418 results about "Aileron" patented technology

An electric tilt rotor aircraft and a control system thereof. The aircraft comprises a fuselage (1), wings (2), horizontal stabilizers (3) and a landing gear system (4), wherein the fuselage (1) comprises a fuselage body and a rudder (10) for controlling and adjusting a flight direction of the aircraft; the wings (2) each comprise an integral wing box section (21), an aileron (22), a wing left / right nacelle (23) and a supporting rod left / right nacelle (24); left and right supporting rods are integrated on front sides of the integral wing box sections (21); the wing left / right nacelle (23) is connected to the integral wing box section (21) by means of a tilting shaft (25), and the wing left / right nacelle (23) rotates around the tilting shaft (25); the horizontal stabilizers (3) each comprise a horizontal stabilizer box section (31) and a horizontal stabilizer left / right nacelle (32); the horizontal stabilizer left / right nacelle (32) is connected to the horizontal stabilizer box section (31) by means of the tilting shaft (25), and the horizontal stabilizer left / right nacelle (32) rotates around the tilting shaft (25); the landing gear system (4) uses a tricycle landing gear system, and comprises a nose landing gear (41) and two main landing gears (42); brake systems are integrated in the main landing gears (42); and the nose landing gear (41) plays a role in balancing and controlling direction. By using the electric tilt rotor aircraft, urban air traffic can be realized.

The invention relates to the technical field of aircrafts, and discloses an aircraft attitude control method, device and equipment and a storage medium, and the method comprises the steps: collecting flight state parameters of an aircraft; the pitch angle data in the flight state parameters are compared with a preset angle threshold value, corresponding flight mode switching is executed, and the attitude deviation value is calculated; power distribution calculation is conducted on the three ducts, and the target rotating speed of each duct paddle is obtained; performing position optimization calculation on the guide plates below the duct to obtain the position control quantity of each guide plate; the three flow guide plates are configured as triangular control points, the position compensation amount of an aileron, a yaw rudder and an elevator is calculated and determined through vector synthesis, and the angle control amount of each control surface is obtained; according to the method, the attitude of the aircraft is converted into an execution signal, execution feedback data is acquired to correct a control instruction in real time, closed-loop attitude control of the aircraft is realized, the attitude control precision of the aircraft is improved, and the stability of the flight process is ensured.

The invention discloses a distributed electric propulsion box type wing ground effect aircraft which comprises a main fuselage, the two sides of the main fuselage are connected with one ends of sweepback wing main wings, one ends of two empennages above the tail of the main fuselage are connected with the main fuselage through vertical tails, and the other ends of the two empennages are connected with the other ends of the two main wings through end plates respectively. The empennage is connected with a plurality of high ducted fans directly driven by a motor, an elevator and an aileron are arranged at the rear part of the empennage, a rudder is arranged at the rear part of the vertical fin, and the elevator, the aileron and the rudder are controlled by a flight control system and are mutually independent. The ground-effect aircraft provided by the invention adopts a box-type wing structure formed by combining the sweepback main wing and the sweepforward empennage, has better aerodynamic performance in a take-off and climbing low-speed flight state, and can provide larger lift force and better maneuvering characteristics.

The invention belongs to the technical field of glide control, and relates to a direct lift active disturbance rejection control method suitable for flying wing layout. According to the method, a trailing edge flap is introduced to serve as a main lift control surface, meanwhile, a constant attack angle is kept through a lifting aileron so as to balance part of pitching moment generated by the trailing edge flap, and therefore the attitude of the aircraft is stabilized, and the constant speed is maintained. In the actual flight process, external disturbance is inevitable, so that an original glide control strategy is improved and optimized in combination with an active disturbance rejection control method. Compared with a traditional glide control method, the flying wing layout direct lift active-disturbance-rejection control method provided by the invention has remarkable anti-interference capability and excellent glide control performance, and shows a wide application prospect.

