311 results about "Wingspan" patented technology

The invention provides a morphing wing integrating telescopic variable-elongation and rotary variable-sweepback functions and a morphing method. The morphing wing comprises an inner wing section, an outer wing section and an elastic skin, the inner wing section and the outer wing section are positioned in the same spanwise plane; the inner wing section stretches and retracts in the spanwise plane along the spanwise direction of the wing, so that the span length of the inner wing section is adjusted; the outer wing sections rotate in the chordwise direction relative to the inner wing sections in the spanwise plane, and then the sweepback angles of the outer wing sections are adjusted. By adjusting the extension length and the sweepback angle of the wings, the lift force of the aircraft during low-speed flight is improved, and the resistance of the aircraft during high-speed flight is reduced. Specifically, the extension length of the wings is increased during low-speed flight to improve lift force, and the extension length is reduced and the sweepback angle is increased during high-speed flight to reduce resistance. The optimal aerodynamic performance can be achieved no matter in a low-speed flight state or a high-speed flight state, and the adaptability and the flight performance of the aircraft are remarkably improved.

An aircraft is disclosed having a device and an aircraft wing. The aircraft wing includes a main wing element and a movable wing tip device attached to a tip end of the main wing element. The movable wing tip device includes an accumulator configured to store energy and the movable wing tip device can move relative to the main wing element to vary a span of the aircraft wing. The accumulator is configured to transmit power to the device. The accumulator enables energy to trickle between the main wing element and the movable wing tip device whilst still providing a suitable power source to the device. Additional embodiments include a method of maintenance, and a method of operating the aircraft wing to store energy in the movable wing tip device.

The invention discloses a hypersonic variable wingspan aircraft attitude control method based on reinforcement learning, aiming at the characteristics of complex flight environment, strong nonlinearity of control input and aerodynamic characteristics, strong coupling and the like of a hypersonic variable configuration aircraft. The method comprises the following steps: acquiring discrete relationships between variables such as wingspan length, rudder deflection angle, flight attitude and flight speed and aerodynamic parameters through a computational fluid mechanics method, and then fitting the discrete relationships by utilizing a neural network method to realize aerodynamic parameter identification under different flight working conditions and appearance conditions; secondly, establishing a dynamic model of the hypersonic variable-wingspan aircraft by using an aerodynamic parameter identification network; and finally, realizing attitude control of the hypersonic variable-wingspan aircraft based on a DQN reinforcement learning method. The method can effectively solve the problems that a traditional control method is complex in controller design and strict in model precision requirement, and simulation experiment results prove that the method can effectively achieve attitude angle tracking.

The invention discloses a multi-degree-of-freedom morphing wing based on a zero poisson ratio superstructure, and belongs to the technical field of morphing wings. The wing comprises a front edge section, a wing body and a rear edge section which are sequentially connected, the front edge section comprises an inner side front edge plate, a middle front edge plate and an outer side front edge plate, the wing body comprises a plurality of zero-Poisson-ratio honeycomb structures, each zero-Poisson-ratio honeycomb structure is formed by arraying a plurality of zero-Poisson-ratio unit cell cubic structures, and the adjacent zero-Poisson-ratio unit cell cubic structures share a connecting rod; the rear edge section comprises an inner side rear edge plate, a middle rear edge plate and an outer side rear edge plate; the inner-side front edge plate and the inner-side rear edge plate, the middle front edge plate and the middle rear edge plate, and the outer-side front edge plate and the outer-side rear edge plate are respectively connected through zero-Poisson-ratio honeycomb structures which are arranged along the chord length direction of the wing; the two sides of the middle front edge plate are connected with the inner side front edge plate and the outer side front edge plate through zero-Poisson-ratio honeycomb structures arranged in the extension direction of the wing correspondingly, and the two sides of the middle rear edge plate are connected with the inner side rear edge plate and the outer side rear edge plate through zero-Poisson-ratio honeycomb structures arranged in the extension direction of the wing correspondingly. The wing has four deformation functions of variable elongation, chord length, dihedral angle and camber, so that the aircraft can adapt to different flight environments and tasks.

