44 results about "Wing root" patented technology

The application discloses a kind of safety evaluation methods of wind turbine generator set, it is related to wind power generator technical field, including the monitoring data of generator blade, the monitoring data includes wind speed, turbulence intensity and temperature, the edge distribution of the monitoring data is coupled by mixed Copula function to the fitting of the monitoring data;According to the shape parameter of generator set, establish three-dimensional model, carry out a variety of scene simulation by analysis software, generate the maximum bending moment of wing root under different scenes, obtain the mapping relationship of scene-bending moment;Based on the mapping relationship of scene-bending moment.The method can more accurately capture the influence of environmental changes on blade bending moment by combining real-time monitoring data such as wind speed, turbulence intensity, temperature and aerodynamic principles, thereby significantly improving the accuracy of safety evaluation.

The invention provides a dual-motor-driven ornithopter, and belongs to the technical field of aircrafts, the dual-motor-driven ornithopter comprises a dihedral angle control mechanism, a wing root angle control mechanism is arranged at the bottom of the dihedral angle control mechanism, and flapping wing mechanisms are symmetrically distributed at the top of the dihedral angle control mechanism; the flapping wing mechanism comprises a fixing frame, a driving assembly and a transmission assembly, and the transmission assembly can drive the first wing and the second wing to symmetrically move under the driving of the driving assembly. According to the scheme, the flapping-wing air vehicle driven by the double motors adopts an X wing type, the lift force is larger than that of a traditional flapping-wing air vehicle, and the energy utilization rate is higher; through cooperation of the dihedral angle control mechanism, the wing root angle control mechanism and the double motors, many high-maneuverability actions can be achieved.

The invention discloses a shape-preserving and resistance-reducing design method for a rotating mechanism under rotor / fuselage coupling, which comprises the following steps of: setting the slenderness ratio of a streamline shaft sleeve based on a shape resistance coefficient formula, and verifying the surface pressure distribution and the separation area of the shaft sleeve through numerical simulation to ensure that the shape resistance coefficient meets a resistance-reducing target; setting the area ratio of an outlet to an inlet of the seam based on a Venturi effect formula, so that a low-pressure area is formed in the seam to suck a wing root boundary layer and delay airflow separation; establishing an aerodynamic load and motion parameter coupling model based on an instantaneous resistance formula, analyzing flow fields under different rotation angular velocities and oblique angles through numerical simulation, and optimizing a rotation track to reduce a resistance peak value and aerodynamic torque fluctuation; the invention aims to solve the problems of airflow vortex interference, large shape resistance, high resistance peak and poor aerodynamic stability by optimizing the shape of the rotating shaft sleeve, the fuselage-wing gap structure and the wing rotating track, and improves the cruise fuel efficiency and flight stability of the aircraft.

The invention discloses a high-aerodynamic-efficiency fusion type wingtip winglet. The wingtip winglet comprises a fusion wing root section, a torsion main section and a wingtip rectification section which are connected in sequence; the wingtip winglet adopts a spanwise variable thickness modified NACA laminar flow airfoil, the relative thickness of a wing root is 12%-15%, the relative thickness of a wingtip is 6%-8%, the wingtip winglet is linearly and gradually changed in the spanwise direction, the radius of the front edge of the wingtip winglet is 0.8%-1.2% of chord length, the rear edge of the wingtip winglet is in forward pressure gradient transition, and the maximum thickness of the wingtip winglet is located at the position of 35%-40% of the chord length; the wingtip winglet main body is of a multi-wall multi-closed-chamber integrated bearing structure and comprises a front beam, a rear beam, an upper skin and a lower skin, and at least three main closed chambers are formed; the high-aerodynamic-efficiency fusion type wingtip winglet has the advantages of being excellent in aerodynamic performance, light in weight, fast to install and adjust and high in universality, and is suitable for civil airliners, transport aircrafts and high-end fixed-wing unmanned aerial vehicles.