A modular unmanned aerial vehicle (UAV) system comprises a body module, a rotor module, and a wing module. The body module includes a flight controller and a power distribution device. The body module is releasably attachable to the rotor module or the wing module, and the body module is releasably attachable to the rotor module. The rotor module includes one or more motors and electronic speed controllers (ESCs), while the wing module includes a wing having a flap, elevator, aileron, or rudder. Various UAV configurations can be formed from the body module, the rotor module, and the wing module. Each configuration includes different advantages for flight time, distance, battery life, and payload capacity. A UAV can be configured to a particular configuration to optimize parcel delivery.

The invention provides a foundation pit supporting system, and belongs to the technical field of constructional engineering. According to the structure, an inner supporting frame is located in an outer supporting frame, a rectangular steel column of a bridging frame is located between the outer supporting frame and the inner supporting frame, the two Y-shaped ends of the bridging frame abut against the inner wall of the outer supporting frame and the outer wall of the inner supporting frame respectively, and a connecting disc of the bridging frame abuts against the inner wall of the inner supporting frame; the reinforcing anchor rods extend into the external contact structure, and a plurality of vertical supporting columns are arranged on the outer wall of the outer supporting frame. A frame structure with an inner layer and an outer layer is constructed, on the basis, a whole is formed through the bridging frames, and the reinforcing anchor rods are installed; the reinforcing anchor rod is wedged into the soil layer of the foundation pit and is fixed with the soil layer by using the anchor bolt with the wing plate and the aileron, so that the stability is fully enhanced; meanwhile, the vertical supporting columns are arranged on the outer side of the outer supporting frame, and the structural strength is guaranteed. The supporting structure has better structural strength and is stable and reliable when used for supporting the foundation pit.

The retractable lift force type wind-assisted propulsion device comprises a sail assembly, a retractable device and a control system, and the sail assembly can rotate relative to a ship body; the sail assembly comprises a main wing sail, a first aileron sail, a second aileron sail and a lower end plate; wherein all the sails are wing-shaped sails, and the main wing sails are elliptical wing-shaped sails; each sail is rotatably connected with the lower end plate; and the rotating axes of the sails are parallel to each other and are not coplanar. Through the arrangement of the multi-body wing-shaped sail, the multi-body wing-shaped sail can adapt to different wind directions to change the attack angle, and the wind power boosting effect is improved; through the arrangement of the folding and unfolding device, the sail can be freely folded and unfolded in different working states.

The utility model discloses a double-layer wing type vertical axis wind turbine, which relates to the technical field of wind power generation and comprises a wind turbine body and at least two wing type blade groups, each airfoil blade group comprises a main airfoil blade, an auxiliary airfoil blade and a plate sheet; the auxiliary wing-shaped blades are arranged on the sides, close to the rotating center line of the wind turbine body, of the main wing-shaped blades, and first wind passing channels are formed between the auxiliary wing-shaped blades and the main wing-shaped blades. The plate sheets are arranged on the sides, close to the rotating center line of the wind turbine body, of the main wing-shaped blades, and second wind passing channels communicated with the first wind passing channels are formed between the plate sheets and the main wing-shaped blades. And a clearance space is formed between one side edge of the plate sheet and the tail edge of the auxiliary airfoil blade. The pressure difference between the upper surface and the lower surface of the main wing-shaped blade can be increased through the design of the plate sheets, the blade tip vortex generated on the lower surface of the main wing-shaped blade can be weakened, then the lift force borne by the main wing-shaped blade is increased, and the purpose of further improving the self-starting performance and the power generation efficiency of the wind turbine is achieved.

The invention belongs to the field of vertical take-off and landing aircrafts, and relates to a distributed contra-rotating rotor hybrid tilting rotorcraft. Comprising a front tilting electric contra-rotating rotor nacelle, a front cabin body, a wing end tilting electric contra-rotating rotor nacelle, a V-shaped empennage, a V-shaped empennage aileron, a vertical empennage, a retractable undercarriage, a green hybrid power energy system and an empennage tilting electric contra-rotating rotor nacelle, the front tilting electric contra-rotating rotor nacelle is hinged with the other end of the front cabin body; an empennage tilting electric contra-rotating rotor nacelle is arranged at the end part of the V-shaped empennage; wing end tilting electric contra-rotating rotor wing nacelles are arranged at the ends of the wings; and rotor wing units are arranged in the rotor wing nacelles, the six rotor wing units all adopt a coaxial reverse upper and lower two-paddle structure, and the rotating directions of the rotor wings in the six rotor wing units are that the adjacent rotating directions and the opposite rotating directions are opposite.