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

The invention discloses a regional unmanned aerial vehicle safety risk assessment method, which comprises the following steps: acquiring parameters such as unmanned aerial vehicle body length, wingspan width and body height, and preset task scene data such as regional tall building area ratio, unmanned aerial vehicle density, specific parameters and quantity of aircrafts; based on the data, regional ground and airspace safety risk values are calculated. And determining visibility, rainfall and wind speed as meteorological risk assessment indexes, performing weighting by using an analytic hierarchy process, and calculating a meteorological risk assessment coefficient. And constructing a regional unmanned aerial vehicle safety risk assessment model by integrating assessment data obtained under three environmental conditions of weather, ground and airspace. And comprehensively evaluating the safety risk of the unmanned aerial vehicle in the region by using the model. According to the technical scheme of the invention, three environmental factors of meteorological conditions, ground conditions and airspace conditions are integrated, and various environmental risk types possibly experienced by the unmanned aerial vehicle in the operation process can be comprehensively evaluated.

According to various embodiments, there may be provided a monocopter. The monocopter may include a body chassis. The monocopter may further include a wing structure extending from the body chassis, the body chassis being at a root of the wing structure. The wing structure may include a plurality of rigid wing segments distributed along a spanwise direction extending between the root and a tip of the wing structure. The wing structure may further include a plurality of flexible wing segments, each flexible wing segment adjoining a pair of adjacent rigid wing segments of the plurality of rigid wing segments. The monocopter may further include a propulsion unit coupled to a rigid wing segment of the plurality of rigid wing segments between a midspan of the wing structure and the tip of the wing structure.

The invention relates to the field of aircraft wing design, in particular to a multi-stage telescopic wing for improving the flight performance of an aircraft, the extension length and the area of the wing are changed through multi-stage stretching of the wing, and different requirements of the aircraft for the wing in low-speed and high-speed flight states are met. When the aircraft takes off, lands and flies at a low speed, the telescopic wings stretch out to increase the area of the wings and enable the aircraft to obtain enough lift force, and when the aircraft flies at a high speed, the telescopic wings retract to reduce the extension length and area of the wings and enable the aircraft to obtain enough lift force and reduce the resistance of the wings, so that the extension length and area of the wings are changed through extension of the wings; and the aircraft can simultaneously meet different requirements on the extension length and the area of the wings in low-speed and high-speed flight states.

Towed aerodynamic platforms approach the behavior of flat plate airfoils to achieve high lift-to-drag (L:D) performance when operated at low pitch angles. A lead aircraft may tow multiple platforms to form train units. Low form drag and high L:D enable solar aircraft that are more robust and faster than approaches extending wingspan rather than extending the length of the aerial train. Applications extend beyond solar aircraft and include vehicles towed along an overhead monorail system and aerial drones.

The invention relates to the technical field of unmanned aerial vehicles, in particular to an inflatable wing type variable wingspan unmanned aerial vehicle. The unmanned aerial vehicle comprises a vehicle body, variable inflatable wings are arranged on the two sides of the vehicle body, the variable inflatable wings are connected with an inflation and deflation mechanism arranged in the vehicle body, each variable inflatable wing comprises a hollow fixed inflatable wing body and an extended inflatable wing body, and an inflatable part of each extended inflatable wing body is arranged outside a hollow cavity of the corresponding hollow fixed wing body; the uninflated part of the extended inflatable wing body is arranged in the hollow cavity, and the extended inflatable wing body is connected with the telescopic part of the extended driving mechanism. According to the inflatable wing type variable wingspan unmanned aerial vehicle, the variable inflatable wings are adopted, the wingspans of the whole wings are adjusted by lengthening the inflatable wing bodies, the unmanned aerial vehicle adapts to different flight states, and the unmanned aerial vehicle is of a relative folding structure, so that multi-stage adjustment is achieved.

