72 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 application discloses a design method of a ducted propeller, comprising the following steps: S1, clearly defining design conditions and initial design parameters; the design conditions comprise tension T generated by the duct and the propeller, and working speed n of the propeller; the initial design parameters comprise propeller diameter D, propeller-duct clearance δ and airfoil; S2, designing the propeller diameter D; S3, designing the propeller-duct clearance δ; S4, designing the airfoil; the prerequisite of parameterization of the propeller blade is to determine an airfoil family used by the propeller blade and a radial position, and to keep constant the chord lengths of a wing root, a middle part of the wing and a wing tip, and the duct; under the prerequisite, the airfoils of the wing root, the middle part of the wing and the wing tip section are determined through a CST type function; except the wing root and the wing tip, the negative torsion angles of 3-7 sections of the radial position are simultaneously taken as optimization variables, and a parameterization model of the propeller blade is generated. The method has the advantages that the iterative optimization design method is based on numerical simulation, can accurately evaluate and improve the aerodynamic efficiency of the ducted propeller.

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 combined wing and an aircraft, and belongs to the technical field of aircrafts. Comprising a trapezoidal wing, a multi-section wing and a spiral winglet, and the trapezoidal wing, the multi-section wing and the spiral winglet are sequentially connected in the direction from a wing root to a wingtip; the multi-section wing comprises a front-end wing and a rear-end wing, a wing gap is formed between the rear edge of the front-end wing and the front edge of the rear-end wing, and the attack angle of the rear-end wing is larger than that of the front-end wing; one end of the spiral winglet is smoothly connected with the wingtip of the front-end wing, the other end of the spiral winglet is smoothly connected with the wingtip of the rear-end wing, and the geometric center of the spiral winglet is located on the axis extension line of the trapezoidal wing; the spiral winglet is smoothly mounted at the wingtip position of the multi-section wing, so that the vortex of the wingtip of the trapezoidal wing can be effectively weakened, and the induced resistance is greatly reduced.

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 application belongs to the field of flight control technology, and particularly relates to a stall angle of attack determination method and device for a small aspect ratio wing. The method comprises the following steps: S1, interpolating a first correction coefficient C1 and a second correction coefficient C2 in a first interpolation table and a second interpolation table respectively according to a wing root ratio; S2, determining a first interpolation parameter CAB1 according to the first correction coefficient and determining a second interpolation parameter CAB2 according to the second correction coefficient; S3, determining a basic value of the stall angle of attack of the wing in a third interpolation table according to the first interpolation parameter CAB1; S4, determining a stall angle of attack correction amount in a fourth interpolation table according to the second interpolation parameter CAB2; and S5, determining the stall angle of attack according to the basic value of the stall angle of attack of the wing and the stall angle of attack correction amount. The application has high calculation efficiency and controllable calculation precision for the stall angle of attack of the small aspect ratio wing.

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 quadrilateral unit structure finite element model automatic construction method and device, equipment and a medium, and relates to the technical field of aerospace, and the method comprises the steps: determining third coordinate data of a wing middle section based on a target control parameter, first coordinate data of a wing root section and second coordinate data of a wing tip section; based on the first coordinate data, the second coordinate data and the third coordinate data, constructing an airfoil geometric parameterized model by utilizing 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, and determining a first grid control edge connecting the wing root section and the wing middle section, the second grid control edge is connected with the wing middle section and the wing tip section; and determining a wing quadrilateral unit structure finite element model file based on the first quadrilateral mesh of the skin, the second quadrilateral mesh of the wing rib and the third quadrilateral mesh of the wing spar. And a high-fidelity structural finite element model is efficiently constructed.

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.

The invention discloses a piezoelectric-driven miniature imitated butterfly robot and a motion control method thereof, and the miniature imitated butterfly robot is characterized in that the miniature imitated butterfly robot comprises a machine body skeleton, and piezoelectric drivers and flexible wings which are arranged on two sides of the machine body skeleton; the flexible wings comprise a flexible main wing and a flexible auxiliary wing, the first end of the piezoelectric actuator is connected with the fuselage framework, and the wing root of the flexible main wing is connected with the second end of the piezoelectric actuator through a displacement amplification mechanism; the wing root of the flexible aileron is connected with the wing root of the flexible main wing, so that a phase delay angle is formed between the flexible main wing and the flexible aileron in the upward flapping process of the flexible wing. The flexible wing is applied to the field of micro bionic robots, the flexible wing comprises the flexible main wing and the flexible aileron, so that a phase delay angle is formed between the flexible main wing and the flexible aileron in the flapping process of the flexible wing, the optimal aerodynamic effect is achieved, the lifting force can be maximized, the cost of needed materials is low, manufacturing and assembling are easy, and the flexible wing can be widely applied to the field of micro bionic robots. And large-batch and low-cost manufacturing can be realized.

