Method and system for controlling ornithopter based on interference compensation

Abstract

The invention discloses an ornithopter control method and system based on interference compensation. The ornithopter control method comprises the following steps: respectively constructing state equations of an interference observer, an interference compensator, a target state controller and an attitude control system; estimating external interference in real time by using an interference observer based on the expected space attitude of the ornithopter to obtain an interference observation value, and inputting the interference observation value to an interference compensator; the interference compensator generates interference compensation according to the interference observation value; based on ideal dynamic output of an attitude control system, target control generated by a target state controller is adopted, the target state controller is designed in a layered mode, sliding mode control is added in the target state controller to enhance the suppression effect on external interference, and rapid and accurate tracking of system output on a target track can be achieved; interference compensation and target control are jointly input into an attitude control system to adjust the actual space attitude of the ornithopter, and control over the ornithopter is achieved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft control, and in particular relates to a flapping-wing aircraft control method and system based on interference compensation. Background Art

[0002] Bionic flapping-wing aircraft offer advantages such as low aerodynamic noise, strong maneuverability, high aerodynamic efficiency, and excellent stealth. Due to their unique design and versatility, their applications encompass military reconnaissance, environmental monitoring, urban management, search and rescue missions, and other scenarios. However, the aerodynamic characteristics generated by wing flapping are very complex, especially under low Reynolds numbers (typically less than 10,000). Therefore, understanding the aerodynamics of flapping wings is crucial for the design of flapping-wing aircraft. Currently, there are many design and control methods for flapping-wing aircraft both domestically and internationally, and their aerodynamic characteristics also vary significantly, resulting in the lack of a standardized and unified dynamic modeling solution for flapping-wing aircraft.

[0003] Aircraft control can generally be divided into position control and attitude control. Position control primarily aims to control the aircraft's spatial position and trajectory, typically involving adjustments to altitude, horizontal velocity, and flight path. Attitude control, on the other hand, requires precise adjustment of the aircraft's pitch, yaw, and roll angles to ensure a stable flight attitude. These two control mechanisms complement each other, collectively determining the aircraft's motion state in three-dimensional space. Because flapping-wing aircraft, as biomimetic aircraft that mimic the flight of birds or insects in nature, exhibit strong nonlinearities and coupling in their aerodynamic characteristics, motion patterns, and structural design, control strategies used for traditional fixed-wing or rotary-wing aircraft are difficult to directly apply. Summary of the Invention

[0004] The object of the present invention is to provide a flapping-wing aircraft control method and system based on interference compensation.

[0005] In a first aspect, the present invention provides a flapping-wing aircraft control method based on interference compensation, the method comprising:

[0006] The state equations of the disturbance observer, disturbance compensator, target state controller and attitude control system are constructed respectively; the target state controller is designed hierarchically and a sliding mode control is added to the target state controller. The state equation of the constructed target state controller is expressed as follows:

[0007]

[0008] Among them, u0 is the target control output by the target state controller; x1 is the attitude angle vector of the flapping-wing aircraft; x2 is the state vector related to the attitude angular velocity vector of the flapping-wing aircraft; g(x1) is the control gain coefficient; is the second derivative of the desired attitude angle vector of the flapping-wing aircraft; c1 is the control gain; c2 is the sliding hyperplane parameter; f(x1, x2) is the system function vector of the attitude control system; K is the switching gain parameter; s is the sliding surface function; z1 and z2 are tracking errors;

[0009] Based on the desired spatial posture of the flapping-wing aircraft, the interference observer is used to estimate the external interference in real time, obtain the interference observation value, and input it into the interference compensator; the interference compensator uses the interference observation value to calculate the external interference. Generate interference compensation u d ; Based on the ideal dynamic output of the attitude control system, the target control u0 is generated by the target state controller; the disturbance compensation u d Together with the target control u0, it is input into the attitude control system to adjust the actual spatial attitude of the flapping-wing aircraft and realize the control of the flapping-wing aircraft.

[0010] Preferably, the state equation of the attitude control system is expressed as:

[0011]

[0012] Among them, u is the control input; d is the external interference; y is the control output vector of the flapping-wing aircraft attitude control system.