The invention discloses a safety interval calculation method for an unmanned aerial vehicle and a manned aircraft, and relates to the technical field of aviation safety. The safe interval calculation method comprises the following steps: step 1, firstly obtaining the length, wingspan and height of a first target, then obtaining the length, wingspan and height of a second target, marking the first target as a mass point at the moment, and making up for an error caused by volume loss when the first target is regarded as the mass point. A new collision model is used for carrying out quantitative analysis on the collision probability of a manned aerial vehicle and an unmanned aerial vehicle so as to find the relation between the flight interval and the collision probability and calibrate a reasonable safety interval, and the requirement of the safety interval with a fixed value for the airspace cost is high. Therefore, the dynamic safety interval is found by adopting interval division iterative calculation, and a more flexible dynamic safety interval between the unmanned aerial vehicle and the manned vehicle is given on the basis, so that the construction of a fusion flight system is accelerated.

The invention belongs to the technical field of aircraft flight load design, and particularly relates to an aircraft pneumatic rolling torque correction method, which comprises the steps of initializing calculation parameters, calculating aircraft vacuum speed, equivalent speed and aircraft speed pressure based on the calculation parameters, then calculating motion parameters based on maneuvering simulation dynamic response, and calculating full-aircraft target pneumatic rolling torque based on the motion parameters. Giving unbalanced torque tolerance; solving an unbalanced moment based on the whole-aircraft target rolling moment; and comparing and analyzing the difference value between the unbalanced moment and the unbalanced moment tolerance, and if the absolute value of the unbalanced moment is smaller than the unbalanced moment tolerance, ending the calculation. By adjusting the local attack angle of each section in the spanwise direction of the wing part and reconstructing the pressure distribution characteristic in the spanwise direction of the wing, the unbalanced pneumatic rolling moment of the aircraft caused by the pressure distribution characteristic is eliminated, the purpose of load balance is achieved, and the design requirements in engineering are met.

An aircraft wing tip that is movable by an actuator between a flight configuration, and a ground configuration which has a shorter wingspan for use in ground-based operations. Lugs that connect to arms of the actuator are integrally formed as part of the primary structure of either the fixed wing or outboard wing section, such as a folding tip section.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The invention provides a horizontal measurement method of a large-wingspan movable wing aircraft, which comprises the following steps of: establishing a coordinate system by taking a virtual coordinate on a tool as a reference, adjusting a basic platform and a supporting arc to be within an allowable deviation range, positioning a middle section of a product by using the supporting arc, measuring the middle section, judging whether the levelness of a mounting surface of the middle section reaches the standard or not, and adjusting. And sequentially assembling the rear section and the front section by taking the middle section as a standard of complete machine assembly, completing assembly and measurement of the fuselage, and adjusting each section of the fuselage in real time according to a measurement result. And then the wing is assembled, and measurement of the wingtip coordinate and the deflection angle is completed. The tool has the advantages that assembly and measurement are carried out on one tool, alternate station change of products between an assembly table and a measurement table is avoided, a machine body does not need to be rotated during measurement, measurement is completed in the assembly state, the assembly efficiency is guaranteed, and the product safety can be guaranteed.

The invention discloses a novel heavy-load aircraft with a stub-wing lift fuselage. Comprising an airfoil-shaped lift fuselage, front stub wings connected to the two sides of the nose of the lift fuselage in a smooth transition mode, rear stub wings connected to the two sides of the tail of the lift fuselage in a smooth transition mode, a propeller power mechanism arranged on the nose of the lift fuselage, a V-shaped empennage arranged on the tail of the lift fuselage and a ventral fin arranged on the belly of the lift fuselage. The wingspan of the novel heavy-load stub wing lift fuselage aircraft is less than or equal to 2m, the wingspan of the front stub wing is smaller than that of the rear stub wing, and the mounting angle of the front stub wing is larger than that of the rear stub wing; after the scheme is adopted, the lift fuselage can provide large lift force, the short wings can provide a part of lift force, meanwhile, longitudinal balancing of the aircraft can be completed by adjusting the mounting angle, the static stability requirement of the aircraft is met, the performance of high-speed cruise is reserved, the wingspan of the wings is reduced, and the stability of the aircraft is improved. Therefore, the unmanned aerial vehicle has flexible performance and can be applied to narrow and complex environments.