The invention provides an unmanned aerial vehicle dot matrix skin fixed wing and a metal laser additive manufacturing method thereof, and belongs to the technical field of unmanned aerial vehicle structure design and additive manufacturing, the unmanned aerial vehicle dot matrix skin fixed wing comprises an integrally formed three-dimensional gradient dot matrix skin, and the dot matrix unit density of the dot matrix skin is non-uniformly distributed based on the stress state of the wing; the non-uniform distribution is determined by performing orthogonal composite division on the inner space of the wing along the radial direction and the axial direction: the radial division is that the inner part of the wing is divided into a radial core low-density filling area, a radial middle medium-density filling area and a radial outer high-density filling area based on the distance from the position to the normal direction of the skin; according to axial division, the interior of the wing is divided into an axial root low-density filling area, an axial middle medium-density filling area and an axial tail high-density filling area based on the axial distance between the position and the root of the wing; and the axial division is realized by adjusting the reference density value of each radial filling area.

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.

The invention discloses a wing of a bionic flapping-wing aircraft capable of inhabiting in a manner of being perpendicular to a wall surface. The wing comprises a connecting component, a supporting framework and a wing membrane, the connecting component is of a four-fork branch supporting arm structure of a forked structure, the tail end of each supporting arm is provided with a long hole used for fixing the supporting framework, and the center boss is provided with a rudder arm connecting hole used for being connected with a driving component. The supporting framework comprises a main force bearing beam, sub-components distributed in a radial mode, transverse fixing wing ribs and rear wing supporting wing ribs distributed in an annular mode. The wing film is adhered to the surface of the supporting framework to form a front wing area and a rear wing area of the wing; wherein the outer contour of the front wing area is constructed to have the curvature radius which is continuously reduced from the wing root to the wing tip, and a variable curvature radius layout is formed; and the hip angle curvature radius of the rear wing area is 50mm. Through structural innovation, the aerodynamic efficiency, the structural stability and the collaboration with inhabiting components of the wing are effectively improved, and the wing is particularly suitable for a bionic flapping-wing aircraft needing to inhabit on a vertical wall surface.

An unmanned aircraft according to the present disclosure is provided with an airframe (2) and a main wing (4). The main wing (4) has: a pair of left and right wing body members (12) that are connected to the airframe (2) at the wing root side, and extend in the wing width direction toward the wing tip side; and at least one movable wing member (14) that is provided at the trailing edge of the wing body members (12), and can be displaced relative to the wing body members (12). The at least one movable wing member (14) has a wing root-side portion disposed at the wing root side, and a wing tip-side portion disposed at the wing tip side. The wing tip-side portion has a smaller lift coefficient than the wing root-side portion.

The application discloses a rear-swept wing distributed non-equal-diameter rotorcraft and a control method thereof, which comprises a fuselage, a wing, a tilt-rotor group, a landing gear, a horizontal tail and a vertical tail; the tilt-rotor group comprises four tilt-rotors which are connected to the leading edge of the wing through a support rod; the rotor disc area near the wing tip is larger than the rotor disc area near the wing root, so as to balance the flight efficiency at low speed and the flight speed at high speed, and the rear sweep of the wing is beneficial to improving the stability of the tilt-rotor aircraft in the fixed-wing mode flight, and the environmental adaptability of the aircraft is improved. The application balances the efficiency of the tilt-rotor aircraft in the hovering and cruising states, the aircraft structure is relatively simple, is beneficial to the attitude control of the aircraft and the flight control software design.

The invention discloses a wing body fusion aircraft and a control method thereof, the aircraft comprises an aircraft body, a left wing and a right wing, and the wing body fusion aircraft is characterized in that the total wingspan length of the aircraft is S; the length L of the machine body is larger than or equal to 0.4 S and smaller than or equal to 0.6 S; the chord length Lr of the wing roots of the ends, connected with the fuselage, of the left wing and the right wing meets the condition that Lr is larger than or equal to 0.2 S and smaller than or equal to 0.4 S; the wing tip chord length Lt of one end, far away from the fuselage, of each of the left wing and the right wing meets the condition that Lt is greater than or equal to 0.1 S and less than or equal to 0.2 S. Through collaborative innovation of pneumatic, power and structure, long-time and low-power-consumption continuous flight of the wing body fusion aircraft is achieved.

