Flight control method and device for folding-wing variant aircraft

Abstract

The present application provides a flight control method and device for a folding-wing morphing aircraft, belonging to the technical field of flight control systems. The method determines an initial dynamic configuration based on the initial static configuration of the folding-wing morphing aircraft, estimates a variable configuration disturbance term in the initial dynamic configuration based on the aircraft's current and previous state signals, obtains a compensation control law instruction based on the variable configuration disturbance term and the pilot's current control command signal, obtains a main control law instruction based on the aircraft's current state signal and the pilot's current control command signal, adds the compensation control law instruction to the main control law instruction, obtains a flight control instruction, and sends the flight control instruction to the servo. This method achieves configuration recognition and stable control during the deformation process, and has the characteristics of fast response, good robustness, high accuracy, and wide applicability.

Description

Technical Field

[0001] The present application belongs to the technical field of flight control systems, and in particular relates to a flight control method and device for a folding-wing variant aircraft. Background Art

[0002] When a folding-wing morphing aircraft's wings fold or unfold, changes in the aircraft's overall structure affect its aerodynamic and mass characteristics, resulting in time-varying dynamics in the aircraft's dynamics model and highly nonlinear kinematic characteristics. Modeling this nonlinear process is challenging, requiring extensive wind tunnel test data and mass center of mass change data. This is complex and often carries significant errors. Folding-wing morphing aircraft experience significant changes in pitch attitude, high speed, and velocity after folding or unfolding their wings, making them unable to rely on their inherent stability to maintain flight during the deformation process. A US folding-wing morphing aircraft previously failed flight tests due to the dramatic changes in its aerodynamic and mass characteristics during wing deformation, preventing the control system from quickly and effectively compensating for the changes and ensuring stable flight. Understanding how to maintain stable flight during the deformation process is a primary concern in the research and application of folding-wing morphing aircraft.

[0003] In related technologies, PID controllers are often used to ensure that folding-wing variant aircraft maintain stable flight during the deformation process. However, these controllers rely on the accuracy of the aircraft model. When faced with folding-wing variant aircraft with large modeling errors and complex aircraft parameter changes, they cannot provide fast-response and high-precision control effects, and cannot guarantee stable flight control effects. Summary of the Invention

[0004] In order to solve the problem in the related art that the PID controller cannot provide a fast response and high precision control effect and cannot ensure a stable flight control effect, the present invention provides a flight control method and device for a folding wing variant aircraft. The technical solution is as follows:

[0005] In a first aspect, a flight control method for a folding-wing deformable aircraft is provided, comprising:

[0006] Determine the initial dynamic configuration according to the initial static configuration of the folding wing variant aircraft, and estimate the configurational disturbance term in the initial dynamic configuration according to the state signals of the aircraft at the current and previous beats;

[0007] According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained;

[0008] According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained;

[0009] The compensation control law command and the main control law command are added to obtain the flight control command, and the flight control command is sent to the servo.

[0010] The initial dynamic configuration of the folding-wing morphing aircraft is determined according to the initial static configuration of the folding-wing morphing aircraft, including:

[0011] A reference system state equation is obtained according to the initial static configuration of the folding wing variant aircraft, wherein the reference system state equation includes a state vector, an input vector, a reference system state matrix and a reference system input matrix;

[0012] Add three disturbance terms, namely state disturbance, input disturbance and constant disturbance, to the reference system state equation to construct the real system state equation of the current aircraft. The state disturbance term is multiplied by the state vector to obtain a first result, and the input disturbance term is multiplied by the input vector to obtain a second result. The first result, the second result and the constant disturbance term are summed and then added to the input vector.

[0013] The real system state equation is used as the initial dynamic configuration.

[0014] Among them, the configuration-varying disturbance term in the initial dynamic configuration is estimated based on the state signals of the aircraft in the current and previous cycles, including:

[0015] Multiply the aircraft's current state signal, the difference between the current state signal and the previous state estimate, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix. Project the multiplication result onto the previous state disturbance estimate as the current state disturbance estimate. The state estimate signal is obtained by inputting the flight control command into the reference system state equation.

