Strong wind disturbance resistant unmanned aerial vehicle and control method thereof

Abstract

The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to a strong-wind-disturbance-resistant unmanned aerial vehicle and a control method thereof.Box-type wings are installed on a fuselage, multiple pairs of coaxial rotors are arranged on a plane forming an inclination angle with the datum plane of the fuselage, and the characteristics of low induced resistance and light weight of a non-planar lifting surface (box-type wings) are combined, so that the strong-wind-disturbance-resistant unmanned aerial vehicle is obtained; the wind disturbance resistance and the stability of the unmanned aerial vehicle in a strong wind environment are improved; when the unmanned aerial vehicle encounters high-speed wind disturbance, the unmanned aerial vehicle adjusts the pitch attitude, and the lift force generated by the box-type wings compensates and balances the gravity and reduces the rotor wing load; meanwhile, under the strong wind working condition, the propellers are used for enhancing wind resistance forward force so as to offset wind disturbance; the empennage adopts a V-shaped design, so that the pitching stability is guaranteed, and the attitude adjustment response speed is optimized; and designing a robust adaptive flight control law, solving the problem of cooperative control of a heterogeneous control system, and realizing smooth state transition. The method has excellent wind disturbance resistance, and realizes high-precision attitude adjustment and energy efficiency optimization in a complex wind field.

Description

Technical Field

[0001] The present invention belongs to the technical field of unmanned aerial vehicles (UAVs), and in particular relates to a UAV capable of resisting strong wind disturbance and a control method thereof. Background Art

[0002] With the continuous advancement of science and technology, drone technology has developed rapidly, with its performance continuously improving, including flight stability, endurance, and payload capacity. Due to their small size, flexible operation, and low cost, they have been widely used in various fields. In addition to traditional fields such as aerial photography and agricultural plant protection, they are also being applied to emerging areas such as logistics and distribution, air traffic management, and smart city development. According to research, mature drone products with strong wind resistance and the ability to withstand strong wind shear in complex environments such as plateaus and sea areas are still relatively rare.

[0003] Low-level flow fields in plateaus, sea levels, and adjacent areas are prone to violent air currents and weather system activity. These can generate severe nonlinear aerodynamic forces on multirotor drones, causing them to be blown out of their hovering positions or even crash. Traditional designs rely primarily on the drone's power to support the load and resist wind forces, but this approach often limits the drone's power potential and results in poor wind resistance. Summary of the Invention

[0004] The purpose of the present invention is to provide a UAV that is resistant to strong wind disturbance and a control method thereof to solve the above-mentioned problems.

[0005] To achieve the above object, the present invention provides the following solutions:

[0006] A strong wind resistant UAV comprises a body and a driving mechanism arranged on the body;

[0007] The fuselage comprises a fuselage, two box-type wings are symmetrically arranged on opposite sides of the fuselage, and a tail fin is arranged at the tail end of the fuselage;

[0008] The driving mechanism includes a plurality of first driving members arranged on the box-shaped wing and a second driving member arranged on the tail wing, the first driving members being drivingly connected to rotors, the rotors providing lift and thrust for the fuselage, and the second driving members being drivingly connected to propellers, the propellers providing thrust for the fuselage;

[0009] The driving mechanism includes a multi-rotor working mode and a fixed-wing working mode. In the multi-rotor working mode, the first driving component controls the high-speed rotation of multiple rotors. In the fixed-wing working mode, the first driving component controls the low-speed rotation of multiple rotors, and the second driving component controls the rotation of the propeller.

[0010] Preferably, one end of a connecting rod is fixedly connected to the tail end of the fuselage. The connecting rod and the tail end of the fuselage have a smooth transition. The connecting rod is cylindrical. The tail wing is fixedly connected to the outer side wall of the connecting rod. The tail wing is located at one end of the connecting rod away from the fuselage. The second driving member is fixedly installed at one end of the connecting rod away from the fuselage.

[0011] Preferably, the tail wing includes two wing pieces. Both of the wing pieces are fixedly connected to the outer side wall of the connecting rod and are symmetrically arranged. The two wing pieces are arranged in a V shape with the opening facing upward.

[0012] Preferably, a power arm is fixedly connected to the box-shaped wing. The power arm is arranged along the axis direction of the unmanned aerial vehicle. An included angle is provided between the power arm and the box-shaped wing. Both ends of the power arm pass through the box-shaped wing and are respectively fixedly connected to the first driving member.

[0013] Preferably, the fuselage is in an arc transition on both sides along its own axis. The connection between the fuselage and the box-shaped wing has a smooth transition.

[0014] Preferably, the top surface of the fuselage and the top surface of the box-shaped wing form a non-planar lift surface.

[0015] A control method for an unmanned aerial vehicle against strong wind disturbances, based on the unmanned aerial vehicle against strong wind disturbances, includes the following steps:

[0016] Based on a preset dynamic model and kinematic equation, obtain an anti-wind flight control algorithm;

[0017] Based on the anti-wind control algorithm, use a cascade closed-loop control architecture to control the drive mechanism to work in a multi-rotor working mode or a fixed-wing working mode;

[0018] Predict the interference force and interference moment, and perform feedforward compensation on the anti-wind flight control algorithm based on the aerodynamic effect information of the box-shaped wing and the aerodynamic effect information of the wind field.