According to the method for determining the maximum hinge moment of the flaperon in the aircraft scheme design stage, the corresponding state of the maximum hinge moment of the flaperon under the landing configuration and the corresponding state of the maximum hinge moment of the flaperon under the cruise type are analyzed and determined; the maximum flaperon hinge moment of the landing configuration and the maximum flaperon hinge moment of the cruising configuration are calculated, the larger value is used as the maximum flaperon hinge moment, high calculation efficiency and accuracy are achieved, in the aircraft scheme design stage, wind tunnel tests do not need to be carried out, and the design efficiency is improved. A large amount of calculation of the hinge moment of the full-envelope full-configuration wing is not needed, the available flaperon maximum hinge moment is rapidly obtained, and the design of a flaperon actuator is supported.

A device may include a locking assembly configured for coupling with an aileron of an aircraft, the locking assembly includes a lock configured for coupling with the aileron. The locking assembly may include a lock mechanism operatively coupled with the lock and in communication with an accelerometer, where the lock mechanism is configured to transition the lock from a locked configuration to a released configuration. The lock mechanism opens the lock based on a specified acceleration measured with the accelerometer, and the lock frees the aileron to a dynamic configuration. A device may include a tailored aileron deployment system including a biasing member coupled with the aileron, in the released configuration the biasing member controls movement of the aileron in the dynamic configuration.

A deformable wing with variable camber of front and rear edges is composed of a driving device, a middle-section fixed wing, a movable rear edge and a movable front edge. Through the combined movement of the wings, the smooth and continuous deformation of the wings can be realized, and the resistance caused by discontinuous deformation and gaps of the traditional wings is eliminated. The deformable wing capable of changing the camber of the front edge and the rear edge has the advantages of being simple in structure, low in cost, light and firm in structure and large in variable angle, the lift-drag coefficient of the wing can be changed under the control of an airborne computer according to different flight tasks and environments through smooth and continuous deformation on the premise of low resistance and low noise, and the wing can be applied to the aircraft. The wing can be used as a flap to improve cruising speed and shorten takeoff and landing taxiing distance, and can also be used as an aileron to ensure that resistance coefficients of a left wing and a right wing are the same when lift coefficients change, a rolling-yawing coupling effect is reduced, and flight quality is remarkably improved.

The utility model discloses a reversible control surface of an unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles. The reversible control surface of the unmanned aerial vehicle comprises an unmanned aerial vehicle body, the number of the fixed control surfaces is two, and one end of each fixed control surface is fixedly connected to the outer surface of one side of the unmanned aerial vehicle body; and the number of the disassembling mechanisms is four, and each disassembling mechanism comprises a mounting box, a rotating component, an adjusting component, a moving component and two pushing components. According to the turnover control surface of the unmanned aerial vehicle, the positions of the two moving plates are adjusted through cooperation of the adjusting component and the moving component, the fixing blocks are separated from the fixing grooves through arrangement of the pushing component, and the movable control surface is taken down from the mounting box and replaced with a new movable control surface; the fixing grooves in the movable control surfaces are connected to the mounting box in a butt joint mode, the movable control surfaces are fixed through the pushing parts, and under the condition that the aileron and the empennage are not replaced, the maintenance cost and the maintenance time are reduced by replacing the single movable control surface.

The invention relates to a ground effect aircraft capable of reducing the steering radius. A main wing of the ground effect aircraft comprises a wing surface, an aileron and a main beam, the tail end of the main beam is fixedly connected with a buoy; the buoy comprises a buoy main body, a spoiler, an aquaplane, a first driving mechanism and a second driving mechanism; one end of each of the spoiler, the water ski, the first driving mechanism and the second driving mechanism is rotationally mounted in the buoy main body; the other end of the first driving mechanism is hinged to the middle of the spoiler. The other end of the second driving mechanism is hinged to the middle of the water ski. According to the ground-effect aircraft, the aircraft body can be pressed down through the additionally arranged spoilers, so that the aircraft can press the water ski into water within a controllable range under the action of the spoilers, the resistance of the single side of the aircraft is increased, and therefore large yawing torque is obtained, and small-radius steering is achieved.