The utility model relates to a gliding type precipitation enhancement bomb which comprises a bomb body, a telescopic wingspan frame, a gravity center seat, an endurance motor and flying wings, the telescopic wingspan frame is arranged on the bomb body in a foldable mode, the gravity center seat is arranged between the bomb body and the telescopic wingspan frame, and the endurance motor is arranged on the telescopic wingspan frame. And the flying wings cover the telescopic wingspan frames and are unfolded along with the telescopic wingspan frames. The rain enhancement bomb is simple in structure, convenient to produce and manufacture and low in manufacturing cost. In addition, besides the advantage that rainfall is sufficient due to long-time hang, the gliding type rain enhancement bomb is low in speed and safer when landing.

The invention provides a conformal structure suitable for a large-wingspan flexible wing and a design method, and relates to the technical field of aircraft mechanical structure design. The unmanned aerial vehicle is provided with a fuselage and inflatable wings; fuselage fixing pieces are installed on the surface of the fuselage, at least one pulling point is arranged on the surface of each inflatable wing in the wingspan direction of the surface of the inflatable wing, at least one pulling point is located at the wingtip of the inflatable wing, wing fixing pieces are installed at the pulling points, and the wing fixing pieces are connected with the fuselage fixing pieces through traction pieces. According to the large-wingspan flexible inflatable wing, the wing can be prevented from generating large structural deformation in the lift force direction under the gravity and pneumatic action, and the engineering application problem that an existing large-wingspan flexible inflatable wing is large in deformation under the gravity and pneumatic load can be solved; the pulling point arrangement along the spanwise direction and the chordwise direction can be flexibly determined according to flight conditions; a certain dihedral angle or dihedral angle is pre-set for the inflatable wing by adjusting the length of the pull rope, so that the transverse stability of the aircraft is improved.

The invention belongs to the technical field of aircraft design, and relates to an aircraft wing rib section mass discretization method, which comprises the following steps of: (1) setting a virtual trapezoidal shape covering a wing, and equidistantly setting a plurality of chordwise grid points on each spanwise tangent plane on the virtual trapezoidal shape along the spanwise direction of the wing; (2) determining the position of the mass gravity center of the rib section on the spanwise tangent plane; (3) dispersing the mass of the rib sections to adjacent spanwise tangent planes along the spanwise direction according to a lever ratio; (4) dispersing the mass of the section to chordwise grid points along the chordwise direction according to trapezoidal distribution; (5) segmental integration is carried out to calculate the quality of chordwise grid points; according to the method, a trapezoidal shape is arranged along the shape of a wing, the mass of a wing rib section is dispersed to adjacent spanwise sections along the spanwise direction of the wing according to a lever ratio, and the mass of the sections is dispersed to chordwise grid points along the chordwise direction of the wing according to trapezoidal distribution. The mass is reasonably dispersed under the condition that the mass gravity center of the rib section is not changed, so that the wing mass distribution is closer to the real condition, the wing load calculation precision is improved, and the structural design is safer and more reliable.

The invention belongs to the technical field of spacecrafts, and particularly relates to a flexible solar wing ground unfolding gravity unloading mechanism. Comprising a guide rail and a plurality of gravity unloading units. The gravity unloading unit comprises pulleys, a supporting beam, a rotating shaft, a swing rod and a balancing weight, the two ends of the supporting beam are connected with the two pulleys, the pulleys are installed on the guide rails, the rotating shaft is rotationally installed on the supporting beam, the lower end of the rotating shaft is connected with the swing rod, and hanging points at the tail end of the swing rod are connected with wing plate mass center hanging points of the solar wing through hanging ropes respectively; the upper end of the rotating shaft is connected with the balancing weight; when the solar wing is unfolded, the tackle moves longitudinally along with the mass center of each wing plate and freely slides along the guide rail, the swing rod swings along with the transverse movement of the mass center of each wing plate, and the balancing weight balances the torque generated by hanging the solar wing. The device adapts to the extremely small distance when the flexible solar wing is folded, and can move along with the mass center in the unfolding process of the solar wing in the longitudinal direction and the transverse direction, and gravity unloading in the ground unfolding process of the flexible solar wing is achieved.