[0019] Preferably, the dynamic model of the unmanned aerial vehicle is:

[0020]

[0021] Among them, f b is the total force on the unmanned aerial vehicle under the rotor system, and m b is the total moment received. are respectively the components of the aerodynamic resultant force generated by the rotor along the x w y w z w three axes of the airflow coordinate system, m x , m y , m z are respectively the components of the aerodynamic resultant force generated by the rotor along the x of the airflow coordinate systemw y w z w Components of the three axes. f D , f Y , f L They are respectively the components of the total aerodynamic force exerted by the airflow on the cassette wing along the x-axis of the airflow coordinate system w y w z w Components of the three axes. L, M, and N are respectively the components of the total aerodynamic moment exerted by the airflow on the cassette wing along the x-axis of the airflow coordinate system w y w z w Components of the three axes.

[0022] Preferably, the kinematic equation of the UAV is:

[0023]

[0024] Wherein, p e and v e are respectively the position and velocity vectors of the UAV in the inertial system; ω b is the angular velocity of the UAV in the rotor coordinate system, [ω b x is the skew-symmetric matrix, m is the total mass of the UAV, J is the inertia matrix, f b is the total force on the UAV in the rotor coordinate system, m b is the total moment received.

[0025] Preferably, the calculation formula for the disturbing force is:

[0026]

[0027] The calculation formula for the disturbing moment is:

[0028] d2 = -J -1 ·ω b ×(J·ω b ),

[0029] Wherein, d1 is the disturbing force and d2 is the disturbing moment.

[0030] Compared with the prior art, the present invention has the following advantages and technical effects:

[0031] By installing a cassette wing on the fuselage and configuring multiple rotors on the cassette wing, the present invention provides thrust and necessary lift for the fuselage by the rotors, improving the anti-wind disturbance performance and stability of the UAV in a strong wind environment.

[0032] ​With a non-planar lifting surface design, the induced drag and structural weight are reduced. When encountering high-speed wind disturbances, the drone adjusts its pitch attitude, causing the box wing (non-planar lifting surface) to increase lift to compensate for and balance gravity, thereby unloading the coaxial rotors. At the same time, the propellers are used to enhance the forward wind resistance, effectively improving maneuverability and saving energy.

[0033] The tail wing adopts a V-shaped design, which while ensuring pitch stability, optimizes the response speed of attitude adjustment and improves the maneuverability of the drone. In addition, by designing a robust adaptive flight control law, the problem of cooperative control of heterogeneous control systems is solved. Compared with traditional drones, the drone of the present invention has better wind disturbance resistance performance, and the state conversion process is simpler, and it can achieve high-precision attitude adjustment and energy efficiency optimization under complex wind fields. Brief Description of the Drawings

[0034] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings:

[0035] Figure 1 Schematic diagram of the overall structure of the drone in the present invention;

[0036] Figure 2 Three-view and isometric view of the drone in the present invention;

[0037] Figure 3 Schematic diagram of the coordinate system and actuators of the present invention;

[0038] Figure 4 Exploded view of the tail wing of the present invention;

[0039] Figure 5 Schematic diagram of the control structure of the present invention;

[0040] Figure 6 Schematic diagram of a 10-level gradient wind field;

[0041] Figure 7 Response result diagram of the hovering attitude position control of the drone of the present invention against a 10-level gradient wind;

[0042] Figure 8 Response result diagram of the hovering attitude control of the drone of the present invention against a 10-level gradient wind;

[0043] Figure 9 Schematic diagram of a 10-level Dryden turbulent wind field;

[0044] Figure 10This is a graph showing the response results of the hovering attitude position control of the UAV of the present invention against level 10 turbulent wind;

[0045] Figure 11 This is a graph showing the hovering attitude control response results of the UAV of the present invention against level 10 turbulent wind.

[0046] Among them, 1. Fuselage; 2. Box wing; 3. Tail; 4. Rotor; 5. Propeller; 6. Connecting rod; 7. Power arm; 8. Motor. DETAILED DESCRIPTION

[0047] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0048] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0049] Reference Figures 1 to 11 , the present invention discloses a UAV capable of resisting strong wind disturbance, comprising a body and a driving mechanism arranged on the body;

[0050] The fuselage comprises a fuselage 1, two box-type wings 2 are symmetrically arranged on opposite sides of the fuselage 1, and a tail fin 3 is arranged at the tail end of the fuselage 1;

[0051] The drive mechanism includes a plurality of first drive members provided on the box-shaped wing 2 and a second drive member provided on the tail wing 3. The first drive members are in driving connection with the rotor 4, which provides lift and thrust for the fuselage 1. The second drive members are in driving connection with the propeller 5, which provides thrust for the fuselage 1.