[0016] Multiply the aircraft's current state signal, the pilot's current control command signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the input disturbance estimation term of the previous cycle as the current cycle input disturbance estimation term;

[0017] Multiply the difference between the aircraft's current state signal and the previous state estimation signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous constant disturbance estimation term as the current constant disturbance estimation term.

[0018] Among them, according to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained, including:

[0019] Determining the sum of the product of the current beat input disturbance estimation term and the current beat control command signal of the pilot, the product of the current beat state disturbance estimation term and the current beat state signal of the aircraft, and the constant disturbance estimation term;

[0020] The product of the preset reference matrix command gain and the driver's current control command signal is subtracted from the sum value to obtain an unfiltered compensation control law command;

[0021] The unfiltered compensation control law instruction is passed through a low-pass filter to obtain the compensation control law instruction.

[0022] Among them, according to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained, including:

[0023] The result of multiplying the aircraft's current state signal by the full-state feedback gain matrix is ​​used as the full-state feedback signal;

[0024] The result of multiplying the driver's current control command signal by the control gain matrix is ​​used as the main control law control signal;

[0025] The full-state feedback signal is summed with the main control law manipulation signal to obtain the main control law instruction.

[0026] The status signal data includes: airspeed, acceleration, altitude, airflow angle, attitude angle, attitude angle rate, attitude angle change rate and three-axis overload.

[0027] In a second aspect, a flight control device for a folding-wing variant aircraft is provided, comprising:

[0028] Identify modules for:

[0029] Determine the initial dynamic configuration according to the initial static configuration of the folding wing variant aircraft, and estimate the configurational disturbance term in the initial dynamic configuration according to the state signals of the aircraft at the current and previous beats;

[0030] According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained;

[0031] According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained;

[0032] The sending module is used to add the compensation control law instruction and the main control law instruction to obtain the flight control instruction, and send the flight control instruction to the servo.

[0033] The determination module is specifically used to:

[0034] A reference system state equation is obtained according to the initial static configuration of the folding wing variant aircraft, wherein the reference system state equation includes a state vector, an input vector, a reference system state matrix and a reference system input matrix;

[0035] Add three disturbance terms, namely state disturbance, input disturbance and constant disturbance, to the reference system state equation to construct the real system state equation of the current aircraft. The state disturbance term is multiplied by the state vector to obtain a first result, and the input disturbance term is multiplied by the input vector to obtain a second result. The first result, the second result and the constant disturbance term are summed and then added to the input vector.

[0036] The real system state equation is used as the initial dynamic configuration.

[0037] The determination module is specifically used to:

[0038] Multiply the aircraft's current state signal, the difference between the current state signal and the previous state estimate, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous state disturbance estimate as the current state disturbance estimate.

[0039] Multiply the aircraft's current state signal, the pilot's current control command signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the input disturbance estimation term of the previous cycle as the current cycle input disturbance estimation term;

[0040] Multiply the difference between the aircraft's current state signal and the previous state estimation signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous constant disturbance estimation term as the current constant disturbance estimation term.

[0041] The determination module is specifically used to:

[0042] Determining the sum of the product of the current beat input disturbance estimation term and the current beat control command signal of the pilot, the product of the current beat state disturbance estimation term and the current beat state signal of the aircraft, and the constant disturbance estimation term;

[0043] The product of the preset reference matrix command gain and the driver's current control command signal is subtracted from the sum value to obtain an unfiltered compensation control law command;

[0044] The unfiltered compensation control law instruction is passed through a low-pass filter to obtain the compensation control law instruction.