The invention provides a flying wing layout aircraft single-shot failure flight quality evaluation method which comprises the following steps: a control surface balance parameter calculation step: ignoring lateral displacement caused by sideslip angle change, and calculating a change curve of asymmetric rolling and yawing moment caused by balanced single-shot failure, required aileron skewness delta a and rudder skewness delta r along with speed; a control surface balance capability evaluation step: judging whether the aircraft can be balanced after single-engine failure at each speed by judging whether the aileron skewness delta a and the rudder skewness delta r exceed limit values or not, and evaluating the flight quality of the aircraft after single-engine failure; a time domain response simulation step: performing time domain simulation after a single aircraft engine fails; and an attitude keeping capability evaluation step: based on a time domain simulation result after the single aircraft engine fails, judging the attitude change degree of the aircraft, and evaluating the flight quality of the single aircraft engine failure. On the other hand, the invention provides a flying wing layout aircraft single-shot failure flight quality evaluation system.

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.

According to the electric tilting helicopter wing, a flap steering engine end adjusting support is connected with a middle hinge structure of a flap conversion rocker arm, one end of a flap control surface connecting rocker arm is connected with the right end of the flap conversion rocker arm, and the other end of the flap control surface connecting rocker arm is connected with a flap control surface fixing connector; the aileron steering engine root mounting frame is located on the wing rib, and a single lug of the aileron steering engine root mounting frame is connected with double lugs of the aileron steering engine root; a middle fixing frame of the aileron steering engine is positioned on the rear beam, and two sides of the aileron steering engine are respectively arranged on the wing ribs; an aileron steering engine end adjusting support is connected with the rear beam, and an aileron steering engine end double-lug connector is connected with the left end of an aileron conversion rocker arm; an aileron steering engine end adjusting support is connected with a middle hinge structure of an aileron conversion rocker arm, one end of an aileron control plane connecting rocker arm is connected with the right end of the aileron conversion rocker arm, and the other end of the aileron control plane connecting rocker arm is connected with an aileron control plane fixing connector. A force transmission path of bending, twisting and shearing loads caused by concentrated mass of aerodynamic loads, a transmission system and a rotor system is provided, and the strength and rigidity of the flap and aileron steering engine connecting structure are guaranteed.

The utility model relates to a static aeroelastic force-measuring wind tunnel test model for a large-flexibility wing structure, which has the capability of testing the bending moment of a wing root and belongs to the technical field of aeroelastic tests of aircrafts. A static aeroelastic force measurement wind tunnel test model of a large flexible wing structure comprises a wing body fusion section, an outer wing section and a winglet, the wing body fusion section is arranged between the outer wing section and the winglet, the outer wing section comprises a wing section, an aileron, a spar and a steering engine, and the wing section is made of foam; the spar is bonded with the wing section, the length direction is consistent with that of the wing section, one end of the spar is connected with the winglet, and the other end of the spar extends out of the end part of the wing section to serve as a mounting end fixed with the side wall of the wind tunnel; the steering engine is connected with one side of the spar; the aileron serves as a part of the wing section, is bonded to the other side of the steering engine and can rotate along with actuation of the steering engine. The utility model has the advantages that: 1, the structural form is simple; 2, the assembly is simple, and the price is low; 3, the overall bearing load of the foam is reduced; and 4, the operation is easy.

The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to a tailstock type overwater vertical take-off and landing fixed-wing unmanned aerial vehicle which can float and park on the water surface, can vertically take off to leave the water and is switched into a fixed-wing mode for cruising and level flight. The unmanned aerial vehicle comprises a four-rotor support buoy and a wing body fusion body fixed to the upper portion of the four-rotor support buoy, small wing type buoys are arranged on wing tips of the two sides of the wing body fusion body respectively, and lifting ailerons capable of deflecting up and down are arranged on the rear edges of wings on the two sides of the wing body fusion body respectively. And four motor rotor propeller propelling devices are uniformly arranged at the tip part of the four-rotor bracket buoy.

