113 results about "Aileron" patented technology

The invention discloses a fixed wing type cross-domain aircraft capable of realizing self-adaptive load replacement, belongs to the technical field of cross-domain aircrafts, and is used for a water-air double-domain operation cross-domain aircraft. Comprising an aircraft body, a main wing, an aileron, a tail assembly, a vertical take-off and landing support, a vertical take-off and landing rotor wing, a fixed wing propeller, an underwater propeller, a load replacement device and a control module. The control module comprises a decision making system, a telemetering communication system, a state sensing system, an underwater water surface load replacement system, a water surface four-rotor vertical take-off and landing control system, an air fixed-wing cruise control system and an underwater vehicle control system. According to the fixed wing type cross-domain aircraft, through collaborative innovation of a multi-mode power system and self-adaptive buoyancy control, the dynamic characteristic adaptation problem in cross-domain navigation is solved, and the fixed wing type cross-domain aircraft has the vertical take-off and landing capacity, the air cruising capacity and the underwater maneuvering capacity.

The invention discloses a tilt rotor aircraft depending on aerodynamic force, and relates to the technical field of aircrafts. The four-axis unmanned aerial vehicle comprises a fuselage, wing assemblies, four-axis unmanned aerial vehicle modules and empennage assemblies. The wing assembly comprises a left wing, a right wing and a wing cross beam which are all provided with ailerons; the four-axis unmanned aerial vehicle module is I-shaped, is hinged to a wing cross beam through a bearing, and can tilt bidirectionally; the empennage assembly comprises a tail support, a horizontal electric variable pitch propeller, a horizontal empennage with a control surface and a vertical empennage. The four-axis unmanned aerial vehicle module serves as a tilting unit, and the structure is simplified; attitude control is realized by virtue of aerodynamic difference, and the efficiency is improved; the tail horizontal electric variable-pitch propeller realizes rapid pitching adjustment; vertical take-off and landing, low-altitude maneuvering and high-speed cruising capabilities are integrated, the application scene is wide, and safe forced landing can be achieved when faults occur.

Various control surfaces are disclosed that can cause a side force along a lateral axis running through a center of mass of a craft when the control surfaces are deployed in a same direction. The side force can be created without creating a yawing moment on the craft. The control surfaces can be on a winglet of the craft. Also disclosed are a split-flap control surface and stand-alone ailerons.

The utility model discloses an integrated aileron driving system, which belongs to the field of aircraft control systems and comprises a fixed wing, an aileron is arranged on one side of the fixed wing, the aileron is rotatably connected with the fixed wing through a first linkage shaft, the upper surface of the aileron is rigidly connected with a first speed reducing plate, and a second speed reducing plate is arranged on the other side of the fixed wing. A power rotating assembly is arranged on the upper surface of the fixed wing, the power rotating assembly is connected with the upper surface of the first air brake and used for driving the first air brake to move, fixed limiting pieces are symmetrically arranged at the two ends of the surface of the side, close to the aileron, of the fixed wing, and a second air brake is arranged between the fixed limiting pieces. The second air brake is located on the lower surface of the aileron, and the first air brake and the second air brake are connected through a transmission assembly. According to the integrated aileron driving system, through the linkage design of the aileron and the air brake, the structure is simplified, pneumatic coordination control is achieved, and the control performance at different flight speeds is improved.

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 application provides a kind of aircraft servo elastic dynamics model correction method considering operating surface gap, first aircraft structure is divided into substructure, and the aircraft dynamics model containing aileron gap nonlinearity is established according to substructure branch modal;And additional inertia force caused by aircraft operating aileron deflection is introduced in dynamics model, and the aircraft open-loop servo elastic motion equation containing aileron gap nonlinearity under independent generalized coordinate is obtained;According to sensor output equation, rudder transfer function and flight control system equation, the aircraft nonlinear servo elastic closed-loop system model containing aileron gap nonlinearity is established, and the model is corrected based on ground vibration test, rudder frequency response test and gap measurement test result, and the stability of the model is judged based on simulation analysis result and open-loop frequency response test result and is further corrected, which improves the accuracy of gap nonlinearity aircraft aerodynamic servo elastic characteristic evaluation, and is helpful to guide aircraft servo elastic ground test engineering practice.