[0013] Preferably, the disturbance observer adopts a nonlinear disturbance observer, and the state equation of the nonlinear disturbance observer is expressed as:

[0014]

[0015] Where z is the state variable of the nonlinear disturbance observer; is the disturbance observation value; x is the state variable; q(x) is the auxiliary function; G(x) is the gain function of the disturbance observer.

[0016] Preferably, the state equation of the interference compensator is expressed as:

[0017]

[0018] Among them, u d is the interference compensation output by the interference compensator.

[0019] Preferably, the method for constructing the state equation of the attitude control system is as follows: construct the dynamic differential equation and kinematic differential equation of the body and the kinematic equation of the wing respectively, and combine the above equations to construct the state equation of the attitude control system.

[0020] Preferably, the dynamic differential equation of the body is constructed based on the three-axis rotation angular velocity and torque of the flapping-wing aircraft relative to the inertial coordinate system.

[0021] Preferably, the torque includes wing aerodynamic torque, body aerodynamic torque and interference torque.

[0022] Preferably, the kinematic differential equation of the body is constructed based on the pitch angle, roll angle and yaw angle of the flapping-wing aircraft.

[0023] Preferably, the kinematic equation of the wing is constructed based on the twist angle and flapping angle of the wing.

[0024] In second aspect, the present invention provides a flapping-wing aircraft control system based on interference compensation, which is used to execute the above-mentioned flapping-wing aircraft control method; the flapping-wing aircraft control system includes an interference observer, an interference compensator, a target state controller and an attitude control system; the interference observer is used to estimate external interference and generate interference observation values; the interference compensator is used to generate interference compensation for the interference observation values; the target state controller is used to generate control inputs to achieve target trajectory tracking and desired attitude control of the flapping-wing aircraft; the attitude control system is used to adjust the spatial state of the flapping-wing aircraft according to the inputs of the interference compensator and the controller.

[0025] The present invention has the following beneficial effects:

[0026] 1. The present invention enhances the suppression effect of external interference by designing a target state controller in a hierarchical manner and adding sliding mode control to the target state controller, thereby enabling the system output to quickly and accurately track the target trajectory. At the same time, the present invention designs a nonlinear disturbance observer combined with a compensation control law, which can not only effectively cope with the nonlinear dynamic characteristics of flapping-wing aircraft, but also improve the system's fault tolerance under uncertain parameters and external disturbances, thereby optimizing the overall performance of the aircraft and providing a more stable and reliable control solution for flapping-wing aircraft in future diversified applications.

[0027] 2. The present invention uses a nonlinear disturbance observer to estimate the uncertainty of the system and external disturbances in real time, and uses the estimation results for compensation control to reduce the impact of disturbances on the system. This not only effectively suppresses the chattering problem of sliding mode control, but also enhances the robustness and anti-interference ability of the system under complex environmental changes, and has strong adaptability to model uncertainty and external disturbances. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 This is an overall flow chart of the flapping-wing aircraft control method of the present invention.

[0029] Figure 2 Schematic diagram of the undisturbed constant attitude pitch angle tracking curve of the present invention and the existing control method.

[0030] Figure 3 Schematic diagram of the undisturbed constant attitude yaw angle tracking curve of the present invention and the existing control method.

[0031] Figure 4 Schematic diagram of the undisturbed constant attitude roll angle tracking curve of the present invention and the existing control method.

[0032] Figure 5 Schematic diagram of attitude angle tracking error of undisturbed constant attitude of the present invention and the existing control method.

[0033] Figure 6 Schematic diagram of pitch angle tracking curve under pulse disturbance of the present invention and the existing control method.

[0034] Figure 7 Schematic diagram of yaw angle response curves under pulse disturbance of the present invention and the existing control method.

[0035] Figure 8 Schematic diagram of the roll angle response curve under pulse disturbance of the present invention and the existing control method. DETAILED DESCRIPTION

[0036] The present invention will be further described below with reference to the accompanying drawings.