[0052] The driving mechanism includes a multi-rotor working mode and a fixed-wing working mode. In the multi-rotor working mode, the first driving member controls the high-speed rotation of multiple rotors 4. In the fixed-wing working mode, the first driving member controls the low-speed rotation of multiple rotors 4, and the second driving member controls the rotation of the propeller 5.

[0053] The first driving member and the second driving member are both motors 8 .

[0054] In the present invention, a cassette wing 2 is installed on the fuselage 1, and a plurality of rotors 4 are arranged on the cassette wing 2. The axis of the fuselage 1 forms an angle of 20° with the reference plane of the rotor 4. In addition to providing thrust and necessary lift for the fuselage 1, the cassette wing 2 can efficiently provide partial lift (unable to balance gravity) in the presence of wind, improving the anti-wind disturbance performance and stability of the unmanned aerial vehicle in a strong wind environment.

[0055] By means of a non-planar lift surface design, the induced drag and structural weight are reduced. When encountering high-speed wind disturbances, the unmanned aerial vehicle adjusts its pitch attitude to increase the lift of the cassette wing (non-planar lift surface) to compensate for and balance part of the gravity, thereby unloading the coaxial rotor; meanwhile, the propeller 5 is used to enhance the forward anti-wind force, effectively improving the maneuverability and saving energy.

[0056] The tail fin 3 adopts a V-shaped design, which not only ensures pitch stability but also optimizes the response speed of attitude adjustment, improving the maneuverability of the unmanned aerial vehicle. In addition, by designing a robust adaptive flight control law, the problem of coordinated control of heterogeneous control systems is solved. Compared with traditional unmanned aerial vehicles, the unmanned aerial vehicle of the present invention has better anti-wind disturbance performance, and the state conversion process is simpler, and it can achieve high-precision attitude adjustment and energy efficiency optimization in a complex wind field.

[0057] In a further optimized solution, one end of a connecting rod 6 is fixedly connected to the tail end of the fuselage 1. The connecting rod 6 is smoothly transitioned with the tail end of the fuselage 1. The connecting rod 6 is arranged in a cylindrical shape. The tail fin 3 is fixedly connected to the outer side wall of the connecting rod 6. The tail fin 3 is located at one end of the connecting rod 6 away from the fuselage 1. The second driving member is fixedly installed at one end of the connecting rod 6 away from the fuselage 1.

[0058] In a further optimized solution, the tail fin 3 includes two wing pieces. Both wing pieces are fixedly connected to the outer side wall of the connecting rod and are symmetrically arranged. The two wing pieces are arranged in a V shape with the opening facing upward.

[0059] In a further optimized solution, a power arm 7 is fixedly connected to the cassette wing 2. The power arm 7 is arranged along the axis direction of the unmanned aerial vehicle. An included angle is provided between the power arm 7 and the cassette wing 2. Both ends of the power arm 7 penetrate through the cassette wing 2 and are respectively fixedly connected to the first driving member.

[0060] In a further optimized solution, the fuselage 1 is arc-shaped and transitioned towards both sides along its own axis. The connection between the fuselage 1 and the cassette wing 2 is smoothly transitioned.

[0061] In a further optimized solution, the top surface of the fuselage 1 and the top surface of the cassette wing 2 form a non-planar lift surface.

[0062] A control method for an unmanned aerial vehicle, based on an unmanned aerial vehicle resistant to strong wind disturbances, includes the following steps:

[0063] Based on a preset dynamic model and kinematic equation, an anti-wind flight control algorithm is obtained;

[0064] Based on the anti-wind control algorithm, a cascaded closed-loop control architecture is adopted to control the driving mechanism to work in the multi-rotor working mode or the fixed-wing working mode;

[0065] Predict the interference force and interference moment, and perform feedforward compensation on the anti-wind flight control algorithm based on the aerodynamic effect information of the box-wing and the aerodynamic effect information of the wind field.

[0066] For the further optimized solution, the dynamic model of the UAV is:

[0067]

[0068] where f b is the total force on the UAV in the rotor system, and m b is the total moment received. are the components of the aerodynamic resultant force generated by the rotor along the x w y w z w axes of the airflow coordinate system, and m x , m y , m z are the components of the aerodynamic resultant force generated by the rotor along the x w y w z w axes of the airflow coordinate system. f D , f Y , f L are the components of the total aerodynamic force exerted by the airflow on the box-wing along the x w y w z w axes of the airflow coordinate system, and L, M, N are the components of the total aerodynamic moment exerted by the airflow on the box-wing along the x w y w z w axes of the airflow coordinate system.

[0069] For the further optimized solution, the kinematic equation of the UAV is:

[0070]

[0071] where p e and v e are the position and velocity vectors of the UAV in the inertial system respectively; ω b is the angular velocity of the UAV in the rotor coordinate system, [ω b x is the skew-symmetric matrix, m is the total mass of the UAV, J is the inertia matrix, f b is the total force on the UAV in the rotor coordinate system, and m b is the total moment received.

[0072] ​For the further optimized solution, the calculation formula for the interference force is:

[0073]

[0074] The calculation formula for the interference torque is:

[0075] d2 = -J -1 ·ω b ×(J·ω b )

[0076] where d1 is the interference force and d2 is the interference torque.