An aircraft featuring at least one lifting rotor, one or more lifting airfoil-shaped wings, and at least one thrust-generating device capable of spinning the wings or the aircraft with the wings around its centre axis, characterised by the aircraft being able to change its orientation during flight by varying the aerodynamic lift of at least one of the airfoil-shaped wings by a device or a combination of devices including but not limited to, aileron or wing tip control surfaces, orwing pitch modulation, or wing profile modulation, or by modulating the lift produced by the lifting rotors around the aircraft centre of rotation, or a combination of said devices, and with a controller for the aerodynamic lift modulating device.

The application provides a swept-back three-beam thin wing composite wing structure, which comprises a wing main body and an aileron installed in a wing main body mounting groove, the wing main body comprises a wing skeleton and a wing outer skin structure outside the wing skeleton; the wing skeleton comprises a longitudinal member unit and a transverse member unit, the longitudinal member unit comprises a front beam, a middle beam and a rear beam, the front beam and the middle beam and the middle beam and the rear beam are arranged at acute angles, the transverse member unit comprises a plurality of wing ribs arranged between the front beam and the middle beam, between the middle beam and the rear beam and on one side of the front beam towards the front of the fuselage; the front beam, the middle beam and the rear beam are all provided with a spar connecting head for connecting with the fuselage. The wing provided by the application adopts a three-beam structure, each wing beam has a mouth-shaped cross section, so that each wing beam forms a multi-wall form, can bear the bending moment and the torque brought by a large wing effective load, and the wing structure is light in weight, high in strength and good in integrity.

Wheels are arranged on the lower portion of a ground carrier in a bilateral symmetry mode, a power system and a communication and automatic driving control system are arranged in a vehicle head on the front portion of the ground carrier, and a connecting clamping groove is formed in a floor of the ground carrier. Electromagnetic spring buckles are arranged on the lower portion of the flight cabin and correspond to the four connecting clamping grooves formed in the ground carrier floor in position. Wings with upward inverted inner sections and straight outer sections are respectively arranged on the left and right sides of the upper part of the flight cabin; the inner section and the outer section of each wing are averagely divided into two sections; when the hydraulic mechanism is opened and closed, the inner section and the outer section can be folded or unfolded around the hinge connecting mechanism, and after folding, the spring positioning pins are inserted into the positioning holes. Oblique empennages are arranged on the left side and the right side of the tail portion of the upper portion of the flight cabin respectively, and deflection rudders are arranged on the rear edges of the outer ends of the oblique empennages. The hovercar can run on the ground, can take off and land vertically, can hover in the air or fly horizontally and can sail underwater; the safety and reliability are high, and the use efficiency is high.

The systems, methods, and other embodiments described herein relate to controlling the elevons of an aircraft to stabilize the fuselage of a free-wing aircraft. In one embodiment, the system includes a processor and a memory storing machine-readable instructions. When an instruction is executed by the processor, the processor 1) measures the fuselage pitch attribute of the aircraft, 2) measures the wing pitch angle of the wing assembly, and 3) controls the angles of the elevons of the wing assembly based on the fuselage pitch attribute and the wing pitch angle, so that the fuselage and wing assembly of the aircraft rotate independently of each other and freely around the pitch axis of the aircraft.

The application discloses a wing, rib and skin integrated forming auxiliary wing mold and a forming control method thereof, and belongs to the technical field of unmanned aerial vehicle composite auxiliary wings, and comprises an upper mold and a lower mold, characterized in that a forming surface of the upper mold is matched with a cavity surface of the lower mold, and a three-dimensional closed auxiliary wing cavity for overall forming of the auxiliary wing is formed after the mold is closed; the auxiliary wing cavity comprises a main beam cavity platform extending along the length direction of the auxiliary wing, a plurality of rib cavity platforms arranged at intervals with the main beam cavity platform and arranged in a vertical direction, and a continuous skin transition curved surface connecting the main beam cavity platform and the rib cavity platform; and a plurality of glue injection ports and exhaust ports in communication with the auxiliary wing cavity are arranged on the lower mold. The application can complete overall manufacturing of the auxiliary wing in one forming process, significantly reduces assembly links, and the formed product has a continuous fiber reinforced overall structure, no mechanical connection holes, uniform stress distribution, and significantly improved structural strength and service life.

