296 results about "Aileron" patented technology

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 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 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 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.

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.

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 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.

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.

The present application belongs to the technical field of aircraft operation and control. Provided is a method for determining the maximum hinge moment of ailerons of a transport aircraft. The method comprises the following steps: determining a hinge moment calculation formula for ailerons of a transport aircraft; determining an aircraft speed when the maximum hinge moment of the ailerons occurs, and, on the basis of the aircraft speed, determining calculation parameters corresponding to the maximum hinge moment of the ailerons, the calculation parameters comprising aileron deflection and aircraft angle-of-attack; and, on the basis of the hinge moment calculation formula for ailerons and the calculation parameters, obtaining the maximum hinge moment of the ailerons. The method in the present application can perform rapid calculation only by obtaining the aircraft speed, aircraft angle-of-attack and aileron deflection corresponding to the maximum hinge moment of the ailerons, can rapidly obtain the maximum hinge moment of the ailerons without performing full-envelope and full-configuration hinge moment calculation, is suitable for determining the maximum hinge moment of ailerons during an aircraft schematic design phase, and achieves high accuracy of results.

[Problem] To provide a flying object, a control method, and a program capable of controlling to achieve an increase in the cruising distance of the flying object and the stability of the flying object during hovering and landing by providing rotatable auxiliary wings to the flying object and controlling the angle of attack. [Solution] Provided is a flying object comprising: a main body part; a plurality of rotary wing parts provided to the main body part; an auxiliary wing part provided to the main body part; and a control part for controlling the operation of the rotary wing parts and the auxiliary wing part. The control part executes a change control of the lift of the auxiliary wing part in accordance with a comparison result between an acquired thrust-related parameter and one or more reference parameters.

The invention relates to the technical field of aircrafts, and discloses a triphibian aircraft and a triphibian form switching method thereof. Comprising an aircraft main body, a left auxiliary body and a right auxiliary body, retractable wings are arranged between the aircraft main body and the left auxiliary body and the right auxiliary body, sliding wheel assemblies are arranged at the bottoms of the three bodies respectively, foldable rotor wing mechanisms are arranged at the tops of the left auxiliary body and the right auxiliary body, a propeller is arranged at the tail of the aircraft main body, and foldable small wings are arranged on the outer sides of the left auxiliary body and the right auxiliary body. During take-off, the wings and the winglets are unfolded, the rotor mechanisms are started, and rising is achieved; during air cruise, the sliding wheel assemblies and the rotor mechanisms are stored, the propellers are started, and smooth flight is ensured; when the unmanned aerial vehicle runs on land, the sliding wheel assemblies are opened, the rotor mechanisms are stored, the wings and the winglets are folded, wind resistance is reduced, and the propellers are started to run; when sailing on the water, the three-body structure ensures buoyancy, and the propeller is pushed to move forwards; through the three-body structure, seamless switching of air, land and water triphibian scenes is achieved, the system adapts to multiple environments, the use flexibility is improved, and the application range is widened.

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 aileron structure of a flapping-wing aircraft, which relates to the technical field of flapping-wing aircrafts and comprises a through-type grid keel, blades and a gear rotating shaft, the blades are arranged in grids of the through-type grid keel and are rotatably connected with the grids, and the gear rotating shaft is arranged on one side of the through-type grid keel and is connected with the flapping-wing aircraft; the blades comprise side edge rotating blades and center rotating blades, the side edge rotating blades are rotationally connected with the side edges of the grids, and the center rotating blades are arranged in the grids and rotationally connected with the side walls of the grids. The through type grating keel and the rotatable wing edge form an aileron structure with the wide middle and the two narrow sides, and wind resistance can be effectively reduced; and the side rotating blades are combined with the central rotating blade, so that the wind resistance can be quickly and effectively increased, and quick sudden stop can be realized in an emergency.

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.

An unmanned aerial vehicle comprises: a fuselage (1) containing a nose fairing (2) and a load compartment (6), a front and a rear wings (3, 7) attached to the fuselage (1), a front camera (15) and a sighting camera (16) mounted to the fuselage (1), an engine with a propeller (14) mounted on the rear of the fuselage (1), a communication module with an antenna (17) mounted on the fuselage (1). The front wing (3) contains two height rudders (4) and two vertical fins (5) at its ends, the rear wing (7) contains ailerons (8). The unmanned aerial vehicle additionally contains a keel (11) with a directional rudder (12) and an altimeter (18). Said rear wing (7) additionally contains two consoles (9) at its ends, made with the possibility of attaching to it in the process of assembling the UAV. The said fuselage (1) containing the nose fairing (2) and a load compartment (6) is made of flat plywood sheets.