[0037] like Figure 1 As shown, a flapping-wing aircraft control method based on interference compensation is adopted, and the flapping-wing aircraft control system adopted includes an interference observer, an interference compensator, a controller and an attitude control system; the interference observer is used to estimate external interference and generate interference observation values; the interference compensator is used to generate interference compensation for the interference observation values; the controller is used to generate control inputs for achieving target trajectory tracking and desired attitude control of the flapping-wing aircraft; the attitude control system is used to adjust the spatial state of the flapping-wing aircraft according to the inputs of the interference compensator and the controller to achieve control of the flapping-wing aircraft.

[0038] The flapping-wing aircraft control method comprises the following steps:

[0039] Step 1: Construct the dynamic equations and kinematic equations of the flapping-wing aircraft

[0040] 1-1. Construct the inertial coordinate system, body coordinate system, velocity coordinate system, path coordinate system, and left and right wing coordinate systems respectively. The inertial coordinate system takes the starting point of the flapping-wing aircraft at takeoff as the origin, the north direction as the x-axis, the direction perpendicular to the ground pointing to the sky as the y-axis, and the east direction as the z-axis; the body coordinate system takes the center of mass of the flapping-wing aircraft as the origin, the direction from the center of mass to the nose as the x-axis, and the direction directly above the body as the y-axis; the velocity coordinate system takes the center of mass of the flapping-wing aircraft as the origin, the instantaneous velocity direction of the flapping-wing aircraft as the x-axis, and the direction directly above the body as the y-axis; the path coordinate system takes the center of mass of the flapping-wing aircraft as the origin, the instantaneous velocity direction of the flapping-wing aircraft as the x-axis, and the direction directly above the body as the y-axis; the path coordinate system takes the center of mass of the flapping-wing aircraft in the inertial coordinate system plane x i o i z iThe projection in the plane is the origin, and the instantaneous velocity direction of the flapping-wing aircraft is in the plane x i o i z i The projection direction is the x-axis, and the direction perpendicular to the ground pointing to the sky is the y-axis; the wing coordinate system takes the root joint as the origin, the direction from the fuselage to the nose as the x-axis, and the direction perpendicular to the symmetry plane of the fuselage as the y-axis.

[0041] 1-2. Construct the following dynamic differential equation for the rotation of a flapping-wing aircraft around its center of mass in the body coordinate system:

[0042]

[0043] Among them, p, q, and r are the angular velocities of the flapping-wing aircraft relative to the xyz axis in the inertial coordinate system; are the derivatives of angular velocity p, q, and r respectively; I x ,I y ,I z are the moments of inertia of the flapping-wing aircraft around the x, y, and z axes respectively; M wx ,M wy ,M wz are the aerodynamic moments of the wing in the xyz axis directions respectively; M bx ,M by ,M bz are the aerodynamic moments of the body in the xyz axis directions respectively; M dx ,M dy ,M dz are the interference moments in the x, y, and z axis directions respectively.

[0044] Wing aerodynamic moment M w =[M wx M wy M wz ] T The calculation method is as follows:

[0045]

[0046] Among them, F lx ,F ly ,F lz are the aerodynamic forces generated by the left wing in the x, y, and z axes respectively; F rx ,F ry ,F rz are the aerodynamic forces generated by the right wing in the x, y, and z axes respectively; ξ is the distance between the aerodynamic center and the leading edge of the wing; ζ is the distance between the aerodynamic center and the wing root; and are the torsion angles of the left and right wings respectively; θ l and θ rare the flapping angles of the left and right wings respectively; η is the distance between the root joint of the wing and the center of mass of the flapping-wing aircraft.

[0047] In this embodiment, the distance ξ is one quarter of the chord length of the wing; and the distance ζ is 0.7 times the wingspan.

[0048] Airframe aerodynamic moment M b =[M bx M by M bz The calculation method is as follows:

[0049]

[0050] Where ρ is the air density; V is the forward flight speed of the aircraft in the velocity coordinate system; S t is the characteristic area of ​​the body, that is, the area of ​​the cross section of the body; l x ,l y ,l z are the characteristic lengths of the aircraft in the roll, pitch and yaw directions respectively; c x ,c y ,c z are the dimensionless drag coefficient, lift coefficient, and lateral force coefficient of the flapping-wing aircraft respectively; α is the flight attack angle of the aircraft.