[0077] One specific example:

[0078] Drone component settings: When the drone is dealing with a strong wind environment, it mainly relies on the fixed-wing mode to maintain a stable anti-wind attitude. Therefore, an airfoil design with a high lift-to-drag ratio is adopted to optimize its aerodynamic performance.

[0079] The box wing 2 needs to reduce the power demand of the rotor 4 through lift augmentation compensation under high anti-wind conditions, thereby reducing the size and weight of the drone. A relatively mature airfoil with a high lift-to-drag ratio is selected to design the box wing 2 to ensure that the overall size of the drone can meet the requirements of the ground support system under the premise of meeting the same lift requirements. The AH79-100B airfoil is selected as the airfoil of the box wing 2. When determining the wing loading, in order to maximize the anti-wind speed of the drone, the wing loading standard that can maximize the lift-to-drag ratio is selected. Theoretically, when the induced drag is equal to the zero-lift drag, the lift-to-drag ratio will reach the optimal state. Based on this principle, the following formula can be further derived:

[0080]

[0081] In the formula, C D0 is the zero-lift drag coefficient, e is the Oswald efficiency factor reflecting the magnitude of the induced drag, Q is the dynamic pressure, and A is the aspect ratio of the box wing 2. By combining these parameters with the relationship that the lift is equal to the gravity in the cruise state, the wing loading (unit: N / m 2 ) that maximizes the range can be derived as:

[0082]

[0083] The zero-lift drag coefficient C D0 is determined to be 0.65 according to the lift and drag characteristic curve of the selected airfoil, and the following formula is used to estimate the value of the Oswald efficiency factor e:

[0084] e = 1.78(1 - 0.045A 0.68 ) - 0.46;

[0085] In the design of the drone, the set maximum wind resistance speed is 28 m / s, and the takeoff weight is 100 kg. Combining the aspect ratio data, the specific value of the wing loading can be calculated. Subsequently, substituting this wing loading value into the relevant formula can solve for the area of the cassette wing 2, which is 1.75 m 2 , the aspect ratio of the cassette wing 2 is defined as the ratio between the span length of the cassette wing 2 and its average geometric chord length. Increasing the aspect ratio of the cassette wing 2 helps reduce the drag generated by lift. However, a large aspect ratio design will also increase the structural weight and weaken the crosswind resistance ability. The selected average aspect ratio is 3.88. Considering all the above parameters of the cassette wing 2, the design of the cassette wing 2 can be regarded as a trapezoidal structure. Under the size limitations imposed by the ground support system, the taper ratio of the cassette wing 2 is set to 4. By using the relevant calculation formula, the specific value of the wing root chord length can be obtained as 1.0 m:

[0086]

[0087] The wingtip chord length is 0.25 m, and the calculation formula is:

[0088] C tip = λC toor

[0089] The distance between the mean aerodynamic chord length of the cassette wing 2 and the fuselage 1 axis is 0.27 m, and the calculation formula is:

[0090]

[0091] The box-wing fuselage integration layout with high aerodynamic efficiency is adopted and optimized on this basis. To improve the wind resistance speed and cruise efficiency, an airfoil transition fuselage 1 design is adopted, and a certain contraction angle is set at the rear section of the fuselage 1 to achieve a smooth connection with the cylindrical connecting rod 6, thus endowing the UAV with a streamlined appearance similar to a cone. In the design of the electromechanical component loading part of the fuselage 1, emphasis is placed on improving the space utilization rate and optimizing the loading efficiency. Its aspect ratio is set to 4.25. At the same time, to facilitate the installation of the V-shaped tail fin 3 and reserve space for the installation of the pusher propeller 5, a fuselage 1 design with an aspect ratio of 10 is adopted. On the premise of ensuring that the fuselage 1 has sufficient volume, the loading area is fully integrated into the wide box-wing fuselage integration, making the entire UAV body an efficient lifting surface, effectively reducing the wetted area and interference drag, and thus capable of generating considerable lift. In the selection of the airfoil for the basic shape of the fuselage 1, the GOE383 airfoil with a relatively large thickness is adopted to meet the actual requirements of the payload volume. To more closely fit the design concept of the box-wing fuselage integration, a V-shaped tail fin 3 is selected, which realizes the functions of the horizontal stabilizer and vertical stabilizer in a traditional tail-fin aircraft with a reduced number of components. The V-tail is arranged at the tail of the fuselage 1 and forms a certain angle with the mid-vertical plane of the fuselage 1. This V-shaped tail fin 3 replaces the traditional vertical tail fin and horizontal tail fin with fewer stabilizer surfaces, thus greatly reducing the number of components and control surfaces, effectively reducing the wetted area and structural weight of the whole aircraft, and further reducing the overall structural weight and flight resistance of the UAV. In the fixed-wing mode, the tail fin 3 located at the tail of the fuselage 1 needs to be in a trim state with the mid-plane of the fuselage 1. Since the tail fin 3 does not need to generate additional aerodynamic force during horizontal flight, the NACA0012 symmetric airfoil is selected. When the UAV is in the fixed-wing mode, the tail fin 3 plays a stabilizing and controlling role. To ensure that the tail fin 3 can fully perform its functions, it must have sufficient vertical tail effect and horizontal tail effect, and both of these effects are closely related to the tail volume. The area required for the V-tail can be accurately determined through the calculation of the tail volume:

[0092]

[0093] In the formula, C v is the tail volume coefficient of the tail fin 3, L v is the moment arm length of the tail fin 3, S is the reference area of the box wing 2. The moment arm length of the tail fin 3 is set to 50% of the length of the fuselage 1. Through calculation in combination with the length data of the fuselage 1, the vertical tail area is obtained as 0.08 m 2 . The design scheme of removing ailerons and flaps is adopted. In the fixed-wing mode, the tail fin 3 will undertake the tasks of stabilizing and controlling pitch, roll and yaw. The aspect ratio of the tail fin 3 is selected as 1.93, the angle of the tail fin 3 is 90°, and the taper ratio is 1.65;

[0094] In the hovering condition, the unmanned aerial vehicle (UAV) is mainly divided into two working modes: calm wind and low wind speed environment, and medium to high wind speed environment. In the calm wind and low wind speed environment, the UAV usually adopts the multi-rotor mode. In this mode, the power provided by the high-speed rotation of the rotors 4 enables the UAV to achieve attitude adjustments such as pitch, yaw, and roll. While in the medium to high wind speed environment, the UAV switches to the fixed-wing mode. In this mode, the propeller 5 of the tail wing 3 combines with the power of the rotors 4 to jointly assist in the control and stability of the pitch and yaw attitudes of the UAV. At the same time, the box-wing 2 and the power of the rotors 4 cooperate with each other to ensure the control and stability of the UAV in the roll attitude. To meet the requirement of the cooperation between the box-wing 2 and the rotors 4 to ensure the control and stability of the UAV in the roll attitude, the rotors 4 are connected to the fuselage 1 through the power arms 7. The power arms 7 form an angle of 20° with the fuselage 1 in the pitch direction, enabling the rotors 4 to not only provide the necessary lift but also generate thrust, endowing the UAV with more excellent wind resistance performance. The aerodynamic layout of the tail wing 3 and the fuselage 1 is a crucial consideration factor, mainly focusing on two core points: the tail capacity of the tail wing 3 and the mutual interference between the tail wing 3 and the flow fields of other components. Therefore, the longitudinal position relationship of the tail wing 3 relative to the fuselage 1 becomes the focus of attention. To optimize this design, the method of increasing the chord length of the airfoil section of the fuselage 1 and simultaneously reducing its thickness is adopted. At the same time, a connecting rod 6 is set at the rear of the fuselage 1, and a certain contraction angle is set at the rear section of the fuselage 1, which is smoothly connected to the connecting rod 6, enabling the tail wing 3 to be installed on the fuselage 1, further reducing the wetted area of the tail wing 3 and reducing the resistance formed by the flow around the fuselage 1 to the tail surface. Although the increase in the angle of the tail wing 3 exacerbates the mutual interference between the fuselage 1 and the tail wing 3 to a certain extent, the increase in the area of the virtual vertical stabilizer effectively offsets this interference and continuously improves the course stability;

[0095] The UAV with a combined configuration of a box-shaped non-planar lift surface and rotors has multiple actuation methods such as rotor power, box-wing aerodynamic force, and propeller power, belonging to an aircraft with heterogeneous multi-control surfaces. The UAV is mainly divided into two parts: the box-wing and the four rotors 4. Therefore, the following coordinate systems need to be defined before control design: the inertial coordinate system F e , in the inertial coordinate system o e x e y e z e Among them, the axis o e z e vertically points to the ground, the axis o e x e points to a definite direction in the horizontal plane, and the direction of the axis o e y e is obtained according to the right-hand rule based on the other two axes. The initial position of the UAV is set as the coordinate origin oe, and this coordinate system is fixed during the flight; the box-wing coordinate system F l, the rotor coordinate system F b , the rotor coordinate system o b x b y b z b is fixed to the rotor 4 part of the UAV, and the centroid of the UAV is selected as the coordinate origin o b , the axis o b x b points to the nose direction in the symmetry plane of the rotor 4, and the axis o b z b is in the symmetry plane, with the direction downward and perpendicular to the axis o b x b , the axis o b y b is obtained by the right-hand rule. In this coordinate system, there is the definition of Euler angles and the rotation matrix R from the rotor coordinate system to the inertial system e b , the box-wing coordinate system F l , the box-wing coordinate system o l x l y l z l is fixed to the box-wing part of the UAV, and the coordinate origin o l is also set at the centroid of the UAV. The axis o l x l points to the nose direction in the symmetry plane of the box-wing, and the axis o l z l is in the symmetry plane, with the direction downward and perpendicular to the axis o l x l , the axis o l y l is obtained by the right-hand rule. The axis o l x l and the plane c b x b y b the included angle between them is the installation angle of the box-wing, denoted as k. The airflow coordinate system F w , the airflow coordinate system o w x w y w z w has its origin also at the centroid of the UAV. The axis o w x w is parallel to the airspeed vector. The axis o w z w is in the symmetry plane of the box-wing, with the direction downward and perpendicular to the axis o w x w , the axis o w y w is obtained by the right-hand rule. The angle of attack α is defined as the airspeed vector in the plane o l x l yl The angle between the upward projection and the axis o l x l is the angle of attack α, and the sideslip angle β is defined as the angle between the airspeed vector and the plane o l x l y l To transform the aerodynamic forces and moments in the coordinate systems F w and F l into the quadrotor coordinate system F b , the following two rotation matrices are defined:

[0096]

[0097] where λ = k - α. The aerodynamic forces and moments acting on the fixed structures such as the box wing 2 and the fuselage 1 by the airflow can be regarded as a whole, acting on the aerodynamic center, replaced by aerodynamic coefficients in the dynamic model, and substituting the prior data obtained from aerodynamic calculations through interpolation method in actual calculations. What needs to be considered emphatically in the dual-system of the UAV is the deflection of the tail wing 3 and the forces and moments generated by the second driving member. Assuming there is a wind speed e in the inertial system F The airspeed V, airspeed vector angle of attack α and sideslip angle β in the box wing system can be defined as:

[0098]

[0099]

[0100] The total aerodynamic force exerted by the airflow on the aircraft along the x w y w z w components of the three axes are f D 、f Y 、f L respectively, and the total moment components along each axis are L, M, N. Taking the lift corresponding to the axis as an example, the total lift includes three parts: the box wing 2, the fuselage 1, and the tail wing 3:

[0101]

[0102] The part of the box wing 2 and the fuselage 1 is expressed as the product of the lift coefficient C L , dynamic pressure Q and effective area S. The lift f L acts on the z w axis, and the moment arm between the acting point and the y w axis causes the aircraft to perform pitching motion, corresponding to the pitching moment M of the y w axis. Considering the acting force on the tail wing 3, the aerodynamic force received by the tail wing 3 during flight is perpendicular to the wing surface, and it is decomposed. Among them, C LN,α,β,δr and C LN,α,β,δlare the aerodynamic control derivatives related to the angle of attack α, sideslip angle β, and wing surface deflection angle δ r , δ l related, and C L,δr , C L,δl , C Y,δr , C Y,δl are the decomposed wing surface lift coefficient and wing surface side force coefficient. The aerodynamic coefficient decomposition formula of the tail wing 3 is as follows:

[0103]

[0104] Denote the upward deflection of the tail wing 3 as the positive direction, and the lift is upward. The total aerodynamic force of the box-wing system of the UAV is:

[0105]

[0106] The total aerodynamic moment is:

[0107]

[0108] where Q is the air dynamic pressure, S is the reference area of the fuselage 1 and the box wing 2, b is the span of the box wing 2, c is the average geometric chord length of the box wing 2, S v is the reference area of the V-shaped wing surface, b v is the effective moment arm of the V-shaped wing surface along the y w axis, c v is the effective moment arm of the V-shaped wing surface along the x w axis, C L , C D , C Y , C m , C n , C l are the dimensionless aerodynamic coefficients of the fixed structure part of the whole machine, C D,δr and C D,δl are the drag control derivatives, C n,δr and C n,δl are the yaw moment coefficients, related to the horizontal component of the aerodynamic force of the V-shaped wing surface, C l,δr , C l,δl , C m,δr and C m,δl are the roll moment and pitch moment coefficients, related to the vertical component of the aerodynamic force of the V-shaped wing surface. The thrust of the second driving part is modeled as:

[0109]

[0110] where, K tail is the thrust coefficient of the second driving part, ω tail is the rotational speed of the second driving part;

[0111] In the multi-rotor coordinate system, the thrust and torque of a single motor 8 are modeled as:

[0112]

[0113] where K T and K M are the thrust and torque coefficients of the motor 8 respectively, ω i is the rotational speed of the i-th motor 8, and the total lift and total side force received by the multi-rotor coordinate system are:

[0114]

[0115] Denote the distances from the front-side rotors 4 to the center of mass of the UAV as d1, and the angles with the positive x b axis of the rotor coordinate system as υ1. Denote the distances from the rear-side rotors 4 to the center of mass of the UAV as d2, and the angles with the negative x b axis of the rotor coordinate system as υ2. The total moment in the multi-rotor coordinate system:

[0116]

[0117] where k α is the anti-torque coefficient of the motor 8, T ij is the combined thrust generated by the ij-th motor 8. Thus, the dynamic model of the UAV can be established based on the forces and moments in the multi-rotor coordinate system and the forces and moments in the box-wing system:

[0118]

[0119] The motion of the UAV is formulated by Newton's equations and Euler's equations as:

[0120]