Disclosed is a wing assembly for an aircraft. The wing assembly comprises a fixed wing, a wing tip device, a wiring harness and a harness guide. The wing tip device is moveably mounted at a joint at an end of the fixed wing. The wing tip device is moveable about the joint between: (i) a flight configuration for use during flight, and (ii) a ground configuration for use during ground-based operations, in which ground configuration the wing tip device is moved relative to the fixed wing such that the span of the wing is reduced. The wiring harness extends between the fixed wing and the wing tip device. The wiring harness comprises a plurality of conductors and is arranged to transmit electrical power and / or data to the wing tip device. The harness guide is located between the fixed wing and the wing tip device. The harness guide comprises a helical channel configured to receive the wiring harness and guide the wiring harness between the fixed wing and the wing tip device.

A method for reducing aerodynamic loads on an aircraft located on the ground includes moving a wing tip relative to a fixed wing to an on-ground configuration in which the span of the wing is reduced. The wing tip includes an airflow channel, which is in an open configuration in which airflow through the channel between upper and lower surfaces of the wing tip is enabled, when the wing tip is in the on-ground configuration. This has been found to reduce aerodynamic loads on the wing tip and / or parts of the aircraft connected thereto during this phase of the aircraft's use.

Closed wing aircraft design and configuration with three wings on either side of the fuselage wherein at least one closed frame is established between the 1st wing and the 2nd wing, by using at least one bracing entity. A separate closed frame is established between the 2nd wing and the 3rd wing, by using at least one bracing entity. Each of said closed frames defines its own aerodynamic channel. The framework is strong and stiff, the mutually supported parts level out the stress. The fuselage is lifted in three points. The Aspect Ratio is high and the wingspan is relatively short but there are embodiments wherein the 3rd wing has an extended folding tip section.

The invention provides a patrol flight unmanned aerial vehicle with W-shaped folding wings. The patrol flight unmanned aerial vehicle comprises a fuselage, inner-section wings, inner-section wing limiting structures, inner-side folding and unfolding mechanisms, outer-section wings, ailerons, outer-side folding and unfolding mechanisms, empennages, empennage folding and unfolding mechanisms and empennage protection structures. The fuselage is designed by adopting a flat lifting body configuration, the inner and outer sections of wings are designed by adopting low-speed friendly plane shapes such as rectangles and trapezoids, and the empennage is designed by adopting a V-shaped empennage / inverted V-shaped empennage. The inner section wing and the outer section wing of the unmanned aerial vehicle are integrally in a W shape in the unfolding process, and a folding wing layout with the larger aspect ratio and the higher aerodynamic efficiency can be derived based on the W-shaped layout. According to the layout, the extension length of the wings is slightly restrained by the length of the fuselage, the contradiction between the aerodynamic characteristic design of the wings and the storage space of the launch canister is well balanced and solved under the condition that the length of the fuselage is fixed, and the layout wings are large in aspect ratio, high in aerodynamic efficiency and simple in structural design.

The invention discloses a distributed power fixed-point take-off and landing portable solar unmanned aerial vehicle, and belongs to the technical field of overall comprehensive design of unmanned aerial vehicles. Comprising a front wing, a rear wing and a folding assembly, the front wing and the rear wing are arranged in a tandem wing mode, the folding assembly is connected between the front wing and the rear wing, the front edge of the front wing and the front edge of the rear wing are each provided with a plurality of rotors in the spanwise direction, and the four rotors located in the spanwise middle of the front wing and the spanwise middle of the rear wing are located at the four vertexes of a quadrangle respectively to form the four-rotor power system of the unmanned aerial vehicle. The other rotor wings form a fixed wing power system of the unmanned aerial vehicle; the front wing and the rear wing are connected into a tandem wing with a space height difference through the folding assembly, and the front wing and the rear wing can be in a stacked state through deformation. The problems that a single-type propeller of a tailstock type configuration machine consumes much energy, the efficiency of the cruising stage is low, and the endurance is short are solved.

A wing structure is disclosed having a first span portion including a depth that reduces in a spanwise outward direction, and a second span portion outboard of the first span portion and having a substantially uniform spanwise depth D3. The second span portion may be non-tapered, and may include parallel fore and aft spar portions. In some embodiments the non-tapered portion supports a movable wing tip device. A wing assembly and an aircraft are also disclosed.