[0051] 1-3. Construct the kinematic differential equation for the flapping-wing aircraft's rotation around its center of mass in the body coordinate system as follows:

[0052]

[0053] in, is the pitch angle; φ is the roll angle; ψ is the yaw angle; is the perturbation of the pitch angle; is the perturbation of the roll angle; is the perturbation of the yaw angle.

[0054] 1-4. Construct the kinematic equations of the flapping-wing aircraft wing in the body coordinate system as follows:

[0055]

[0056] θ l (t) = -θ L cos(2πft)-Δθ l

[0057] θ r (t) = θ R cos(2πft)+Δθ r

[0058] in, and are the torsion angles of the left and right wings respectively; θ l (t) and θ r (t) are the flapping angles of the left and right wings, respectively; and are the maximum values ​​of the torsion angles of the left and right wings respectively; θ L and θ R are the maximum flapping angles of the left and right wings respectively; Δθ l and Δθ r are the correction values ​​of the flapping angles of the left and right wings respectively; f is the frequency of the wing movement relative to the body.

[0059] In this embodiment, the maximum values ​​of the torsion angles of the left wing and the right wing are the same; the maximum values ​​of the flapping angles of the left wing and the right wing are the same.

[0060] Step 2: Combine the dynamic equations and kinematic equations of the flapping-wing aircraft and the kinematic equations of the wing to construct the state equation of the attitude control system:

[0061]

[0062] in, is the attitude angle vector of the flapping-wing aircraft; x2=g'(x1)ω; ω=[pqr] T is the attitude angular velocity vector of the flapping-wing aircraft; f(x1,x2) is the system function vector of the attitude control system; g(x1) is the control gain coefficient, g(x1)>0; u=[M wx M wy M wz ] T is the control input; d is the external disturbance, satisfying |d|≤D; d is the upper bound of the external disturbance; y is the control output vector of the flapping-wing aircraft attitude control system.

[0063] The system function vector f(x1,x2) and the control gain coefficient g(x1) are expressed as follows:

[0064]

[0065] in,

[0066] Step 3: Build the target state controller

[0067] The tracking errors z1 and z2 of the attitude control system are set as follows:

[0068]

[0069] in, is the desired attitude angle vector of the flapping-wing aircraft, that is, the desired spatial attitude; α1 is the virtual control quantity.

[0070] Derivative of the dynamic equation of tracking error z1, we get:

[0071]

[0072] In order to make the tracking error z1 converge, it is necessary to design a first-level virtual control law Where c1 is the control gain, which is a diagonal matrix and c1>0; the derivative of the tracking error obtained based on the tracking error z2 and the virtual control amount α1 is expressed as:

[0073]

[0074] When z2=0, Then the system is asymptotically stable. However, in general, z1≠0. Define the first-order Lyapunov function Taking the derivative and sorting it out we get:

[0075]

[0076] From the above results, we can see that the derivative of the first-order Lyapunov function contains the second-order error z2=x2-α. At this time, Negative definiteness cannot be guaranteed, so it is necessary to further design a two-stage controller to stabilize the tracking error z2. By deriving the dynamic equation of the tracking error z2, we obtain:

[0077]

[0078] Ignore the external disturbance d in the attitude control system and set the attitude control system Substituting the expression of into the above formula, the derivative of the tracking error is:

[0079]

[0080] In order to further stabilize the secondary error z2, the extended secondary Lyapunov function is defined Taking its derivative we get:

[0081]

[0082] Since flapping-wing aircraft controllers usually require real-time, high-precision adjustments, in order to achieve faster dynamic response and stronger anti-interference capability, and to be easy to implement, the following glide surface function s is designed:

[0083] s=z2+c2z1

[0084] Where c2={c21 ,c 22 ,c 23}>0, is the sliding hyperplane parameter, which together determines the shape of the sliding surface, affects the error convergence rate and the switching characteristics of the sliding mode control, and satisfies the Hurwitz stability condition. Derivative of the sliding surface function yields:

[0085]

[0086] To ensure Design equivalent control law u eq for:

[0087]

[0088] In order to ensure that the control system can quickly enter the sliding surface and stabilize on the sliding surface, the switching control law u is designed. sw To enhance the robustness to interference d and meet the reachability conditions, that is, Switching control law u sw As shown in the following formula:

[0089]

[0090] Where K is the switching gain parameter; sgn(s) is the switching function,

[0091] Based on the equivalent control law u eq and switching control law u sw Get the sliding mode control law u smc for:

[0092]

[0093] Based on the sliding mode control law smc , sliding surface function s and the derivative of the sliding surface function Will Expressed as:

[0094]

[0095] Where K = D + η.