This aims to create an attitude control system that can respond appropriately to pilot inputs during maneuvers such as turning and landing. [Solution] An attitude control device installed in a radio-controlled model airplane, which is a controlled object that is operated based on control signals received from an external device, includes a calculation unit that selectively outputs a first aileron control signal input as a control signal and a second aileron control signal calculated to maintain the horizontal position of the controlled object based on the roll angle of the controlled object. The calculation unit selects and outputs the second aileron control signal when the function is instructed to be ON by a function setting signal received from an external source, the first aileron control signal is a signal indicating neutral, and the roll angle of the controlled object is within a set angle.

The utility model provides an aircraft airfoil skin repair training rack. The aircraft airfoil skin repair training rack comprises a mounting rack, an aileron skin assembly structure, a wing skin assembly structure and an empennage skin assembly structure. An operation table is arranged on the top of the mounting frame. The aileron upper skin wraps the upper surfaces of the aileron framework and the flap accessory, the aileron lower skin wraps the lower surfaces of the aileron framework and the flap accessory, the wing skin wraps the upper and lower surfaces of the wing framework through the outer end of the wing framework, and the empennage skin wraps the upper and lower surfaces of the empennage framework through the outer end of the empennage framework. The device is complete in parts, rich in structure, convenient to disassemble and assemble, and convenient for teaching and daily airplane airfoil skin repair practical training operation of trainees.

The application discloses a device for inhibiting the rotating movement of the aileron of an aircraft wing, and belongs to the field of aircraft structure design. The device can inhibit the rotating movement of the aileron of the aircraft wing. The device comprises a support frame, a limiting block, a limiting rod, a driving rod, a control motor and a support arm. The support frame and the support arm are installed on the rear wall of the wing through the support arm. When the device is started, the limiting block contains the limiting rod; when the device is restored, the limiting block is separated from the limiting rod. When the motor is started, the limiting block moves to the limiting rod along the axis of the rotating shaft. In the process that the limiting block moves to contain the limiting rod, the rotating angle of the limiting rod (the aileron) is gradually reduced until the rotation is stopped. The aileron flutter of the aircraft is caused by the coupling of the rotating mode of the aileron and other modes of the structure, so that the rotating movement of the aileron is inhibited, and the rotating flutter of the aileron of the aircraft is inhibited. The application can be used in the aircraft flutter test to inhibit the aileron flutter of the wing.

This invention belongs to the field of aircraft manufacturing technology, and provides an integral molding die for an aileron, a molding method, and an aileron. The die includes a lower die, a core die, side dies, and an upper die. The lower die includes a lower die pressing part and a lower fixing part; the lower fixing part is located on both sides of the lower die pressing part and is a protruding structure relative to the lower fixing part. The upper die includes an upper die pressing part and an upper fixing part; the upper fixing part is located on both sides of the upper die pressing part and is a protruding structure relative to the upper die pressing part. The core die is installed at the lower die pressing part and corresponds to the mounting shaft of the aileron. The side die is installed at the lower fixing part. After the lower die, upper die, and side die are closed, the inner surface of the side die is in contact with the ends of the lower die pressing part and the upper die pressing part, and the outer surface is in contact with the inner surface of the upper fixing part, forming a cavity structure corresponding to the shape of the aileron. This invention has the advantages of low cost, simple process, short time, and lighter aileron weight in aileron manufacturing.

An attitude control device mounted on a radio-controlled model airplane, which is a controlled object controlled based on a control signal received from an external device, comprises a calculation part configured to selectively output a first aileron control signal, which is input as the control signal, and a second aileron control signal, which is calculated based on a roll angle of the controlled object to maintain the controlled object in a horizontal state. The calculation part selects and outputs the second aileron control signal when the following conditions are satisfied: turning on of a function is instructed by a function setting signal received from the external device; the first aileron control signal is a signal indicating neutral; and the roll angle of the controlled object is within a set angle.