The invention discloses an unpowered micro aircraft imitating seeds of a bombycaceae plant, and belongs to the technical field of micro aircrafts. The micro aircraft comprises a load cabin and at least two fruit wings, wherein the fruit wing roots of the at least two fruit wings are uniformly distributed at the top of the load cabin in the circumferential direction; the at least two fruit wings are the same in appearance; a fruit wing center line in the length extending direction of the fruit wing is constructed to be in a continuous or discontinuous bending form which firstly extends upwards in the direction of a central vertical longitudinal axis of the load cabin and then obliquely extends towards the oblique upper portion away from the central vertical longitudinal axis in sequence in the direction from the root of the fruit wing to the wing tip of the fruit wing. Due to the bent shape, stable vortex is formed when airflow flows through the upper surfaces of the fruit wings, lift force is provided for the micro aircraft, the falling speed of the aircraft is reduced in the falling process, and long-time hang flight is achieved. Meanwhile, the device is simple in structure, light in weight and suitable for application scenes such as air-drop delivery, environment detection and special operation.

The invention discloses a method for automatically generating a finite element model of a wing hexahedron element structure, and belongs to the field of multi-physics field coupling numerical simulation. Comprising the steps that a radial basis interpolation model driven by a curve distance is adopted, topologically consistent wing root and wingtip quadrilateral grids are generated, geometric and topological mapping from two dimensions to three dimensions is achieved through control point alignment and spanwise over-limit interpolation, a wing is divided into 10 parameterized blocks, internal nodes are generated through tensor product interpolation, and the internal nodes are divided into 10 parameterized blocks. And combining repetitive nodes based on a spatial hash algorithm, and constructing a globally consistent hexahedral unit topology. According to the method, full hexahedron unit generation can be ensured, and the generation efficiency is remarkably improved compared with that of a manual interactive method.

The invention provides a wing structure with distributed fan ducts, which comprises a wing main body, the wing main body comprises a wing spar extending along the wingspan direction and a wing rib extending from the wing spar to the front-back direction of a fuselage, the wing spar comprises a front beam positioned at the front edge of the wing, a rear beam positioned at the rear edge of the wing and a middle beam between the front beam and the rear beam, the wing ribs comprise an inner side rib close to the wing root, an outer side rib close to the wing tip and a middle rib between the inner side rib and the outer side rib, the inner side rib and the middle beam form a first angle, and the inner side rib extends to the front beam and the rear beam from the middle beam, so that the inner side rib is parallel to the longitudinal axis of the fuselage; the outer side rib extends from the middle beam to the front beam at a second angle with the middle beam, so that the outer side rib is perpendicular to the middle beam; the middle rib extends from the middle beam to the front beam and the rear beam, and the included angle between the middle rib and the middle beam is gradually reduced from a first angle to a second angle. According to the wing structure, the function and structural integrity of the wing main body, the fan duct and other parts can be realized, and the requirements of integration of the wing structure and all systems are met.

The utility model discloses an eddy current array generator of a light moving aircraft, which comprises an aircraft fuselage, two aircraft wings are symmetrically fixed on two sides of the aircraft fuselage, and four groups of eddy current generator components are fixedly mounted on two side surfaces of the aircraft fuselage and close to wing roots of the aircraft wings. The four vortex generator assemblies on the two sides of the aircraft fuselage are symmetrically arranged. According to the utility model, the triangular vortex generator a with the filleted corners, the rectangular vortex generator b with the filleted corners, the rectangular vortex generator c with the filleted corners and the rectangular vortex generator d with the filleted corners are arranged, so that the airflow flowing characteristic at the cross-linking part of the wings and the fuselage of the light moving aircraft is improved, and the airflow separation of the wings is delayed when the aircraft flies at a low speed; stall is delayed, and the stall speed is reduced; pilots and passengers can fly up and down conveniently, and efficiency is improved; the crosswind resistance of the aircraft during low-speed flight is improved, and the transverse course stability is improved.

The utility model belongs to the technical field of unmanned aerial vehicles and provides a wing folding mechanism and an unmanned aerial vehicle. The wing root of the first wing is connected with a first adapter; the wing root of the second wing is connected with a second adapter; the first adapter and the second adapter are rotationally connected and are used for folding the first wing and the second wing; the torsional spring is located between the first adapter and the second adapter, and the two ends of the torsional spring are installed in the first adapter and the second adapter respectively. When the fixed-wing unmanned aerial vehicle is used and the wings need to be folded, the first wing and the second wing are rotated inwards through the rotating shaft, so that the first wing and the second wing are mutually attached and folded to achieve the purpose of convenient transportation, the folded fixed-wing unmanned aerial vehicle can be loaded into a launching cylinder, and when the fixed-wing unmanned aerial vehicle is launched out of the cylinder, the unmanned aerial vehicle can be conveniently transported. The wings are automatically unfolded under the torque action of the torsional springs, and the aim of flexible and convenient launching is achieved.