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 utility model relates to the technical field of automobile spoilers, in particular to an adjustable automobile spoiler metal casting which comprises a main wing plate, and sliding grooves are formed in the two end faces of the main wing plate in the length direction. Auxiliary wing plates are slidably mounted at the two sliding grooves, and a plurality of fixing grooves are formed in the bottoms of the two auxiliary wing plates in the length direction; side wing plates are fixed to the ends, away from each other, of the two auxiliary wing plates. Fixing parts are fixed at the two ends of the bottom of the main wing plate; the top end of the fixing plate penetrates through the fixing base and sequentially penetrates through the top of the fixing base and the top of the support in a sliding mode, and the top end of the fixing plate can penetrate through the bottom of the main wing plate in a sliding mode and is matched with the corresponding fixing groove in an inserted mode. The auxiliary wing plates are installed in the sliding grooves in the two ends of the main wing plate in a sliding mode, the auxiliary wing plates on the two sides can be conveniently adjusted to stretch out and draw back, and the using requirements of different vehicle types are met; and meanwhile, the fixed plate is matched with the fixed grooves in different positions in an inserted manner, so that the adjusted aileron plate can be limited, and the mounting stability of the aileron plate is ensured.

The invention discloses a vertical take-off and landing control method for a distributed power tandem wing unmanned aerial vehicle, and belongs to the technical field of unmanned aerial vehicle flight control. An inner ring power unit and an outer ring power unit of the unmanned aerial vehicle are arranged at the front edges of the front wing and the rear wing of the tandem wing unmanned aerial vehicle; lifting ailerons are symmetrically hinged to the rear edges of the two ends of the rear wing; the front wing and the rear wing are connected into a tandem wing with a space height difference through a connecting assembly; the control method comprises the steps of controlling jumping take-off, transition flight, cruise flight and vertical landing, and performing accelerator quantity distribution on the inner ring power unit and the outer ring power unit for jumping take-off, cruise flight and vertical landing; a continuous climbing transition mode is adopted in the transition flight stage, and the flight state is controlled to be converted into a cruise level flight state after transition is completed. According to the unmanned aerial vehicle, the functions of large-angle jump take-off and 90-degree autonomous vertical landing of the tandem wing unmanned aerial vehicle are achieved, distributed power of the unmanned aerial vehicle is controlled and distributed, the energy utilization rate is increased, and the endurance time is prolonged.

The invention discloses a hydrogen-powered unmanned aerial vehicle, and belongs to the field of unmanned aerial vehicles, the hydrogen-powered unmanned aerial vehicle comprises a fuselage structure and a power system, the fuselage structure comprises a nacelle fuselage provided with a hydrogen storage bottle, a propeller is installed at the front end of the nacelle fuselage, the power system is installed in the nacelle fuselage, a supporting plate is fixed above the nacelle fuselage, and the two sides of the supporting plate are connected with unmanned aerial vehicle wings; the unmanned aerial vehicle wings are connected with ailerons, the unmanned aerial vehicle wings are connected with the unmanned aerial vehicle empennage through connecting rods, the unmanned aerial vehicle empennage is connected with an elevator and a rudder, and an undercarriage is fixed below the nacelle fuselage. The power system adopts a proton exchange membrane hydrogen fuel cell as an energy conversion device, the fuel cell can efficiently convert hydrogen and oxygen into electric energy and only generate water and heat, the efficiency of the conversion process is far higher than that of a traditional internal combustion engine or a battery, and therefore the cruising ability of the unmanned aerial vehicle is improved, and the unmanned aerial vehicle is reasonable in design, long in service life and low in cost. The energy is saved.

A control system for rotating an airfoil along one axis that is offset from a perpendicular axis that varies the angle of attack (AOA), dihedral, and center of lift based on the position of the attachment to precisely maneuver an aircraft. This technique can eliminate the need for traditional control surfaces such as ailerons, elevators, and rudders. allowing for an aerodynamically advantaged shape.