------ Vector-Tripping Aircraft The present invention relates to a vector-tripping aircraft (1), comprising: a wing in the form of a wing (2); a tail in the form of at least one rear aileron (3) carried by said wing (2); and an engine assembly (4) connected to said wing (2) for thrusting the aircraft (1), characterized in that said engine assembly (4) is disposed in front of said wing (2) and is connected to it by at least one arm (6), and that it comprises at least one turbine (10) configured to create, in operation, a jet of air discharged towards said wing (2) by a steerable nozzle (11). Figure to be published with the abbreviation: Figure 1

A wing simulation device comprises: a main wing simulation member; an aileron simulation piece which is installed on the main wing simulation piece and comprises an upper wing surface and a lower wing surface which are opposite to each other; and the mounting part is arranged at the front edge of the aileron simulation piece and is provided with an upper surface and a lower surface which are opposite to each other. Wherein the aileron simulation piece further comprises at least one pressure measuring hole, and the pressure measuring holes are arranged to penetrate through the upper wing surface or the upper surface of the mounting part; the at least one butt joint hole is arranged to penetrate through the lower surface of the mounting part; and at least one pressure measuring pipeline, each of the pressure measuring pipelines is arranged in the aileron simulation piece and is connected with one of the at least one pressure measuring hole and the corresponding one of the at least one butt joint hole. According to the wing simulation device disclosed by the utility model, the phenomenon that a pressure measuring pipeline is inserted into a groove formed in the outer part of the aileron simulation piece is avoided, so that the influence on the aerodynamic shape of the aileron simulation piece is avoided, and the angle of the aileron simulation piece can be conveniently replaced.

A rotorcraft (10) has a body (11) and a main rotor (12) provided on an upper part of the body. The main rotor rotates in the counterclockwise direction or the clockwise direction when viewed from above the body. The rotorcraft has an auxiliary blade (14) provided to the body. The auxiliary blade is provided at a port front part of the body if the rotation direction of the main rotor is the counterclockwise direction. The auxiliary blade is provided at a starboard front part of the body if the rotation direction of the main rotor is the clockwise direction.

The aircraft can include: an airframe, a tilt mechanism, a payload housing, and can optionally include an impact attenuator, a set of ground support members (e.g., struts), a set of power sources, and a set of control elements. The airframe can include: a set of rotors and a set of support members. By utilizing a larger rotor blade area (and / or larger rotor disc area) and adjusting the blade pitch and RPM, the rotors can augment the lift generated by the aerodynamic profile of the aircraft in the forward flight mode in addition to providing forward thrust. Variants generating lift with the rotors can reduce or eliminate additional control surfaces (e.g., wing flaps, ailerons, ruddervators, elevators, rudder, etc.) on the aircraft since the thrust and motor torque is controllable (thereby indirectly controlling lift) at each rotor, thereby enabling pitch, yaw, and / or roll control during forward flight.

The application provides an electric-duct vertical-lift co-rotating-wing aircraft, which comprises a fuselage, a main wing arranged in the middle of the fuselage, and engines connected to the fuselage through engine flight attack angle electric adjusting mechanisms respectively arranged on the front and rear sides of the main wing; the engine flight attack angle electric adjusting mechanisms are provided with ailerons; the front and rear engines are staggered vertically; and the front aileron, the main wing and the rear aileron are arranged in sequence from bottom to top. The electric-duct vertical-lift co-rotating-wing aircraft provided by the application is arranged at three heights, i.e. the lower, middle and upper heights, so that the engines and the wings are prevented from affecting each other when exerting their respective functions, and the stability of flight is improved.

The application discloses a kind of aileron surface failure under airplane lateral fault-tolerant control and reconstruction method, belong to aircraft fault diagnosis field.The application discloses a kind of aileron surface failure under airplane lateral fault-tolerant control and reconstruction method, for fixed-wing aircraft actuator common fault, an active fault-tolerant control scheme of integrated fault estimation algorithm and reconfigurable control is presented, and an integrated fault-tolerant control strategy based on multi-model adaptive unknown input observer and model reference adaptive control method combination is designed.To realize the fault estimation of flight control system, a robust unknown input observer is designed to estimate the actuator additive fault and multiplicative fault.According to the characteristics of the flight control system fault model, the additive fault is treated as external disturbance, and the multiplicative fault is used to design the adaptive law.Finally, the effectiveness of the proposed integrated fault-tolerant control strategy is verified through actuator stuck fault, failure fault and offset fault.