[0096] Design of an extended three-level Lyapunov function Taking its derivative we get:

[0097]

[0098] According to the Lyapunov stability theorem, we know Therefore, the state equation of the designed target state controller is expressed as:

[0099]

[0100] Among them, u0 is the output of the target state controller.

[0101] Step 4: Construct a nonlinear disturbance observer

[0102] For the above second-order system, the linear disturbance observer is designed as follows:

[0103]

[0104] in, is the expected external interference of the flapping-wing aircraft, that is, the interference observation value; x = [x1, x2] T is the state variable; G(x) is the gain function of the disturbance observer; its value determines the speed of system convergence. Usually G(x) is selected to be greater than 0 and moderately large to ensure rapid convergence of the estimation error while avoiding excessive response of the observer.

[0105] The state variable z of the nonlinear disturbance observer is defined as:

[0106]

[0107] Where q(x) is an auxiliary function that satisfies the following dynamic equation:

[0108]

[0109] Taking the derivative of the state variable z, we get:

[0110]

[0111] In summary, the state equation of the constructed nonlinear disturbance observer is expressed as:

[0112]

[0113] It can be seen from the formula that the nonlinear disturbance observer does not need to directly obtain the acceleration signal Its implementation process is more in line with the actual needs of the system and has stronger adaptability.

[0114] Step 5: Build an Interference Compensator

[0115] In order to eliminate the external disturbance d in the dynamic equation of the whole machine, there is a state variable z→0 (error convergence) of the nonlinear disturbance observer, that is: Therefore, the disturbance can be estimated by using the nonlinear disturbance observer Estimate the external interference d and derive it to obtain:

[0116]

[0117] According to the composition of the controller, the control input u is composed of the target control u0 and the disturbance compensation u d It consists of two parts, and the control input u is substituted into the attitude control system From the dynamic equation of :

[0118]

[0119] According to the updated formula Expressed as:

[0120]

[0121] In order to achieve complete compensation of interference, that is, the influence of external interference d on the system is completely eliminated, it is necessary to satisfy: -g(x)u d +d=0, so we can get:

[0122]

[0123] That is, the state equation of the designed disturbance compensator is expressed as:

[0124]

[0125] In summary, it can be seen that the interference compensator interferes with the observation value Adjust the control input u d , so that the impact of external interference d on the system is weakened or even completely eliminated.

[0126] Step 6: Use a nonlinear disturbance observer to estimate the external disturbance in real time based on the desired spatial attitude of the flapping-wing aircraft to obtain the disturbance observation value. And input to the interference compensator; the interference compensator is based on the interference observation value Generate interference compensation u d , so that the influence of external disturbance on the system is weakened or even completely eliminated. The target state controller generates the target control u0 based on the ideal dynamic model of the attitude system (ignoring unknown disturbances and uncertainties); the disturbance compensation u d Together with the target control u0, it is input into the attitude control system to adjust the actual spatial attitude of the flapping-wing aircraft and realize the control of the flapping-wing aircraft.

[0127] Step 7: Effect Verification

[0128] 7-1. The present invention and the NDO-BC control method are used to perform undisturbed and disturbed simulation analysis on a flapping-wing aircraft, respectively. In the undisturbed simulation analysis, each simulation analysis is performed based on the time-varying nature of the target attitude (constant value or time-varying). In the disturbed simulation analysis, simulated external disturbances (pulse disturbances and time-varying disturbances) are introduced to evaluate the robustness and adaptability of the present invention and the NDO-BC control method in complex environments. In addition, during the simulation process, in addition to comparing the control effects of the two methods, various key performance indicators such as the system's response time, steady-state error, and the range of control input variation will be analyzed to ensure feasibility and effectiveness in actual engineering applications. The model parameters of the flapping-wing aircraft in the simulation analysis are shown in Table 1.