The utility model discloses a flying wing layout miniature unmanned aerial vehicle capable of realizing vertical take-off and landing by tilting ducted fans, which can realize vertical take-off and landing and hovering as well as high-speed cruise, and comprises an integrated fuselage and wings, flaps positioned at the inner sections of the wings are used for providing lift force and resistance for runway take-off and landing, and the flaps are used for providing lift force and resistance for runway take-off and landing. The ailerons located on the inner sections of the wings are used for providing rolling torque during flight, the ducts located at the two ends of the wings are used for controlling transverse stability in the vertical take-off and landing and cruising process, the two main ducted fans are longitudinally distributed, and the main ducted fan on the front portion of the fuselage is vertically fixed to provide lift force. The main ducted fan at the rear part of the fuselage tilts by 90 degrees through the tilting mechanism to realize power required by vertical take-off and landing and cruising, and the wing body fusion technology of the flying wing layout has good aerodynamic characteristics and reasonable mass distribution.

The utility model discloses a control surface angle measuring device, which relates to the technical field of angle measurement, and comprises a measuring tool dial and a measuring tool pointer, the measuring tool dial is fixed on a main wing of a wing, the measuring tool pointer is fixed on an aileron of the wing, and the measuring tool pointer is fixed on the aileron of the wing. Angle measurement is realized through the position where the measuring tool pointer points to the scales on the measuring tool dial. The utility model discloses a control surface angle measuring device, which is characterized in that a measuring tool dial is fixed on a main wing of a wing, a measuring tool pointer is fixed on an aileron of the wing, the aileron of the wing drives the measuring tool pointer to deflect, the deflection angle of the aileron of the wing is converted into the deflection angle of the measuring tool pointer, and the measuring diameter is enlarged. The angle reading is easier, the reading and measuring accuracy is greatly improved, the limitation of the measuring space is small, and the method can be suitable for measuring the deflection angles of various airplane control surfaces.

The utility model discloses a connecting structure of an aileron and a wing. The connecting structure comprises an outer aileron, a knuckle bearing, a detachable check block, a detachable rotating shaft, a hinge rotating shaft, a hinge fixing detachable support, an aileron driving pull rod, the wing, an inner aileron and a rotating shaft. The knuckle bearing is embedded into the outer end face of the inner aileron and the inner end face of the outer aileron; the detachable rotating shaft is mounted on the outer side of the outer aileron; the outer end of the outer aileron is close to the position area of the detachable rotating shaft; the hinge rotating shaft is mounted between the inner aileron and the outer aileron and is embedded into the wing end rib through the knuckle bearing; the wing end rib is embedded into the inner side close to the inner aileron; the hinge fixing detachable support is mounted on a wing rear beam; the aileron driving pull rod is mounted on the wing; the inner aileron and the outer aileron are positioned on the rear edge of the wing and close to the wing tip; and the rotating shaft is mounted on the inner side of the inner aileron. The aileron hinge is simple in structure, and the aileron hinge joint can be hidden to the maximum extent, so that the influence on the pneumatic surface of the aileron is reduced.

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.

Provided are a method and a system for cooperative airborne meteorological perception with multiple aircraft. The method includes: determining a training set based on a spatial position, motion data, weather, an observation image, and static information of each of multiple aircraft; determining a fusion performance function based on an objective of a cooperative perception network; generating a fusion model based on the training set and the fusion performance function; generating a perception result image based on the fusion model; determining flight risk information based on the perception result image; determining a route alarm command and sending the route alarm command to a management user interface and an interactive device of a target aircraft; after receiving the route alarm command, generating, by the management user interface, a route adjustment parameter; in response to obtaining the route adjustment parameter, performing at least one of the following operations: adjusting an engine velocity, adjusting an aileron pressure, adjusting an elevator deflection, and adjusting a rudder deflection.

In the event of a failed engine, an automatic takeoff thrust asymmetry compensation system (“ATACS”) for an aircraft improves capabilities to reduce VMCG and deal with the potential side-effects simultaneously. The system commands selected control surfaces (which can be e.g., rudder and / or ailerons and / or spoilers or any combinations thereof) for a short period of time, improving the capability to reduce the VMCG without increasing the penalty on system failures or poor handling qualities.