[0045] The present invention proposes a flight control method and device for a folding-wing variant aircraft. The method and device are designed to configure a characteristic structure with main control law instructions and compensation control law instructions, identify the aircraft state online, and perform compensation control according to a set reference system, thereby achieving stable flight control during the wing folding process and ensuring that the real-time state of the aircraft is close to the ideal state. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Figure 1A schematic diagram illustrating the principle of the flight control method for a folding-wing variant aircraft provided in an embodiment of the present application;

[0047] Figure 2 A schematic diagram illustrating the flight process effect of a flight control method for a folding-wing variant aircraft provided in an embodiment of the present application;

[0048] Figure 3 A schematic diagram of the flight state effect of a flight control method for a folding-wing variant aircraft provided in an embodiment of the present application;

[0049] Figure 4 A schematic diagram of the wing folding and flight phases of a flight control method for a folding-wing variant aircraft provided in an embodiment of the present application shows the flight state effects. DETAILED DESCRIPTION

[0050] The present application is further described in detail below through specific implementation methods and drawings.

[0051] See Figure 1 , which is a schematic diagram of the principle of the flight control method for a folding-wing variant aircraft provided in this application, the method comprising:

[0052] Step 1: The aircraft's current state signal x k , multiplied by the previous state estimation signal The difference Multiply the solution P of the Lyapunov equation of the reference system state matrix by the reference system input matrix b, and then move the result up to the state disturbance estimate term Projection, and then multiply by the preset adaptive gain Γ to get the current state disturbance estimation term Among them, the initial state disturbance estimation term Set as needed. The state estimation signal is obtained by inputting the flight control command into the reference system state equation.

[0053] Step 2: The aircraft's current state signal x k , minus the previous state estimation signal Then multiply it by the current control command signal r of the aircraft pilot k , multiplied by the solution P of the Lyapunov equation of the reference system state matrix, multiplied by the reference system input matrix b, the multiplication result is the input disturbance estimation term of the previous beat Projection, get the current shot input disturbance estimation term Among them, the initial input disturbance estimation term Set up as needed.

[0054] Step 3: Based on the current state signal x of the aircraft k , minus the previous state estimation signal Multiply the solution P of the Lyapunov equation of the reference system state matrix and the reference system input matrix b, and the disturbance estimate is projected onto the constant disturbance estimate of the previous beat. Get the current beat constant disturbance estimate Among them, the initial constant disturbance estimate and the state estimation signal Set up as needed.

[0055] Step 4: Set the preset reference matrix instruction gain k g The control command signal r currently taken by the aircraft pilot k The product of the current beat input disturbance estimation term The product of the current control command signal r of the aircraft pilot and the current state disturbance estimation term The current status signal x of the aircraft k The product of and the constant perturbation estimate of the current beat The sum of the three is the unfiltered compensation control law command.

[0056] Step 5: The unfiltered compensation control law command is passed through a low-pass filter D(z) to obtain the compensation control law command.

[0057] This application designs the robust conditions of the compensation controller based on the L1 adaptive control method, simplifies the complexity of the control system, and improves system performance.

[0058] Step 6: Based on the current state signal x of the aircraft k , multiplied by the full-state feedback gain matrix K, to obtain the full-state feedback signal.

[0059] Step 7: Based on the control command signal r currently taken by the aircraft pilot k , multiplied by the control gain matrix G, to obtain the main control law control signal.

[0060] Step 8: Sum the full-state feedback signal and the main control law manipulation signal to obtain the main control law instruction.

[0061] Step 9: The flight control command is obtained by summing the compensation control law command and the main control law command.