The application provides a small-sized separated flapping-wing aircraft for reconnaissance, and relates to the technical field of bionic flapping-wing aircrafts, characterized in that the small-sized separated flapping-wing aircraft comprises a modular fuselage, a quick-detachable wing assembly and a detachable tail wing, and the components are connected through a tool-free buckle type quick-assembly structure; the modular fuselage comprises two left-right split housings and a core cabin, the housings are internally provided with mounting slots, and the core cabin is arranged in the mounting slots; the quick-detachable wing assembly is arranged above the modular fuselage, and the detachable tail wing is arranged behind the modular fuselage; the quick-detachable wing assembly comprises a wing, and the wing root is provided with a quick-assembly buckle. The application has the advantages of compact structure, mature manufacturing process, low cost and convenient batch production. The module interfaces are standardized, and the supply chain can be divided for production. The application can be used for military individual reconnaissance and special operations, and is also applicable to the fields of scientific research and teaching, education and outdoor entertainment, and has wide market prospects and practical value.

Wing assembly with slats and aircraft. A wing assembly comprising: a swept wing body, a leading edge of the wing body extending outwardly and rearwardly from a wing root to a wing tip; a first slat selectively movably connected to the wing body; and a second slat selectively movably connected to the wing body, the second slat disposed outwardly of the first slat, a flexible seal member disposed and connected between the first slat and the second slat; at least a portion of the first slat, at least a portion of the second slat, and at least a portion of the flexible seal member defining therebetween a slat gap, at least a majority of the slat gap being substantially parallel to a predetermined local airflow direction. Also disclosed is an aircraft comprising a fuselage; and two oppositely disposed wing assemblies connected to the fuselage.

The utility model discloses a novel folding rudder, relates to the technical field of aircraft structures, and can solve the problems that in the prior art, due to a rotating shaft and the like, an abnormal protruding structure appears on the surface of the folding rudder, and the wind resistance coefficient is high. The novel folding rudder provided by the embodiment of the utility model comprises a wing root, a wing tip and a rotating shaft structure, the rotating shaft structure is transversely arranged in the wing root; the wingtip is connected to the wing root through a rotating shaft structure, and the wingtip can rotate around the central axis of the rotating shaft structure to be in a folded state and an unfolded state; the planes, close to each other and parallel to the rotating shaft, of the wing root and the wing tip are inclined planes; the locking assembly is transversely arranged in the wing root and is used for locking the state of the wing tip.

The application discloses a design scheme of a composite bistable flexible wing, which is composed of a wing vein, a reticular wing vein, a front edge sleeve and a wing root sleeve, and is different from the design scheme of a flexible wing used by a current micro flapping-wing aircraft in that the composite bistable flexible wing is designed through an included angle between the front edge sleeve and the wing root sleeve and the reticular wing vein, is a planar structure before being installed to the micro flapping-wing aircraft, and becomes a spatial curved surface structure with a bistable characteristic after being installed.The composite bistable flexible wing for the micro flapping-wing aircraft has the characteristics of low flapping noise, large lift, high aerodynamic efficiency, difficulty in fatigue failure and high reliability.A preparation method of the composite bistable flexible wing has five steps, and has the advantages of simple operation, good processing consistency and good engineering application value.

The application discloses a bionic dragonfly flapping-wing aircraft with a double sliding sleeve flapping mechanism and belongs to the technical field of aircrafts. The bionic dragonfly flapping-wing aircraft comprises a bottom plate support, a transmission mechanism, a double sliding sleeve flapping mechanism, a cylindrical cam steering sweeping mechanism, one pair of front flapping wings and one pair of rear flapping wings. The transmission mechanism comprises front air source driving transmission mechanisms and rear air source driving transmission mechanisms which are arranged on the front and rear sides of the bottom plate support and are of the same structure. Each flapping wing root is provided with double sliding sleeve flapping mechanisms and cylindrical cam steering sweeping mechanisms which are of the same structure. The front air source driving transmission mechanism drives the circumferential rotation of a first sliding sleeve through an air source rotary vane engine, and then drives the reciprocating rotation of a wing root rotating shaft fixed frame and a wing root rotating shaft through a second sliding sleeve, so as to drive the flapping of the flapping wings. The rotation of the cylindrical cam of the cylindrical cam steering sweeping mechanism can drive the reciprocating sliding of a reciprocating top transmission part, push or pull the reciprocating rotation of a wing root connecting part, and then drive the sweeping movement of the flapping wings in space.

