99 results about "Wing root" patented technology

The invention discloses a positioning tool for aircraft wing installation, and belongs to the technical field of aircraft wing installation, the positioning tool comprises three hydraulic telescopic cylinders and a supporting mechanism for installing and supporting the hydraulic telescopic cylinders; the three hydraulic telescopic cylinders are distributed on the supporting mechanism in a triangular shape. The output end of the hydraulic telescopic cylinder is provided with a rotary adjusting supporting assembly used for supporting wing installation. The rotary adjusting and supporting assembly comprises a supporting rod fixedly connected to the output end of the hydraulic telescopic cylinder; the rotating sleeve is rotationally connected to the upper end part of the supporting rod; the supporting rod rotatably supports the rotating sleeve; the bottom bracket is fixedly connected to the rotating sleeve; and the top bracket is mounted on the bottom bracket. According to the positioning tool, the wing is lifted through the whole positioning tool, and it is ensured that the position shape of the root of the wing meets the position requirement of the opening in the fuselage, so that the wing is more stable and accurate in the installation process, and the installation efficiency and the installation effect are improved.

The invention provides a folding wing force measuring device, a calibration structure of the force measuring device and a calibration method, and relates to the technical field of wind tunnel tests. The folding wing force measuring device comprises a wing root connecting piece, a folding wing inner section, a folding wing force measuring balance, a folding wing outer section, a plurality of angle connecting blocks and a plurality of folding patch blocks, wherein the angle connecting blocks and the folding patch blocks are matched with the folding angle of a folding wing. According to the invention, the balance and the wing part are fused into a whole and are integrally designed and processed, the connecting surface of the balance and the wing is canceled, the measuring end of the balance, namely the folding wing part section, is shaped by the cover plate, and the shape is kept consistent with that of the wing part section. Therefore, the load of the force measuring balance tends to be reasonable, the balance coordinate system changes along with the folding wing, the precision of the force measuring balance with the folding wing is improved, the data conversion difficulty of a wind tunnel test is reduced, and the precision of the wind tunnel test is improved.

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 present invention discloses a composite material winged projectile structure forming method, including: S1, pre-fixing the projectile and the wing; pre-assembling the projectile and the wing through a fixing tool, preliminarily determining the relative position of the upper fixing frame and the lower fixing frame; applying structural glue to the wing root surface where the wing and the projectile are in direct contact, determining the final position of the wing, and the wing inclination α must be less than 0.2 degrees; S2, secondary fixing the projectile and the wing; determining the side structural glue application radius R according to the projectile diameter, curing the side structural glue, and secondary fixing the wing and the projectile; S3, forming the projectile and wing reinforcement area; laying fabric prepregs of different sizes and quantities in the reinforcement area in sequence, and then heating and curing. The process controllable and reliable forming method provided by the present invention improves the safety and reliability of the winged projectile structure by controlling the application radius of the structural glue on the wing root surface and the accuracy and consistency of the laying of the reinforcement area, which is conducive to exerting the ultimate performance of the aircraft.

The invention discloses a retractable flapping wing mechanism based on a miniature flapping wing air vehicle, which simulates the flapping wing folding and unfolding motion of ladybirds at wing roots, replaces a common rigid connection form with a compliant mechanism, and realizes the folding and unfolding motion of a flapping wing leading edge beam of the miniature air vehicle by utilizing the deformation of a strip-shaped spring. According to the mechanism, the flapping wings are folded to the two sides of the aircraft body through bending deformation of the strip-shaped springs, the spanwise size of the aircraft during storage can be effectively reduced, and the portability is improved; before the aircraft takes off, the flapping wings can be automatically unfolded by releasing the strain energy of the strip-shaped spring. According to the design, the occupied space of the aircraft is reduced, the launching distance of the unmanned aerial vehicle is increased, the task execution radius is increased, and the application scene is expanded.

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 discloses a three-degree-of-freedom miniature bionic flapping-wing air vehicle for operation in a rugged environment of Mars, and the three-degree-of-freedom miniature bionic flapping-wing air vehicle comprises a rack, wings, a circuit control part, a flapping mechanism wing root regulation and control part and a landing part, the two pairs of wings are connected with the flapping mechanism through wing root pieces, and the wing roots of the two pairs of wings are connected with the wing root regulation and control part through titanium-nickel alloy wing roots; the wing root regulation and control part is used for carrying out three-degree-of-freedom regulation and control on pitching, rolling and yawing by controlling a titanium-nickel alloy wing root; the circuit control part is connected with the wing root regulation and control part through a carbon rod; the landing part is installed at the end, away from the wing root regulation and control part, of the circuit control part. The titanium-nickel alloy material is used as the titanium-nickel alloy wing root, better regulation and control performance and sensitivity are achieved, and meanwhile three-degree-of-freedom control over the ornithopter is achieved in combination with the rotary steering engine regulation and control mechanism and the linear steering engine regulation and control mechanism.