[0129] Table 1 Simulation parameters of flapping-wing aircraft

[0130]

[0131] Set the flapping-wing aircraft's flight speed to Air density Flapping frequency f = 12.5 Hz; flight attack angle α = 10°; θ L =θ R =60°;

[0132] 7-2. Simulation Analysis of Non-Disturbance Controller

[0133] Assume that the initial value of the attitude angle x1 is [0 0 0] T , the desired attitude angle y d [20° 20° 20°] T , and simulated the tracking performance of each controller's pitch angle, yaw angle, and roll angle, and the obtained tracking curves are as follows Figure 2 、 Figure 3 and Figure 4 As shown in the figure, when there is no disturbance and the attitude angle is a constant value, both control methods can effectively track the target attitude angle without overshoot and eventually reach a stable state, which shows that both control methods have high accuracy in target tracking. However, in terms of attitude angle control effect, the present invention is better than the NDO-BC control method, especially in terms of roll angle tracking. Compared with the NDO-BC control method, the present invention is closer to the target value and has smaller error fluctuations, which proves that the present invention has a slight advantage over the NDO-BC control method in tracking accuracy and stability. It also proves that the target state controller can track the target attitude angle more accurately.

[0134] To more clearly compare the tracking performance of the two control methods, we used the key performance metric, root mean square error (RMSE), for quantitative analysis. This RMSE is often used to evaluate the tracking accuracy of an algorithm. The RMS attitude angle errors of each control algorithm are shown in Table 2.

[0135] Table 2 RMS error of attitude angle at undisturbed constant attitude

[0136]

[0137]

[0138] As can be seen from Table 2, although the RMSE values ​​of the attitude angles of the two control methods are not much different, the RMSE values ​​of the present invention are all smaller than those of the NDO-BC control method, indicating that the tracking effect of the present invention for the three attitude angles is still better than that of the NDO-BC control method, with smaller errors.

[0139] When there is no disturbance and the target attitude is a constant value, the attitude angle tracking errors of each controller are as follows: Figure 5 As shown. Figure 5 As can be seen from the simulation results, the tracking error of the NDO-BC control method is slightly larger in the initial stage, and the convergence speed is slower, demonstrating that its interference compensation capability is slightly inferior to that of the present invention. Ideally, the tracking error should approach zero, indicating that the control system has adjusted the attitude angle to the desired value. Simulation results show that both control methods can bring the attitude angle error close to zero, but the present invention achieves a smaller error and a faster convergence speed. Therefore, the present invention outperforms the NDO-BC control method in terms of control accuracy and response speed.

[0140] 7-3. Simulation Analysis of Controller with Disturbance

[0141] In actual flight missions, flapping-wing aircraft are often affected by a variety of external disturbances, such as airflow changes and wind speed fluctuations. These disturbances may not only cause instantaneous impacts on the attitude and heading control of the aircraft, but may also cause the system performance to gradually decline, especially during long-term flight. Therefore, in order to verify the robustness and adaptability of the control algorithm in complex dynamic environments, simulation analysis is carried out by introducing pulse disturbances and time-varying disturbances, exploring the impact of different types of disturbances on the attitude control system, and evaluating the effectiveness of the two control methods in resisting interference. Assume that the initial value of the attitude angle of the flapping-wing aircraft is 0rad, and the desired attitude angle is [20° 20° 20°] T .

[0142] When the flapping-wing aircraft flies for 5 seconds, a pulse disturbance with an amplitude of 100 N.m and a duration of 1 second is applied simultaneously in the pitch, yaw and roll directions. The attitude angle tracking curves of each controller under the pulse disturbance are shown as follows: Figure 6 、 Figure 7 and Figure 8 The root mean square error of the attitude angle is shown in Table 3.