[0121] where p e and v e are the position and velocity vectors of the UAV in the inertial system respectively; ω e is the angular velocity of the UAV in the rotor coordinate system, [ω e x represents the skew-symmetric matrix; m is the total mass of the UAV, and J is the inertia matrix; f b is the total force on the UAV in the rotor coordinate system, m b is the total moment received. When conducting control design, considering that the angle of attack α and the sideslip angle β are difficult to obtain during actual flight, it is assumed that the angle of attack α≈k + θ and the sideslip angle β≈0. The actual aerodynamic influence is regarded as a disturbance during control design. The traditional dynamic model contains five control variables, namely the three-dimensional forces in the rotor coordinate system and the two Euler angles of pitch and roll. It is somewhat difficult to directly determine these five quantities based on the acceleration. First, select the force f rz ​and two Euler angles to uniquely determine the desired acceleration, and a desired yaw angle is designed separately to meet the wind resistance requirement, and the remaining side force f ry and aerodynamic effects are regarded as disturbances. This adopts the control idea of a quadrotor aircraft, but due to the existence of aerodynamic forces, the calculation method is different. Based on the above ideas, the system of the UAV is rewritten as follows:

[0122] p e =v e ,v e =u + g + d1

[0123]

[0124] where g is the gravitational acceleration, and d1 and d2 are the estimated disturbing force and the estimated disturbing torque respectively. The specific expressions of the control input u and the disturbances d1 and d2 are:

[0125]

[0126] d2 = -J -1 ·ω b ×(J·ω b )

[0127] The cascaded closed-loop control is a common control architecture for UAVs, generally consisting of an outer loop for controlling the position and an inner loop for controlling the attitude. The position controller calculates the desired acceleration after receiving the desired position, maps the required acceleration to the total thrust and attitude, the attitude command is sent to the inner loop to solve for the desired torque, and the thrust command is directly sent to the control allocation module. Finally, the control allocation algorithm coordinates and allocates the torque command from the inner loop and the force command from the outer loop to the motors 8;

[0128] Based on the control method of a UAV with a box-shaped non-planar lifting surface and rotor combination configuration, the aerodynamic effects of the box wing and the wind field are estimated to achieve the feedforward compensation of the control. A time-varying uncertainty and disturbance estimator is used to estimate the disturbing force and the disturbing torque. Based on the control method of the UAV, it is necessary to estimate the aerodynamic effects of the box wing and the wind field to achieve the feedforward compensation of the control. A time-varying uncertainty and disturbance estimator is used to estimate the disturbing force and the disturbing torque. It is necessary to achieve the estimation of the disturbance term d1 The disturbing force estimator based on the time-varying UDE is:

[0129]

[0130] where u0 is the controller under the undisturbed nominal system. For the attitude control link, there is a system:

[0131] ω b =J -1 ·m b +d2

[0132] It is necessary to realize the estimation of the interference term d2 The interference torque estimator based on time-varying UDE is as follows:

[0133]

[0134] where m b 0 is the controller under the non-disturbed nominal system. A separate yaw control is designed to achieve the control objective. Before giving the desired yaw angle of the UAV, the wind field is first judged. The aerodynamic interference and aerodynamic torque interference caused by the wind field on the UAV provide information about the wind direction to a certain extent. For the estimation of the interference force The horizontal disturbing force received by the UAV can be used as a reference for the desired yaw angle:

[0135]

[0136] where and are respectively the lateral disturbance and the forward disturbance in For the estimation of the interference torque The yaw moment is mainly caused by the wind field blowing the V-shaped tail surface, and it can itself be used as the angle error between the wind field vector and the box-wing system. According to this error, a dead zone is set to distinguish between windy and windless environments. In a windless scenario, the nose of the UAV should be in the same direction as the current airspeed vector to reduce the sideslip angle. In a windy environment, an upwind surface is actively formed according to the disturbance information. The desired yaw angle Ψ d can be obtained by the following formula similar to a PD controller:

[0137]

[0138] where k d1 and k d2 are respectively the weight coefficients of the yaw interference and the wind field thrust, and k ap and k ad are the control gain coefficients. Finally, the yaw angle Ψ d, it enters the angle control together with the position control output. A PID controller is used to achieve position control. By adjusting the control gain diagonal matrix, the position control accuracy of the UAV is ensured, and it is ensured that the position vector of the UAV in the inertial coordinate system can converge to the desired second-order differentiable position trajectory. First, the desired acceleration is calculated through the dynamic model of the UAV, and then this acceleration is mapped to the desired forces and desired attitudes of the three axes of the UAV. Finally, through the control allocation link, these forces and torques are allocated to specific actuators. The attitude error is represented by quaternions, and a corresponding controller is designed to ensure the uniform convergence of the attitude error. When flying at high speed, coordinated turning is also considered to reduce the sideslip angle and achieve fast and accurate tracking of the UAV's attitude, including the control of pitch angle, roll angle, and yaw angle. The attitude control is divided into an inner loop and an outer loop. The outer loop is responsible for position control, and the inner loop is responsible for attitude control. By solving the desired Euler angles and converting them into the desired angular velocities in the rotor coordinate system, and then the angular velocity error is eliminated through the angular velocity controller. Considering that the UAV has multiple actuators, the control allocation problem is formulated as an optimization problem to be solved, and the reasonable allocation of control quantities is achieved by minimizing the slack variables and weighted errors, and the virtual control expectations are allocated to specific actuators.