[0062] Figures 2 to 4 This is a schematic diagram of the flight control effect of a folding wing variant aircraft obtained by the method provided in this application, wherein: Figure 2 This is the flight process effect. The XYZ axes represent the displacement of the aircraft relative to the ground origin, in meters. It can be seen that the aircraft's flight process is very stable without any unexpected movement. Figure 3is the flight status effect, which are flight altitude HB, unit: m, flight speed VIAS, unit: m / s, aircraft yaw angle YAW, unit: °, aircraft angle of attack AOA, unit: °, aircraft longitudinal overload NZ, unit: g, aircraft pitch angle, unit: °, aircraft sideslip angle AOS, unit: °, aircraft lateral overload NY, unit: g, aircraft roll angle, unit: °, aircraft pitch angular rate ROP, unit: ° / s, aircraft roll angular rate ROR, unit: ° / s, aircraft yaw angular rate ROY, unit: ° / s. The horizontal axes are all flight time t, unit: s. It can be seen that the flight status of the aircraft changes smoothly and rapidly, with small overshoot and error, and can achieve the expected effect. Figure 4 The folding angle of the aircraft's wings is ftheta, in degrees. The flight stages Fstage, from 1 to 7, represent climb, transition to level flight, left roll, right roll, jump, descent, and yaw. It can be seen that the aircraft maintains level flight and deforms after reaching the specified speed and altitude. The wings fold from 0° to 120° when the flight time is 120s to 140s.

[0063] The flight control method for a folding-wing variant aircraft provided in this application has the following innovations:

[0064] This flight control method for a folding-wing morphing aircraft only requires aerodynamic modeling of the statically configured folding-wing morphing aircraft, eliminating the need for aerodynamic modeling of the dynamic wing deformation process. This significantly reduces the burden of aerodynamic modeling and avoids the reduction in control accuracy caused by deviations in aerodynamic modeling during the dynamic process.

[0065] The flight control method for a folding-wing variant aircraft controls the folding process using only the airspeed, angle of attack, pitch angle, and pitch rate feedback signals required by conventional flight control systems. It eliminates the need for additional folding angle and wing folding rate feedback signals that characterize the dynamic process of wing deformation, simplifying the structure of the flight control system.

[0066] This flight control method for a folding-wing morphing aircraft uses an eigenstructure configuration method as the primary control law. This method allows the eigenstructure of the folding-wing morphing aircraft to be configured as needed. This method not only conveniently considers the robustness design indicators and flight quality requirements of the system itself, but also ensures the robustness of the compensation controller designed using the L1 adaptive control method, thus simplifying the complexity of the control system and improving system performance.

[0067] This flight control method for a folding-wing morphing aircraft converts the aircraft's aerodynamic and mass characteristics during the wing-folding dynamic process into three disturbance terms in the aircraft's state equation. This method not only fully defines the changes in the aircraft's aerodynamic and mass characteristics during the wing-folding dynamic process, but also encompasses the changes in the aerodynamic and mass characteristics of other types of morphing aircraft. It is universal and can be extended to other morphing aircraft or aircraft with varying aerodynamic and mass characteristics.

[0068] This flight control method for a folding-wing variant aircraft utilizes the state equations of a fixed-configuration aircraft to design a reference system. This method eliminates the need for mathematical modeling of the wing folding dynamic process of the folding-wing variant aircraft, does not overly rely on the modeling accuracy of the controlled object, and does not require controller design for the wing folding process. This method simplifies the complexity of the control system and ensures that the aircraft's flight performance closely matches that of the reference system.

[0069] This flight control method for a folding-wing variant aircraft compensates for differences between the real aircraft system and the reference system through the L1 adaptive law. Besides providing conventional stable control during deformation, it also enhances its adaptability to unexpected disturbances such as wind disturbances and unmodeled modes. The system also possesses command tracking capability with zero static error.

[0070] The control law of the folding-wing variant aircraft flight control method avoids the high-frequency changes of the control signal caused by rapid adaptation in conventional adaptive control methods by designing a low-pass filter, which does not affect the robust performance of the system.

[0071] The flight control method for the folding-wing variant aircraft adopts the L1 adaptive control method to design a compensation controller. This method has the characteristics of identifying and compensating nonlinear links, such as the rate limit, position limit and dead zone of the servo, and can improve the flight performance of the folding-wing variant aircraft.