[0143] Table 3 RMS attitude angle error under pulse disturbance

[0144]

[0145] from Figure 6 、 Figure 7 and Figure 8 As can be seen in Table 3, the deviation amplitude of the present invention under interference is significantly smaller than that of the NDO-BC control method, demonstrating that the present invention has a stronger ability to suppress large-amplitude pulse interference. After being subjected to pulse interference at 5 seconds, the present invention is able to recover to the target attitude more quickly, with a faster response speed and shorter adjustment time. The present invention has a smaller deviation from the target value during the steady-state phase and a smaller steady-state error, indicating that the algorithm has higher control accuracy. In addition, the NDO-BC control method has a more significant overshoot, with larger deviations and larger oscillations after being subjected to interference. In summary, it can be concluded that under large-amplitude (100 N.m) pulse interference, the present invention has stronger anti-interference ability, faster recovery speed, smaller steady-state error, and lower overshoot. Therefore, the control performance of the present invention is significantly superior to the NDO-BC control method.

Claims

1. A flapping-wing aircraft control method based on interference compensation, characterized in that: The method comprises: The state equations of the disturbance observer, disturbance compensator, target state controller and attitude control system are constructed respectively; the target state controller is designed hierarchically and a sliding mode control is added to the target state controller. The state equation of the constructed target state controller is expressed as follows: Among them, u0 is the target control output by the target state controller; x1 is the attitude angle vector of the flapping-wing aircraft; x2 is the state vector related to the attitude angular velocity vector of the flapping-wing aircraft; g(x1) is the control gain coefficient; is the second derivative of the desired attitude angle vector of the flapping-wing aircraft; c1 is the control gain; c2 is the sliding hyperplane parameter; f(x1, x2) is the system function vector of the attitude control system; K is the switching gain parameter; s is the sliding surface function; z1 and z2 are tracking errors; Based on the desired spatial posture of the flapping-wing aircraft, the interference observer is used to estimate the external interference in real time, obtain the interference observation value, and input it into the interference compensator; the interference compensator uses the interference observation value to calculate the external interference. Generate interference compensation u d ; Based on the ideal dynamic output of the attitude control system, the target control u0 is generated by the target state controller; the disturbance compensation u d Together with the target control u0, it is input into the attitude control system to adjust the actual spatial attitude of the flapping-wing aircraft and realize the control of the flapping-wing aircraft.

2. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The state equation of the attitude control system is expressed as: Among them, u is the control input; d is the external interference; y is the control output vector of the flapping-wing aircraft attitude control system.

3. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The disturbance observer adopts a nonlinear disturbance observer, and the state equation of the nonlinear disturbance observer is expressed as: Where z is the state variable of the nonlinear disturbance observer; is the disturbance observation value; x is the state variable; q(x) is the auxiliary function; G(x) is the gain function of the disturbance observer.

4. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The state equation of the interference compensator is expressed as: Among them, u d is the interference compensation output by the interference compensator.

5. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The method for constructing the state equation of the attitude control system is as follows: construct the dynamic differential equation and kinematic differential equation of the body and the kinematic equation of the wing respectively, and combine the above equations to construct the state equation of the attitude control system.

6. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The dynamic differential equations of the body are constructed based on the three-axis rotation angular velocity and torque of the flapping-wing aircraft relative to the inertial coordinate system.

7. The flapping-wing aircraft control method based on interference compensation according to claim 6, characterized in that: The torque includes wing aerodynamic torque, body aerodynamic torque and interference torque.

8. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The kinematic differential equations of the body are constructed based on the pitch angle, roll angle and yaw angle of the flapping-wing aircraft.

9. The flapping-wing aircraft control method based on interference compensation according to claim 1, characterized in that: The kinematic equations of a wing are constructed based on the twist and flap angles of the wing.

10. A flapping-wing aircraft control system based on interference compensation, characterized in that: Used to execute the flapping-wing aircraft control method based on interference compensation as described in claim 1; the flapping-wing aircraft control system includes a disturbance observer, a disturbance compensator, a target state controller and an attitude control system; the disturbance observer is used to estimate external disturbances and generate disturbance observations; the disturbance compensator is used to generate disturbance compensation based on the disturbance observations; the target state controller is used to generate control inputs for achieving target trajectory tracking and desired attitude control of the flapping-wing aircraft; The attitude control system is used to adjust the spatial state of the flapping-wing aircraft based on the input from the disturbance compensator and the controller.

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