[0139] In the description of the present invention, it should be understood that the orientation or positional relationship indicated by the terms "longitudinal", "transverse", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings, and is only for the convenience of describing the present invention, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation to the present invention.

[0140] The embodiments described above are only descriptions of the preferred modes of the present invention, and do not limit the scope of the present invention. Without departing from the design spirit of the present invention, various deformations and improvements made by those of ordinary skill in the art to the technical solutions of the present invention shall fall within the protection scope determined by the claims of the present invention.

Claims

1. An unmanned aerial vehicle resistant to strong wind interference, characterized in that, It comprises a body and a driving mechanism arranged on the body; The fuselage comprises a fuselage (1), two box-type wings (2) are symmetrically arranged on opposite sides of the fuselage (1), and a tail wing (3) is arranged at the tail end of the fuselage (1); The driving mechanism comprises a plurality of first driving members arranged on the box-type wing (2) and a second driving member arranged on the tail wing (3), the first driving member being in driving connection with a rotor (4), the rotor (4) providing lift and thrust for the fuselage (1), and the second driving member being in driving connection with a propeller (5), the propeller (5) providing thrust for the fuselage (1); The driving mechanism includes a multi-rotor working mode and a fixed-wing working mode. In the multi-rotor working mode, the first driving member controls the plurality of rotors (4) to rotate at high speed. In the fixed-wing working mode, the first driving member controls the plurality of rotors (4) to rotate at low speed, and the second driving member controls the propeller (5) to rotate.

2. The anti-strong wind interference drone according to claim 1, characterized in that: The tail end of the fuselage (1) is fixedly connected to one end of a connecting rod (6), and the connecting rod (6) smoothly transitions to the tail end of the fuselage (1). The connecting rod (6) is configured to be cylindrical. The tail wing (3) is fixedly connected to the outer side wall of the connecting rod (6). The tail wing (3) is located at the end of the connecting rod (6) away from the fuselage (1). The second driving member is fixedly mounted at the end of the connecting rod (6) away from the fuselage (1).

3. The anti-strong-wind-disturbance drone according to claim 1, characterized in that: The tail wing (3) comprises two winglets, both of which are fixed to the outer side wall of the connecting rod (6) and are symmetrically arranged, and the two winglets are arranged in a V shape and opened upward.

4. The anti-strong wind interference UAV according to claim 1, characterized in that: A power arm (7) is fixedly connected to the box-type wing (2), and the power arm (7) is arranged along the axis direction of the drone. An angle is set between the power arm (7) and the box-type wing (2), and both ends of the power arm (7) pass through the box-type wing (2) and are respectively fixedly connected to the first driving member.

5. The drone resistant to strong wind disturbances according to claim 1, characterized in that: The fuselage (1) transitions in an arc shape to both sides along its own axis, and the connection between the fuselage (1) and the box-type wing (2) transitions smoothly.

6. The anti-strong wind interference UAV according to claim 1, wherein: The top surface of the fuselage (1) and the top surface of the box-type wing (2) are combined to form a non-planar lifting surface.

7. A method for controlling a drone against strong wind disturbances, based on the drone against strong wind disturbances according to any one of claims 1-6, characterized in that, The following steps are involved: Obtain wind-resistant flight control algorithm based on preset dynamic model and kinematic equations; Based on the wind resistance control algorithm, a cascade closed-loop control architecture is adopted to control the drive mechanism to operate in a multi-rotor working mode or a fixed-wing working mode; The interference force and interference torque are predicted, and the wind-resistant flight control algorithm is fed-forward compensated based on the box-wing aerodynamic effect information and the wind field aerodynamic effect information.

8. The anti-strong wind interference unmanned aerial vehicle control method according to claim 7, wherein The dynamic model of the UAV is: Among them, f b is the total force on the drone under the rotor system, and m b is the total moment received, are the components of the aerodynamic resultant force generated by the rotor along the x w y w z w axes of the airflow coordinate system, and m x , m y , m z are the components of the aerodynamic resultant force generated by the rotor along the x w y w z w axes of the airflow coordinate system, and f D , f Y , f L are the components of the total aerodynamic force exerted by the airflow on the box wing along the x w y w z w axes of the airflow coordinate system. L, M, and N are the components of the total aerodynamic moment exerted by the airflow on the box wing along the x w y w z w axes of the airflow coordinate system.

9. The anti-strong wind interference unmanned aerial vehicle control method according to claim 7, wherein The kinematic equation of the drone is: where p e and v e are the position and velocity vectors of the UAV in the inertial frame, respectively; ω b is the angular velocity of the UAV in the rotor coordinate system, [ω b x is the skew-symmetric matrix, m is the total mass of the UAV, J is the inertia matrix, f b is the total force acting on the UAV in the rotor coordinate system, and m b is the total moment received.​ 10. The anti-strong wind interference UAV control method according to claim 7, characterized in that The calculation formula of interference force is: The calculation formula of the interference torque is: d2 = -J -1 ·ω b ×(J·ω b ) Among them, d1 is the interference force and d2 is the interference torque.

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