[0072] The present application also provides a flight control device for a folding-wing variant aircraft, comprising:

[0073] Identify modules for:

[0074] Determine the initial dynamic configuration according to the initial static configuration of the folding wing variant aircraft, and estimate the configurational disturbance term in the initial dynamic configuration according to the state signals of the aircraft at the current and previous beats;

[0075] According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained;

[0076] According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained;

[0077] The sending module is used to add the compensation control law instruction and the main control law instruction to obtain the flight control instruction, and send the flight control instruction to the servo.

[0078] The determination module is specifically used to:

[0079] A reference system state equation is obtained according to the initial static configuration of the folding wing variant aircraft, wherein the reference system state equation includes a state vector, an input vector, a reference system state matrix and a reference system input matrix;

[0080] Add three disturbance terms, namely state disturbance, input disturbance and constant disturbance, to the reference system state equation to construct the real system state equation of the current aircraft. The state disturbance term is multiplied by the state vector to obtain a first result, and the input disturbance term is multiplied by the input vector to obtain a second result. The first result, the second result and the constant disturbance term are summed and then added to the input vector.

[0081] The real system state equation is used as the initial dynamic configuration.

[0082] The determination module is specifically used to:

[0083] Multiply the aircraft's current state signal, the difference between the current state signal and the previous state estimate, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous state disturbance estimate as the current state disturbance estimate.

[0084] Multiply the aircraft's current state signal, the pilot's current control command signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the input disturbance estimation term of the previous cycle as the current cycle input disturbance estimation term;

[0085] Multiply the difference between the aircraft's current state signal and the previous state estimation signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous constant disturbance estimation term as the current constant disturbance estimation term.

[0086] The determination module is specifically used to:

[0087] Determining the sum of the product of the current beat input disturbance estimation term and the current beat control command signal of the pilot, the product of the current beat state disturbance estimation term and the current beat state signal of the aircraft, and the constant disturbance estimation term;

[0088] The product of the preset reference matrix command gain and the driver's current control command signal is subtracted from the sum value to obtain an unfiltered compensation control law command;

[0089] The unfiltered compensation control law instruction is passed through a low-pass filter to obtain the compensation control law instruction.

[0090] The flight control method and device for a folding-wing morphing aircraft proposed in this paper can online identify the aircraft's configuration state and rapidly compensate for it, ensuring both aircraft stability and performance close to that of a reference system. Furthermore, this method does not rely on wing deformation and aerodynamic modeling, is unaffected by model accuracy, and can reduce the burden of dynamic process modeling and wind tunnel testing, thus possessing significant engineering application value.

[0091] The above merely describes the embodiments of the present application, and while the description is relatively specific and detailed, it should not be construed as limiting the scope of the invention. It should be noted that a person skilled in the art could make various modifications and improvements without departing from the spirit of the present application, all of which fall within the scope of protection of the present application. Furthermore, any portions not described in detail herein are conventional techniques.

Claims

1. A flight control method for a folding-wing deformable aircraft, characterized in that: include: Determine the initial dynamic configuration according to the initial static configuration of the folding wing variant aircraft, and estimate the configurational disturbance term in the initial dynamic configuration according to the state signals of the aircraft at the current and previous beats; According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained; According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained; Add the compensation control law command and the main control law command to obtain the flight control command, and send the flight control command to the servo; Determine the initial dynamic configuration based on the initial static configuration of the folding wing variant aircraft, including: A reference system state equation is obtained according to the initial static configuration of the folding wing variant aircraft, wherein the reference system state equation includes a state vector, an input vector, a reference system state matrix and a reference system input matrix; Add three disturbance terms, namely state disturbance, input disturbance and constant disturbance, to the reference system state equation to construct the real system state equation of the current aircraft. The state disturbance term is multiplied by the state vector to obtain a first result, and the input disturbance term is multiplied by the input vector to obtain a second result. The first result, the second result and the constant disturbance term are summed and then added to the input vector. The real system state equation is used as the initial dynamic configuration; The configuration-varying disturbance term in the initial dynamic configuration is estimated based on the state signals of the aircraft in the current and previous cycles, including: Multiply the aircraft's current state signal, the difference between the current state signal and the previous state estimate, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix. Project the multiplication result onto the previous state disturbance estimate as the current state disturbance estimate. The state estimate signal is obtained by inputting the flight control command into the reference system state equation. Multiply the difference between the current state signal and the previous state estimation signal, the driver's current control command signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the input disturbance estimation term of the previous beat as the current beat input disturbance estimation term; Multiply the difference between the aircraft's current state signal and the previous state estimation signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous constant disturbance estimation term as the current constant disturbance estimation term.