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 invention belongs to the technical field of aviation aircrafts, and particularly discloses a multi-dimensional continuous variant wing structure which comprises a fixed plate, an inner wing section, an outer wing section, a sweepback actuating mechanism and a telescopic actuating mechanism, the inner wing section is driven by a sweepback actuating mechanism to rotate around the wing root in a spanwise plane so as to change the sweepback angle of the wing; the outer wing section is slidably connected to the inner wing section and is driven by a telescopic actuating mechanism to stretch in the spanwise direction so as to change the extension length of the wing; the front end and the tail end of the rear side of the inner wing section are connected with the flexible rear edge and the rigid rear edge correspondingly, and the flexible rear edge is used for changing the bending degree of the rear edge. The rigid trailing edge is used for changing the trailing edge deflection angle; the sweepback actuating mechanism comprises a rotating disc, a push rod motor and a rocker arm, and the push rod motor drives the rocker arm to enable the inner wing section to rotate around the rotating disc to achieve sweepback changing. According to the invention, the multi-dimensional variation of the wings can be realized through simple mechanism combination, and the aerodynamic performance of the aircraft is improved, so that the aircraft can adapt to different flight conditions, and various flight task requirements are realized.

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 present application belongs to the technical field of flapping-wing aircraft, and particularly relates to a micro flapping-wing aircraft, which comprises a flapping mechanism, an attitude adjusting mechanism, a wing, an overall frame and a controller for controlling the flight of the micro flapping-wing aircraft, the flapping mechanism is fixed at the top end of the overall frame, the attitude adjusting mechanism is connected to the bottom of the flapping mechanism and fixed on the overall frame, and the wing is fixed through the flapping mechanism and the attitude adjusting mechanism; the flapping mechanism comprises a frame, a transmission structure and a speed change structure, the frame is installed at the top end of the overall frame, the speed change structure is fixed on the upper surface of the frame, and the transmission structure is connected to the speed change structure and fixed on the lower surface of the frame; the wing comprises two front edge rods, a plurality of wing veins, a wing membrane and two wing root shafts; the micro flapping-wing aircraft overall architecture is reasonably designed, the center of gravity is approximately located at the average aerodynamic center, the horizontal and vertical force arms from gravity are almost zero, the parasitic moment is reduced, and therefore the aircraft has the advantages of stable flight attitude, long flight time and the like.

This invention belongs to the field of aircraft technology, specifically disclosing a multi-dimensional continuously variable wing structure, including a fixed plate, an inner wing section, an outer wing section, a sweep actuation mechanism, and a telescopic actuation mechanism. The inner wing section is driven by the sweep actuation mechanism to rotate around the wing root in the spanwise plane to change the wing sweep angle. The outer wing section is slidably connected to the inner wing section and is driven by the telescopic actuation mechanism to extend or retract along the spanwise direction to change the wing span. The leading and trailing edges of the rear side of the inner wing section are respectively connected to a flexible trailing edge and a rigid trailing edge. The flexible trailing edge is used to change the trailing edge camber; the rigid trailing edge is used to change the trailing edge deflection angle. The sweep actuation mechanism includes a turntable, a push rod motor, and a rocker arm. The push rod motor drives the rocker arm to rotate the inner wing section around the turntable to achieve variable sweep. This invention can achieve multi-dimensional wing variability through a simple combination of mechanisms, improving the aerodynamic performance of the aircraft and enabling the aircraft to adapt to different flight conditions and meet various flight mission requirements.

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 present application relates to a kind of micro bionic flapping wing aircraft based on gear pair amplification mechanism, including frame, wing and tail, the frame is connected with tail by carbon rod framework, and the number of wing is provided with two, and is connected to the two sides of carbon rod framework by wing root.This application relates to the technical field of bionic aircraft.The aircraft uses crank rocker mechanism to convert hollow cup motor rotary motion into reciprocating rotation of double-layer gear, further transmission is the reciprocating rotation of left wing pinion and right wing pinion larger amplitude, the aircraft has the characteristics of compact structure, small size, light weight, easy to assemble, high reliability etc.The aircraft uses the combination of wing root and tail control, and gives consideration to the efficient use of lift and the miniaturization design of aircraft.

This invention discloses a structural scheme for actively altering the torsional shape of an airfoil, belonging to the field of variator aircraft structural design technology. Specifically, when the drive device rotates the gear screw, the gear screws on the upper and lower flanges generate axial forces in opposite directions, forming a couple that causes the wing spars of the torsional deformation section to deflect around the hinge point. The front and rear wing spars deflect upward and downward around the hinge point respectively, while the skin on the upper and lower surfaces of the airfoil restricts the distance between the front and rear spars, thus causing torsional deformation of the airfoil. The torsional deformation angle is measured by a cable displacement sensor, which measures the distance the gear screw on the wing spar moves relative to the nut after deformation. The deflection angles of the front and rear spars are calculated based on the height of the spar web, and then combined with the chordal distance between the front and rear spars to obtain the torsional shape of the airfoil. The structure for actively altering the torsional shape of the airfoil proposed in this invention effectively improves the roll maneuverability of the aircraft, improves the distribution of aerodynamic loads, and reduces the load at the wing root.