The application discloses a wing quadrilateral element structure finite element model automatic construction method and device, equipment and medium, and relates to the field of aerospace technology, and comprises the following steps: determining third coordinate data of a wing middle section based on target control parameters, first coordinate data of a wing root section and second coordinate data of a wing tip section; constructing an airfoil geometric parameterized model based on the first coordinate data, the second coordinate data and the third coordinate data and by using a preset interpolation method, generating target grid control points and target grid control edges corresponding to the wing root section, the wing tip section and the wing middle section respectively, determining a first grid control edge connecting the wing root section and the wing middle section, and a second grid control edge connecting the wing middle section and the wing tip section; and determining a wing quadrilateral element structure finite element model file based on a first quadrilateral grid of a skin, a second quadrilateral grid of a wing rib and a third quadrilateral grid of a wing spar. The method is used for efficiently constructing a high-fidelity structural finite element model.

The application provides a low-subsonic to high-subsonic wide-speed-range deformable aerodynamic layout of a cruise missile, comprising double-axis variable sweep wing mechanisms arranged at the bottom of a fuselage, capable of arbitrarily changing and locking a sweep angle in a range of 0-90 degrees, each double-axis variable sweep wing mechanism comprising an inner wing segment and an outer wing segment, the inner wing segment being pivotally connected with the fuselage at a wing root, and the inner wing segment being adjusted at a front sweep angle in a range of 0-90 degrees; the outer wing segment being pivotally connected with the inner wing segment at a wing tip, and the outer wing segment being adjusted at a rear sweep angle in a range of 0-90 degrees; and the nose being blunt. In the embodiment, the double-axis variable sweep wing mechanism has opposite influence trends of the inner wing segment and the outer wing segment on the center of pressure of the cruise missile during stretching, folding and flying, so as to control the change range of the center of pressure of the cruise missile in a very small interval, and make the wing sweep process have good maneuvering and stability characteristics and flight stability. In addition, the double-axis variable sweep wing mechanism is parallel to the fuselage when being folded, so that the cruise missile is convenient to store and transport, and the storage and deployment capability is improved.

The present disclosure provides a wing-sail type vortex generator for improving the rudder effect of a flying wing at a large attack angle, the wing of the flying wing comprising a leading edge, a wing root arranged with a body protruding surface, and a trailing edge arranged with a deflected flap, the leading edge, the wing root and the trailing edge being sequentially connected, the wing-sail type vortex generator comprising a wing-sail type vortex generator body, the wing-sail type vortex generator body being arranged on the leeward surface of the wing and close to the leading edge; wherein when at a large attack angle, the wing-sail type vortex generator body is used to combine with the protruding surface of the wing root to form a local area of a close-to-leeward surface partial nozzle, so that the local flow field of the deflected flap surface of the trailing edge changes from a separation state to an attached state. In the wing-sail type vortex generator for improving the rudder effect of a flying wing at a large attack angle, the arrangement of the wing-sail type vortex generator body on the wing can effectively improve the rudder effect of the flap and achieve the effect of reducing the lifting moment.

The invention discloses an aircraft wing body quick connecting device, and relates to the field of aircrafts, the aircraft wing body quick connecting device comprises an aircraft body and wings mounted on the two sides of the aircraft body, and insertion rods are fixedly connected to the root positions of the wings; inserting barrels are fixedly connected to the positions, in butt joint with the wings, of the aircraft body, and the inserting rods are matched with the inserting barrels in an inserting mode; the opening area of the insertion cylinder is larger than that of the insertion end of the insertion rod; a clamping assembly is arranged in the machine body and used for automatically clamping the inserting rod into the inserting barrel when the inserting rod is connected with the inserting barrel in an inserted mode. Positioning assemblies are arranged in the wings, and the positioning assemblies are used for fixing the wings at a preset angle relative to the aircraft body after the insertion rods and the insertion barrels are connected in an inserted mode; the wing can be quickly mounted and dismounted, accurate butt joint is not needed during mounting, and operation is convenient.

