastened through a binder and a bolt; a weight reduction structure is designed in a non-bearing area of the wing handle, so that the overall weight of the structure is reduced; the modulus of the high-modulus carbon fiber main beam can reach 200Gpa, so that the deformation in a bearing state is effectively controlled, and the structural modal of the airfoil is improved; the high-modulus carbon beams and the wing handles are connected in an inserted mode, bonded through epoxy glue and connected in a fastened mode through bolts, and the bearing force of the wing surfaces is effectively transmitted to wing handle connecting points; the outer skin of the airfoil is made of high-strength carbon fiber one-way material, the laying direction and number of fibers are designed according to calculated stress, the 0-degree fiber direction, the 90-degree fiber direction and the + / -45-degree fiber direction are configured, and the torsional strength of the airfoil is improved on the premise that the bearing requirement is met.

The invention provides a multi-stage unfolding and folding device based on multi-cylinder bidirectional movement. The multi-stage unfolding and folding device comprises inner supporting rods, outer supporting rods, inner cylinders and the like. The inflatable wings on the two sides are each provided with a set of inner supporting rods and outer supporting rods, the inner supporting rods and the outer supporting rods are located on the inner sides and the outer sides of the inflatable wings respectively, and the ends of the inner supporting rods and the outer supporting rods are rotationally connected with the aircraft structure. The inner air cylinder is installed on the inner supporting rod, the inner piston rod is installed in the inner air cylinder, and the inner piston rod is pressed to the initial limiting position through the inner spring. Two ends of the inner push rod are respectively connected with the inner piston rod and the outer support rod; an inner cylinder inflation port is connected with an inflation port in the inner support rod; the outer cylinder is connected with a mounting bracket structure in the aircraft, the outer piston rod is mounted in the outer cylinder, and the outer piston rod is pressed to an initial limiting position through the outer spring; the two ends of the T-shaped connector are connected with one ends of the outer push rods on the two sides respectively, the third end in the middle is connected with the outer piston rod, and the other ends of the outer push rods on the two sides are connected with the inner supporting rods on the two sides respectively.

This invention relates to an integrated variable camber wing based on a piezoelectric-driven torsional cubic superstructure, comprising a wing skeleton and a skin. The wing skeleton includes multiple torsional cubic superstructures arrayed along the wing chord direction. Each torsional cubic superstructure is a cube formed by an array of multiple unit-cell torsional superstructures. Adjacent unit-cell torsional superstructures in the wing length and thickness directions share a straight rod, while adjacent unit-cell torsional superstructures in the wing width direction share four Z-shaped piezoelectric rods. Each unit-cell torsional superstructure is a cube formed by an array of four cell units, with adjacent cell units sharing a straight rod. Each cell unit includes Z-shaped piezoelectric rods and straight rods. Two Z-shaped piezoelectric rods are connected to the two ends of two straight rods. Each Z-shaped piezoelectric rod at both ends of the torsional cubic superstructure is connected to the skin. When all the Z-shaped piezoelectric rods of the unit-cell torsional superstructure undergo compressive deformation in the height direction, the unit-cell torsional superstructure undergoes torsional deformation, thus causing the torsional cubic superstructure to undergo torsional deformation. By converting the compressive deformation of the cellular unit into the torsional deformation of the compressive-torsional cubic superstructure, the wing camber is changed, and the drive components are eliminated, thus realizing the lightweight design of the morphing wing.

The present invention relates to a split rib for securing a leading edge assembly. The rib includes a first rib portion having a first rear mounting portion for attaching the first rib portion to a wing box assembly and further having a front mounting portion opposite the first rear mounting portion. The rib further includes a second rib portion having a second rear mounting portion for attaching the second rib portion to the front mounting portion of the first rib portion.

The invention belongs to the technical field of aircraft structure design, and particularly relates to a winglet structure. A winglet upper wall plate is installed on the upper side of a wing box section butt joint rib, a winglet upper wing surface covering cap is arranged at the position, close to a wing box section, of the winglet upper wall plate, a winglet lower wall plate is installed on the lower side of the wing box section butt joint rib, and a winglet front beam and a wing box section front edge partition plate are installed on the front side of the wing box section butt joint rib. The winglet rear beam and the wing box section rear edge partition plate are mounted on the rear side of the wing box section butt joint rib; the winglet front edge partition plate is connected with the winglet front beam, the front edge skin is connected with the winglet front edge partition plate, the winglet front beam and the wing box section front edge partition plate, and a front edge covering cap is arranged at the position, close to the wing box section, of the front edge skin. The winglet rear edge comprises a winglet rear edge partition plate, and the winglet rear edge partition plate is connected with the winglet rear beam; the winglet tip is located at the position away from the wing box section and connected with the winglet upper wall plate, the winglet lower wall plate and the leading edge skin.