208results 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.

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.

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.

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.

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 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.

Systems, devices, and methods including one or more rib mounting flanges, where each rib mounting flange comprises: a spar opening configured to receive a main spar of a wing panel; one or more holes for receiving cross-bracing cables; and one or more holes for receiving cross-bracing cables; and one or more holes for connecting the rib mounting flange to an adjacent rib mounting flange.

The application discloses a two-stage rotating-open wing driven by a metal air bag and relates to the field of aviation technology, which solves the problems of heavy driving and transmission device of a deformation wing, low deformation and retraction ratio of a wing surface and poor continuity of a multi-stage deformation wing surface aerodynamic shape. The aircraft base body is provided with a first-stage rotating-open wing, a second-stage rotating-open wing and an air path control system. The leading edge end of the first-stage rotating-open wing is hinged to the aircraft base body, a first-stage air bag is arranged at the hinged position, the first-stage air bag drives the first-stage rotating-open wing to unfold, the rear spar of the first-stage rotating-open wing is movably connected to the second-stage rotating-open wing, the leading edge end of the second-stage rotating-open wing is hinged to a sliding block, the sliding block can slide on a sliding rail fixed to the aircraft base body, a second-stage air bag is arranged at the connecting position of the two, and the second-stage air bag drives the second-stage rotating-open wing to unfold. The application has simple structure, small mass and space occupation of the aircraft, high output power volume ratio, fast response and easy control.

A method of stiffening the integral ribs of a panel to prevent crippling of the ribs during forming of the panel. The method provides that anti-crippling stiffeners are attached to the ribs, such as by using clamps. Then, the panel is formed while the anti-crippling stiffeners are attached to the ribs. The anti-crippling stiffeners add stiffness to the ribs and prevent the ribs from crippling during forming of the panel. The anti-crippling stiffeners could be C-shaped and be attached to the panel using C-shaped clamps. Both the panel and the anti-crippling stiffeners could be formed of aluminum. Another embodiment involves clamping at least one anti-crippling strap to the flange of a wing stringer to prevent the flange from crippling during forming of the wing stringer.

The utility model discloses a detachable wing structure of a light sports aircraft, which comprises an aircraft body, a mounting plate, wings and a fixing piece, and slots which are symmetrically arranged are formed in wing tables of the aircraft body; the mounting plate is fixed between the opposite side walls of the two wing tables; a plurality of first through holes are formed in the plate surface of the mounting plate; the middle part of the wing root rib extends outwards and is fixedly provided with a main beam; the main beam is provided with a plurality of second through holes corresponding to the first through holes; the two main beams located on the two sides of the machine body are oppositely inserted into the inserting grooves. And the main beam and the mounting plate are fastened by the fixing piece. The left wing and the right wing on the two sides of the fuselage are of independent structures, the wings and the fuselage are integrated in an opposite insertion mode, on the premise that the structural strength of the aircraft is not lost, the disassembly and assembly portability of the wings is improved, the connecting mode is simple in structure, high in stability and convenient and fast to operate, and the problems that the aircraft wings are not easy to disassemble or cannot be disassembled and assembled, and the aircraft is inconvenient to disassemble and assemble are solved. The transportation cost and the storage cost are too high.

A flip-and-swing multi-stage deformable wing suitable for cross-domain flight belongs to the field of aircraft technology. The present invention solves the problem that the existing cross-domain flight deformable wing technology cannot simultaneously solve the problems of heavy driving and transmission devices, low wing surface deformation expansion and contraction ratio, and poor continuity of the aerodynamic shape of the multi-stage deformable wing surface. It includes an aircraft base and a deformable wing skeleton, and the aircraft base and the deformable wing skeleton are connected by a flip assembly. The deformable wing skeleton includes a bottom beam, a first-level wing leading edge, a first-level folding assembly, a second-level wing leading edge and a second-level folding assembly. The front end of the first-level wing leading edge is hinged to the front end of the bottom beam. Based on the geometric layout of the folding assembly combined with the wing leading edge and the idea of ​​graded arrangement and continuous expansion of the second-level swinging wing, a second-level swinging wing surface structure is proposed. The deformable wing has aerodynamic shapes under four different flight conditions: fully folded wing, wing surface flipping, first-level swinging wing, and second-level swinging wing, and is suitable for cross-domain flight.

The invention relates to the field of aerospace, in particular to a telescopic composite wing which can be used for aerospace equipment such as hovercars and carrier-based aircrafts. The telescopic composite wing comprises a wing framework, a wing shell and a linear actuator, and further comprises a hidden propeller mechanism, the telescopic composite wing disclosed by the invention has the following beneficial effects that the shape and the size of the wing can be customized according to requirements; in a contraction state, the wings can be completely hidden, so that a limited space is reasonably utilized, and the whole wings are lighter than traditional wings; in the aspect of structural strength, the wing shell provides secondary bending torsion, the framework provides main bending torsion, and the framework is connected with the wing shell at each section and can synchronously stretch and retract, so that the synchronization degree is high, and the stability is strong; the rotary propellers are installed in the wing sections at the two ends of the wings, and the left and right lift force can be adjusted under the condition that the wings on the two sides are equal in length, so that the aircraft can achieve flight actions such as rolling, and the flight attitude controllability is high.

The invention relates to an aircraft assembly (1) comprising an upper fitting (104c) fixed to a wing structure (104a), a lower fitting (104d) fixed to a wing underside panel, an engine pylon (106) with two side panels hinged to the lower fitting (104d), a starboard connecting rod (252a) and a port connecting rod (252b) hinged between the side panel (208a-b) on the same side and the upper fitting (104c), and a mounting block having a proximal end integral with a rear rib of the engine pylon (106) and a distal end mounted in a window of the lower fitting (104d) via an annular linear connection. With such an assembly, the occurrence of out-of-plane forces is limited.

An aircraft wing including a skin and a framework for supporting the skin. The framework includes a spar extending along a length of the wing, the length extending between a root end and a tip end of the wing, a plurality of ribs extending chordwise between a leading edge of the wing and a trailing edge of the wing; and a plurality of stringers attached to the skin. The plurality of stringers includes a first stringer extending along the length of the wing, wherein a proximal end of the first stringer is arranged nearer the root end and a distal end of the stringer is arranged nearer the tip end; and an auxiliary stringer extending non-parallel with the first stringer and arranged such that a portion of the auxiliary stringer is proximal to the distal end of the first stringer.

A method of forming a structural component is provided. The method includes forming, by arranging at least one composite material to a first determined configuration, a first uncured composite laminate beam structure. The method also includes forming, by arranging at least one composite material to a second determined configuration, a second uncured composite laminate beam structure. The method further includes co-curing a first surface of the first uncured composite laminate beam structure to a first surface of the second uncured composite laminate beam structure to form a composite laminate beam structure, where the composite laminate beam structure is a single continuous structure. The method additionally includes coupling a surface of the composite laminate beam structure to an internal surface of a skin.