The invention discloses an atmospheric parameter sensing system and an atmospheric parameter resolving method, and relates to the technical field of aircrafts, the system comprises a pressure value acquisition module, a vector flow velocity sensor and a resolving module; the pressure value acquisition module and the vector flow velocity sensor are both connected with the resolving module; the pressure value acquisition module is used for acquiring pressure data on the surface of the aircraft; the vector flow velocity sensor is used for collecting spanwise flow velocity components on the surface of the aircraft; the spanwise flow velocity component is the component of the airspeed in the spanwise direction of the wing; and the resolving module is used for resolving the actual attack angle, the actual sideslip angle and the actual air speed of the aircraft according to the pressure data and the spanwise flow velocity component. According to the application, the resolving workload is reduced.

The telescopic variable-pitch cross-medium propeller mechanism comprises a vane seat shell, a supporting frame is arranged in the vane seat shell, an electric telescopic rod assembly is arranged on the supporting frame, vane assemblies are connected to the left end and the right end of the electric telescopic rod assembly in a threaded mode, each vane assembly is formed by sleeving a plurality of multi-section vanes, and the vane seat shell and the electric telescopic rod assembly are connected in a threaded mode. The front end of the electric telescopic rod assembly is connected with a wing piece at the foremost end of the wing piece assembly, and the wing piece assembly is driven by the electric telescopic rod assembly to stretch and retract. Appropriate lift force is provided for the cross-medium aircraft or the aircraft, modular design is adopted, wing panel serialization design can be carried out according to the load requirement of the cross-medium aircraft or the aircraft, the wing panel extends into a large wingspan during air flight, the wing panel contracts into a small wingspan during underwater navigation, and the screw pitch of the propeller is increased.

The utility model discloses a full-lifting-body perovskite solar unmanned aerial vehicle which comprises a large-chord-length lifting body fuselage and wings on the two sides, ailerons are arranged on the rear edges of the outer sections of the wings to serve as rolling control faces, and an elevator is arranged on the rear edge of the fuselage to serve as a pitching control face. A pair of vertical stabilizers are arranged on the outer side surface of the rear edge of the fuselage to enhance the course stability; perovskite solar cells with flexible substrates are distributed on the surfaces of the fuselage and the wings in an array mode and provide power for the unmanned aerial vehicle. The large-chord-length lifting body fuselage is adopted, the lifting area of the unmanned aerial vehicle is increased, meanwhile, the wingspan length of the unmanned aerial vehicle is reduced, loads borne by wings of the unmanned aerial vehicle are reduced, the weight of the wing structure is reduced, and effective loads are increased; the flexible perovskite solar cells are adopted for array, the damage rate of the cells in the installation process is reduced, and meanwhile due to the advantages of low cost, excellent flexibility, foldability, light weight and mass production of the perovskite solar cells, the service life of the unmanned aerial vehicle is prolonged, the cost is reduced, and mass production of the unmanned aerial vehicle is facilitated.

The invention discloses a flying wing and buoyancy lift integrated layout mooring aerostat type high-altitude wind power generation system which comprises an air aerostat wind power generation system, an air-ground electric energy transmission and mooring cable and a ground anchoring system. The air aerostat wind power generation system comprises a flying wing configuration thin-wall truss force bearing structure, a plurality of buoyancy air bags, two empennages, a skin component and three wind power generator sets, and the flying wing configuration thin-wall truss force bearing structure is provided with a plurality of constraint spaces distributed in the wingspan direction; the plurality of buoyancy airbags are respectively arranged in the plurality of constraint spaces; the two empennages are arranged at the two ends of the rear portion of the flying wing type thin-wall truss force bearing structure respectively. The skin component is arranged on the flying wing configuration thin-wall truss force bearing structure and the two empennages in a sleeving manner; the three wind generating sets are installed at the bottom of the flying wing type thin-wall truss force bearing structure through a supporting structure. The device is stable in structure, can stably generate power in a long-time air-staying mode, is low in cost, and is autonomous and controllable.