The invention discloses a follow-up fatigue test method and system for an aileron of an unmanned aerial vehicle. The method comprises the following steps: setting a loading point on a skin according to an aileron design aerodynamic load and an acting position, and enabling a resultant hinge moment to be consistent with a design value; and calculating the torsion spring rigidity of the simulation steering engine according to the hinge moment and the designed torsion angle. And the aileron hinge and the torsion spring are arranged on a clamp to form a follow-up support. The rigidity of the torsion spring is adjusted through counterweight calibration, so that the actual torsion angle of the aileron reaches the standard, and the rear edge calibration displacement is recorded. A loading system based on the liquid pressure equivalence principle is installed, and the loading system is driven by an actuator cylinder to evenly transmit loads to all loading points. And the calibration system is used for adjusting the displacement of the actuator cylinder to enable the trailing edge displacement to reach the calibration displacement and recording the displacement as a control parameter. And controlling the loading system to circularly load the aileron according to the parameters to finish the fatigue test. According to the invention, the stress state of the aileron under the real working condition can be accurately reproduced, and the fatigue performance of the aileron, the aileron connecting hinge and the transmission mechanism can be comprehensively assessed.

This invention provides a high-load thin-wing integral airfoil and its design method. The high-load thin-wing integral airfoil includes: a main wing, an aileron, and a wing-body fairing; the main wing is connected to the fuselage via a wing-body joint assembly; the aileron is fastened to the main wing via a servo mount and aileron inner and outer joints; the wing-body fairing is fixed to the main wing. The present invention features a rectangular composite material main spars, improving the load-bearing capacity of the wing structure. The load-bearing components, the main wing and aileron, are assembled using co-curing and secondary co-curing molding methods, reducing the number of wing structural parts, simplifying the production process, shortening the wing structure manufacturing cycle, and lowering costs. Compared with existing technologies, the technical solution of this invention can solve the technical problems of large number of parts, complex manufacturing processes, long production cycles, and high costs associated with traditional assembled wing structures.

The unmanned aerial vehicle aileron structural member comprises an unmanned aerial vehicle body, a bottom plate is fixedly arranged on the inner bottom wall of the unmanned aerial vehicle body, an empennage is fixedly arranged at the rear end of the top face of the bottom plate, and a pair of L-shaped supports is fixedly arranged on the front side of the top face of the bottom plate; horizontally-distributed canards are mounted in openings of the L-shaped supports in a hinged mode, and the L-shaped supports are connected with the canards on the same side through hinge mechanisms; two pairs of L-shaped side plates are fixedly arranged on the rear side of the top face of the bottom plate, a first main wing is installed between the top ends of each pair of L-shaped side plates in a hinged mode, a second main wing is installed between the bottom ends of each pair of L-shaped side plates in a hinged mode, and the bottom plate is connected with the first main wings and the second main wings through limiting mechanisms. The storage convenience and flight reliability are both considered, through collaborative optimization of all the components, the problems of lift force adjustment, attitude control and storage of the unmanned aerial vehicle are solved, the structure is compact, transmission is stable, the practicability of the unmanned aerial vehicle is remarkably improved, and the adaptive scene range of the unmanned aerial vehicle is remarkably widened.

The application belongs to the technical field of aircraft design, and particularly relates to a high-lift-coefficient layout aircraft and a use method thereof. The aircraft comprises a fuselage, a hump arranged on an upper middle part of the fuselage, a lower wing connected to the hump, an upper wing located above the lower wing and detachably connected to the lower wing, the upper wing being connected to the lower wing through an actuator and a four-bar linkage mechanism, the separation or adhesion of the upper wing relative to the lower wing being controlled by the actuator, a trailing edge flap and aileron arranged on a trailing edge of the lower wing, a V-tail stabilizer connected to a rear part of the fuselage and a V-tail control surface connected to a tail part of the V-tail stabilizer, and an engine connected to the fuselage. The application can greatly increase wing area and effective angle of attack, and significantly improve take-off and landing lift.