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.

The present invention proposes a method for quickly applying a follower load for a static test of an ultra-large flexible wing, comprising the following steps: Step 1: fixing the wing root on a load-bearing wall, fixing one end of a steel wire rope at a loading point, and hanging a weight on the other end; arranging a fixed pulley on a horizontally movable stand; connecting a strain gauge to a strain meter, and connecting a camera and a strain meter to a computer; Step 2: preloading the wing using a weight of a set mass, visually observing the deformation of the wing and moving the fixed pulley until the deformation is stable, so that the loading direction is approximately perpendicular to the wing beam axis, using a computer to determine whether the loading direction is perpendicular to the wing beam axis, adjusting the pulley position until the two are perpendicular to each other, comparing the difference between the test strain and the theoretical strain, and proceeding to the next step if the difference is less than a set threshold; Step 3: loading step by step according to a percentage according to the load application requirement, the loading process of each level is the same as that of Step 2, and recording the strain after each level of loading; Step 4: completing all loading and ending the test.

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 present invention belongs to the technical field of unmanned aerial vehicles (UAVs) and discloses a folding wing release and locking mechanism, comprising a wing root torsion spring (4), a folding wing connector (6), a wing root tension spring (7), a base (8), a rotating shaft (9), a wing folding locking pin spring (12), a wing root locking pin (13), a sales promoter (14), a locking pin push rod (15), a wing folding locking pin seat (16), a wing tip locking pin (18) and a wing folding locking spring (19). The present invention solves the problem of insufficient torque during the unfolding process of the folding wings of large-sized and heavy UAVs by using a wing root torsion spring-tension spring combination scheme, and effectively controls the unfolding timing of the folding wings by arranging a sales promoter mechanism at the wing tip.

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 belongs to the technical field of aircraft lifting design, and particularly relates to a full-aircraft lifting sling based on wing root bearing and a lifting method thereof.The full-aircraft lifting sling based on wing root bearing comprises a lifting frame, a top lifting ring and a bottom lifting plate; the hanging frame is rectangular; the top hoisting ring is connected with hoisting equipment; the top lifting ring is positioned above the lifting frame, is positioned above the central part of the lifting frame, and is connected with four corner parts of the lifting frame through cable-stayed steel cables; wing-shaped openings are formed in the two bottom hanging plates and are used for being connected to wing root parts of wings on the two sides of an aircraft in a sleeving manner; a rubber plate is laid in the wing-shaped opening; the two bottom lifting plates are located below the lifting frame and located on the edges of the two sides of the lifting frame, and the two ends of the two bottom lifting plates are connected with the two ends of the edges of the two sides of the lifting frame through vertical pulling steel cables.

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 present invention discloses a low-noise bionic flapping wing and a preparation method thereof. The flapping wing includes a wing membrane and wing veins. The wing membrane of the flapping wing is formed by laminating a low-permeability material membrane and a porous fiber material membrane. The two material membranes are placed one above the other, and the wing veins are placed in the middle to form a composite sandwich structure of the low-permeability material membrane - wing veins - porous fiber material membrane. The wing veins are divided into main wing veins, bifurcated wing veins, contour wing veins and arc wing veins; in order to improve the bending resistance of the wing, contour wing veins and arc wing veins are respectively arranged at the boundary and in the middle of the wing membrane. There are five main wing veins, which are radially and evenly distributed from the wing root to the wing tip. The bifurcated wing veins are in a three-bifurcation form and bifurcate from the corresponding main wing veins. The design of this flapping wing maintains the deformation mode for efficient generation of high lift and further restricts the wrinkling deformation during wing flapping, achieving low-noise generation.

The ultrahigh component assembling system comprises a bottom base, the front end of the bottom base is slidably provided with a front sailing stand column in the length direction of the bottom base, the rear end of the bottom base is fixedly provided with a rear sailing stand column, and a top cross beam is arranged between the top end of the front sailing stand column and the top end of the rear sailing stand column in a spanning mode; a side framework is arranged at the bottom of the top cross beam, and the side framework is arranged on the bottom base in a sliding mode in the width direction of the bottom base; a plurality of wing root positioning devices are distributed on the top of the bottom base in the length direction, an aileron positioning device is arranged on one side of the aft stand column, a front part positioning device is arranged at the bottom of the top cross beam, and a beam positioning device is arranged on the side face of the side framework. According to the ultrahigh part assembling system, the ultrahigh part can be integrally positioned and fixed in a surrounding manner at a time, so that the ultrahigh part is integrally positioned, and the position precision of the ultrahigh part in the subsequent assembling process is effectively guaranteed.