2. The method according to claim 1, characterized in that According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained, including: Determining the sum of the product of the current beat input disturbance estimation term and the current beat control command signal of the pilot, the product of the current beat state disturbance estimation term and the current beat state signal of the aircraft, and the constant disturbance estimation term; The product of the preset reference matrix command gain and the driver's current control command signal is subtracted from the sum value to obtain an unfiltered compensation control law command; The unfiltered compensation control law instruction is passed through a low-pass filter to obtain the compensation control law instruction.

3. The method according to claim 1, characterized in that According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained, including: The result of multiplying the aircraft's current state signal by the full-state feedback gain matrix is ​​used as the full-state feedback signal; The result of multiplying the driver's current control command signal by the control gain matrix is ​​used as the main control law control signal; The full-state feedback signal is summed with the main control law manipulation signal to obtain the main control law instruction.

4. The method according to claim 1, wherein The status signal data includes: airspeed, acceleration, altitude, airflow angle, attitude angle, attitude angle rate, attitude angle change rate and three-axis overload.

5. A flight control device for a folding-wing deformable aircraft, characterized in that: A flight control method for a folding-wing variant aircraft according to any one of claims 1 to 4, wherein the device comprises: Identify modules for: Determine the initial dynamic configuration according to the initial static configuration of the folding wing variant aircraft, and estimate the configurational disturbance term in the initial dynamic configuration according to the state signals of the aircraft at the current and previous beats; According to the variable configuration disturbance term and the driver's current control command signal, the compensation control law command is obtained; According to the current state signal of the aircraft and the current control command signal of the pilot, the main control law command is obtained; The sending module is used to add the compensation control law instruction and the main control law instruction to obtain the flight control instruction, and send the flight control instruction to the servo.

6. The device according to claim 5, characterized in that Identify the module, specifically for: A reference system state equation is obtained according to the initial static configuration of the folding wing variant aircraft, wherein the reference system state equation includes a state vector, an input vector, a reference system state matrix and a reference system input matrix; Add three disturbance terms, namely state disturbance, input disturbance and constant disturbance, to the reference system state equation to construct the real system state equation of the current aircraft. The state disturbance term is multiplied by the state vector to obtain a first result, and the input disturbance term is multiplied by the input vector to obtain a second result. The first result, the second result and the constant disturbance term are summed and then added to the input vector. The real system state equation is used as the initial dynamic configuration.

7. The device according to claim 6, characterized in that Identify the module, specifically for: Multiply the aircraft's current state signal, the difference between the current state signal and the previous state estimate, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous state disturbance estimate as the current state disturbance estimate. Multiply the aircraft's current state signal, the pilot's current control command signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the input disturbance estimation term of the previous cycle as the current cycle input disturbance estimation term; Multiply the difference between the aircraft's current state signal and the previous state estimation signal, the solution of the Lyapunov equation of the reference system state matrix, and the reference system input matrix, and project the multiplication result onto the previous constant disturbance estimation term as the current constant disturbance estimation term.

8. The device according to claim 7, characterized in that Identify the module, specifically for: Determining the sum of the product of the current beat input disturbance estimation term and the current beat control command signal of the pilot, the product of the current beat state disturbance estimation term and the current beat state signal of the aircraft, and the constant disturbance estimation term; The product of the preset reference matrix command gain and the driver's current control command signal is subtracted from the sum value to obtain an unfiltered compensation control law command; The unfiltered compensation control law instruction is passed through a low-pass filter to obtain the compensation control law instruction.

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