The invention discloses an invisible foldable swarm unmanned aerial vehicle, which relates to the technical field of unmanned aerial vehicles, and comprises a fuselage, two wings symmetrically arranged on two sides of the upper part of the middle part of the fuselage, two empennages arranged on two sides of the tail part of the fuselage, a thrust propeller arranged on the tail part of the fuselage, and ailerons arranged on the rear sides of the wings; folding mechanisms with elastic memory recovery are arranged at the joints of the wings and the fuselage, the joints of the empennages and the fuselage and the joints of the thrust propellers and the fuselage; the overall cross section of the fuselage is of a big-end-up trapezoidal structure used for reducing the radar reflection coefficient. According to the invisible foldable unmanned aerial vehicle, the wings and the empennage are designed to be folded and unfolded, honeycomb type storage of the unmanned aerial vehicle can be achieved, and rapid maneuvering and dense launching are achieved. The radar discovery rate can be reduced through the accurate appearance design, noise can be reduced through propeller propulsion, the concealment of the unmanned aerial vehicle is higher, and the battlefield survival rate and the saturation strike capacity of the unmanned aerial vehicle can be improved through the bee colony technology.

The utility model discloses an angle measuring device for an aircraft aileron steering wheel. The angle measuring device comprises an angle measuring scale 4 and a fastening assembly, the aircraft aileron steering wheel 1 is coaxially connected and in clearance fit with one end of the neck 2, and the other end of the neck 2 is fixedly connected with a steering column; the angle measuring scale 4 is in an open ring shape and is coaxially installed on the neck 2 through a fastening assembly, and scale marks are arranged on the angle measuring scale 4 in the circumferential direction. Zero graduation lines are arranged on the outer wall of a center shaft of the steering wheel 1 and the outer wall of the neck 2; the fastening assembly comprises a fastening bolt 5, a butterfly nut 8 and a filler block 6, the filler block 6 is located at the opening end of the angle measuring scale 4, and the fastening bolt 5 and the butterfly nut 8 fasten the filler block 6 and the angle measuring scale 4 on the neck 2; the steering wheel is rotated to any angle, and the rotation angle of the steering wheel can be directly and accurately read through the specific scale on the angle measuring scale corresponding to the zero scale line on the steering wheel, so that the debugging efficiency of the aircraft aileron is improved.

The utility model discloses a single-blade rotor type vertical take-off and landing fixed-wing aircraft which comprises a single-blade rotor assembly which penetrates through an aircraft body, is connected with a power system and is used for achieving the functions of vertical take-off and landing and hovering of the aircraft. The rotor wing clasping clamping device is mounted on a vertical tail wing framework of the fuselage main body and is used for clasping or loosening the single-blade rotor wing assembly; the homing driving device is mounted on a middle framework of the fuselage main body and is used for driving the single-blade rotor wing assembly to return to the initial position; the aileron driving devices are mounted on the tail and the side wings of the fuselage main body and are used for assisting the aircraft in completing flight actions such as turning and climbing; the power system is mounted on the fuselage main body and used for driving the single-blade rotor assemblies to move; the propelling assembly is installed on the fuselage and used for providing forward flight thrust. Compared with the prior art, the advantages of long endurance of a fixed-wing aircraft and vertical take-off and landing of a traditional helicopter are combined, and fuel economy and high flight speed are achieved in the cruising process.

The utility model discloses a low-altitude and low-speed unmanned aerial vehicle with a sweepforward wing configuration and a wing body fusion layout. The low-altitude and low-speed unmanned aerial vehicle comprises main wings, a fuselage and an empennage, the main wing adopts a forward swept wing structure, so that spanwise flow is effectively weakened, and the wingtip airflow speed is increased; the main wings and the fuselage adopt a wing body fusion layout, redundant connecting mechanisms are reduced, the lift force of the unmanned aerial vehicle is effectively improved, turbulence is weakened, the lift-drag ratio is increased, and aerodynamic load distribution is improved; the fuselage is provided with a cabin and an equipment mounting platform, can be used for carrying power, control and monitoring equipment, and is suitable for various task requirements; the empennage is fixedly connected to the rear end of the fuselage; the main wing is provided with a motor and an aileron; the empennage is provided with a horizontal tail and vertical tails, the vertical tails are fixedly connected to two sides of the horizontal tail, and the rear ends of the horizontal tail and the vertical tails are hinged to an elevator and a rudder. The low-speed aerodynamic performance of the unmanned aerial vehicle is effectively improved, the aerodynamic load layout of the unmanned aerial vehicle is optimized, and the maneuverability and the loading capacity of the unmanned aerial vehicle are remarkably improved. The utility model is applied to the technical field of aircraft structures.