The invention discloses a foldable wing for a small underwater glider and the underwater glider, and belongs to the technical field of underwater detectors. The core of the invention lies in that a split type folding and unfolding structure is adopted, the wing fixing assembly comprises a fixed rigid wing root part, the wing is used as a movable part, and in an unfolding state, the root part of the wing and a clamping groove of the wing root part are embedded and butted, and are jointly spliced to form a complete rigid wing body. The folding and unfolding mechanism drives gear and rack transmission through spring tension, and automatic unfolding of the wings is achieved; and the limiting mechanism is configured to automatically pop out the pin shaft for mechanical locking after the wings are unfolded in place. The size of the glider is effectively reduced so as to adapt to multi-platform hidden deployment, the underwater impact resistance is improved, and through double guarantee of fixed wing root and mechanical locking, it is ensured that the wings have the same structural strength and underwater navigation stability as traditional fixed wings after being unfolded.

The sub-span and super-span speed domain aerodynamic layout comprises double-shaft variable-sweep-wing mechanisms arranged at the bottom of a fuselage, the sweep angle can be freely changed and locked within the range of 0-90 degrees, each double-shaft variable-sweep-wing mechanism comprises an inner wing section and an outer wing section, the wing root of each inner wing section is in pivot connection with the fuselage, and the wing root of each outer wing section is in pivot connection with the fuselage. The sweep-forward angle of the inner wing section is adjusted within any angle of 0-90 degrees; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, and the sweepback angle of the outer wing section is adjusted at any angle of 0-90 degrees; the machine body is in a flat shuttle shape, and the machine head is in a pointed-end shape. According to the implementation mode, in the stretching, folding and flying processes, the influence trends of the double-shaft variable-swept-wing mechanism on the pressure center focus of the patrolling bomb are opposite, so that the pressure center focus change range of cross-speed domain deformation flight is controlled within an extremely small interval, and the pneumatic control stability is remarkably improved. In addition, when being stored, the double-shaft variable-swept-wing mechanism can be completely folded and parallel to the fuselage and is adaptive to narrow spaces such as a launch canister or a transport case, so that the storage and deployment capabilities are improved.

The utility model relates to the technical field of airplane emergency floating, in particular to an emergency floating device for a light airplane. The device comprises a shell, a stop cover and a cover, wing sliding rails are fixed to the two ends of the bottom of the shell, and the stop cover is fixed to one end of each wing sliding rail. According to the utility model, the cover can be quickly unfolded, the disassembly and assembly are very convenient, and the configuration can be quickly adjusted according to task requirements; the shell is installed at the root of a light aircraft wing or the two sides of a helicopter body, streamline design is adopted in the appearance, separation of local airflow is reduced, the installation process is simple and convenient, the internal structure of the aircraft does not need to be transformed, the installation time is greatly shortened, and the installation cost is greatly reduced; the inflation speed of the folding air bag can be obviously increased, the folding air bag can reach the needed pressure in a shorter time, and in addition, the situation that the whole system fails due to the fault of a single air cylinder is avoided.

The utility model belongs to the technical field of unmanned aerial vehicle wing structures, and particularly relates to an unmanned aerial vehicle wing structure with flexible connection. Comprising a wing flexible connecting structure, a wing root connecting piece, a winglet and a winglet connecting rod, the wing flexible connecting structure comprises a wing upper skin, a main wing lower skin, a flap and a flexible connecting piece; the flap side of the main wing upper skin is connected with the flap through a flexible connecting piece; one end of the main wing upper skin and one end of the main wing lower skin are butted and spliced through a wing root connecting piece to form a main wing; the interior of the main wing is a cavity; the winglet sleeves the other end of the main wing upper skin and the other end of the main wing lower skin, and is connected with the wing root connecting piece through the winglet connecting rod; the unmanned aerial vehicle body is connected with the wing root connecting pieces through bolts. The unmanned aerial vehicle has the advantages that the weight of the main wings is effectively reduced, unmanned aerial vehicle fuel or batteries can be stored in the container cavities in the main wings, and the endurance mileage is increased.

A multi-stage ionic thruster includes a voltage supply, a frame of a wing, the frame defining a first axis between a leading edge of the wing and a trailing edge of the wing and a second axis between a root of the wing and a tip of the wing, and a plurality of ionic thrusters coupled to the frame and spaced apart from each other along the first axis.