The invention provides a small low-cost three-channel steering engine room mechanism. The steering engine room mechanism comprises a first steering engine room assembly, a gear transmission assembly, a second steering engine room assembly and a rudder piece. The first steering engine room assembly, the gear transmission assembly and the second steering engine room assembly are arranged in a stacked mode from top to bottom, the whole steering engine room mechanism is cylindrical, and the first steering engine room assembly, the gear transmission assembly and the second steering engine room assembly are sequentially arranged from top to bottom. Four rudder sheets which are arranged in a cross shape are uniformly distributed on the periphery of the steering engine room mechanism; the first steering engine room assembly and the second steering engine room assembly are the same in structure. The first steering engine room assembly, the gear transmission assembly and the second steering engine room assembly are fixedly connected through circumferential screws. The rudder sheets are circumferentially and annularly distributed along the outer contour of the middle base assembly and fixedly connected with the exposed rudder sheet shaft installation position on the middle base assembly through screws near the wing root.

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 utility model relates to a kind of flow improving devices at wing root side of front edge slat wing root, comprising: front edge slat, by support, rack, gear mechanism and wing leading edge can be retracted connection;Slat channel, be opened in the lower skin of the inner surface of the near wing root cavity of front edge slat, width than baffle joint pillar diameter 1-2mm;Slat side edge baffle, through baffle joint pillar wear slat channel;Baffle storage bin track, welding between two baffle storage bin wallboard, width than slat side edge baffle thickness 2-4mm;Baffle curved protrusion, fixed in the lower side of slat side edge baffle rear, with baffle storage bin track sliding contact;Sealing element is set to the profile of slat side edge baffle top and baffle storage bin opening edge.The utility model, using movable baffle and front edge slat coordinated motion, break through the design limit of fixed isolation piece to slat side edge, realize the drag reduction of downflow layout, cooperate sealing element to reduce airflow interference, increase stall angle of attack and maximum lift coefficient, improve aerodynamic performance.

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 relates to the technical field of aircraft design, in particular to a battery and wing integrated structure which comprises a wing structure, a battery module heat management structure and a battery module, the wing structure comprises a wing front beam, a wing middle beam, a wing rear beam, a wing root wing rib, a wing inner box section wing rib, a wingtip wing rib, a laminated battery cabin cover and a cooling fan; the wing front beam, the wing middle beam and the wing rear beam are arranged in parallel in the wing spanwise direction; the wing ribs of the wing inner box section respectively surround the wing front beam, the wing middle beam and the wing rear beam to form a battery cabin, and the battery module heat management structure and the battery module are fixed in the battery cabin; the wing structure is provided with a battery compartment opening, and the laminated battery compartment cover is matched with the battery compartment opening; the wing root wing ribs and the wing tip wing ribs are connected with cooling fans; the weight of the wing can be reduced, and heat dissipation of the battery module is guaranteed.

The invention belongs to the technical field of airfoil design, and discloses a C-shaped beam design method, a C-shaped beam and an airfoil structure. The cross section of the C-shaped beam is equal in thickness and is gradually thinned in the spanwise direction, namely, the cross section is gradually thinned in the direction from the wing root to the wing tip. The inner molded surface of the web is a plane, and the outer molded surface of the web is provided with a web lost layer; the extension direction of the edge strip is the same as that of the skin, and an edge strip lost layer is arranged on the inner molded surface of the edge strip, namely, the web lost layer and the edge strip lost layer are arranged on different sides. The material sheets do not need to be overturned after being paved by the automatic tape paving machine and are directly transferred to the pre-forming tool to be formed, and the risk that fibers are bent or deformed in the material sheet overturning process can be effectively avoided. The web lost layer is arranged on the web outer molded surface, so that the web surface of the male die tool is a plane, the characteristic matching between the web of the C-shaped beam and the web of the tool is reduced, and the forming quality of the C-shaped beam can be improved; and meanwhile, the demolding scheme of pushing from the wing root to the wing tip is supported, the design difficulty of a male mold tool can be effectively reduced, and the machining efficiency is improved.

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 invention belongs to the technical field of aircraft manufacturing, and discloses a wing root fairing body and a manufacturing method thereof. The wing root fairing body comprises a core material, a potting piece and a skin. The core material comprises a first split body and a second split body; the first split body and the second split body are plate-shaped and are connected with each other to form the core material with an L-shaped cross section; the first split body is provided with a filling and sealing piece mounting groove, and the filling and sealing piece is arranged in the filling and sealing piece mounting groove; and the skin is coated and packaged on the outer surface of the core material. According to the wing root fairing body, the manufacturing method is relatively simple, the machining difficulty is small, the production period is short, and the machining cost is relatively low.