Rotor wing waving suppression method and device for multi-propeller tilt-rotor aircraft

Abstract

The invention provides a rotor flapping suppression method and device for a multi-propeller tilt-rotor aircraft, and the method comprises the steps: in a transition flight mode, feeding back a rotor flapping state quantity to an elevator, adjusting the pitching moment characteristics of the whole aircraft, and tracking an operation input instruction through designing a self-adaptive sliding mode controller and a control distribution scheme, thereby achieving the suppression of the flapping of the rotor of the multi-propeller tilt-rotor aircraft. And forming a transition flight mode rotor flapping suppression control law. In a fixed-wing flight mode, the pitch angular velocity of an aircraft body and the instantaneous peak value of rotor waving are reduced through rate limitation and a longitudinal instruction, and an AFCS instruction model and a stability compensation mechanism are corrected, so that the sensitivity of control is reduced, the damping of a closed-loop system is increased, meanwhile, rotor waving is limited by feeding back the pitch angular velocity of the aircraft body, and the control accuracy is improved. And forming a rotor flapping suppression control law in the fixed-wing flight mode. The multi-propeller tilt-rotor aircraft can realize rotor flapping suppression in both a transition flight mode and a fixed-wing flight mode.

Description

Technical Field

[0001] This invention belongs to the field of helicopter flight dynamics and flight control technology, and particularly relates to a method and device for suppressing rotor flapping in a multi-rotor tiltrotor aircraft. Background Technology

[0002] The flight process of a tiltrotor aircraft is a typical variable-speed, variable-configuration process, during which the flight mechanical characteristics change drastically. The multi-rigid-body coupled motion during tilting adds additional hub loads to the rotor system and exhibits strong nonlinear characteristics, making it highly susceptible to flight safety accidents. To reduce rotor flapping loads and improve the flight safety of tiltrotor aircraft, the main technical means include active and passive protection measures. The primary goal is to keep the aircraft away from the design boundaries and fly within the design envelope.

[0003] In existing technologies, passive measures include embedding dynamic vibration absorbers inside the blades in the flapping direction to suppress the second-order flapping alternating load of rigid rotor blades; active measures include utilizing the lift offset characteristics of coaxial rigid rotors to suppress the overall flapping-shimmy bending moment of the rotor by adjusting the lift offset. These measures rely on additional component design (dynamic vibration absorbers) or are based on the unique properties of the configuration (lift offset). Due to their high dependence on specific environments such as installation or configuration and their lack of flexibility, their versatility and portability are poor.

[0004] For multi-rotor tiltrotor aircraft, the rotor flapping load characteristics are different in different configurations. In helicopter mode, rotor flapping can be used as a load reduction measure, while rotor flapping in transition flight mode and fixed-wing mode needs to be suppressed by load reduction measures. The rotor flapping suppression method based on flight control law has a universal and portable control architecture and can be used as an important means to ensure flight safety. Summary of the Invention

[0005] To address the problems of poor versatility and portability in existing technologies that rely on additional component designs (powered vibration absorbers) or are based on the unique properties of the configuration (lift offset), this invention provides a method and device for suppressing rotor flapping in multi-rotor tiltrotor aircraft. The technical solution is as follows: Firstly, a method for suppressing rotor flapping in a multi-rotor tiltrotor aircraft is provided. In the transitional flight mode, the rotor flapping state is fed back to the elevator to adjust the pitch moment characteristics of the entire aircraft. The control input commands are tracked by designing an adaptive sliding mode controller and control allocation scheme to form a rotor flapping suppression control law for the transitional flight mode.

[0006] Furthermore, the method also includes: In fixed-wing flight mode, the pitch rate and rotor flapping instantaneous peak values ​​are reduced by limiting the rate and using longitudinal commands. The AFCS command model and stability compensation mechanism are also modified to reduce the sensitivity of the control and increase the damping of the closed-loop system. At the same time, rotor flapping is limited by feedback of the pitch rate, forming a rotor flapping suppression control law for fixed-wing flight mode.

[0007] Optionally, in transitional flight mode, the rotor flapping suppression process specifically includes: Step 1a: Obtain the vertical instructions from PFCS and AFCS; Step 2a: Input the longitudinal command to the adaptive slicker distribution module, and redistribute the longitudinal command to the elevator and rotor control through the adaptive slicker distribution module; Step 3a: Use the rotor flapping state quantity in the aircraft dynamics as feedback, and obtain a feedback quantity within a suitable range after passing through the dead zone and gain circuit; Step 4a: The feedback quantity is further processed through filtering, rate limiting, and saturation circuits to obtain a control input command adapted to the control mechanism; Step 5a: Input the control input commands to the adaptive gliding membrane allocation module, and then distribute the control input commands and longitudinal commands to the elevator controller and rotor controller. The elevator controller and rotor controller output commands to the aircraft dynamics module to complete the flight dynamics response calculation, and then adjust the pitch moment characteristics of the entire aircraft.

[0008] Optionally, in fixed-wing flight mode, the rotor flapping suppression process specifically includes: Step 1b: Obtain the vertical instructions from PFCS and AFCS; Step 2b: Limit the rate of vertical commands to control the rate of change of manipulation commands; Step 3b: The longitudinal command enters the forward channel and the command model respectively to obtain the first longitudinal command and the second longitudinal command. The second longitudinal command is compared with the response obtained by the sensor to form an error signal. The error signal is corrected by the stability compensation mechanism to form a new second longitudinal command. Step 4b: The pitch rate of the airframe fed back by the sensor is processed by frequency limiting filter and amplitude limiting filter, and then input to the rotor controller along with the new second longitudinal command and the first longitudinal command. The new second longitudinal command and the first longitudinal command are input to the elevator controller. Step 5b: The commands output by the rotor controller and the elevator controller are entered into the aircraft dynamics module to complete the flight dynamics response calculation and transmit it to the sensors.

[0009] Optionally, the vertical instructions obtained from PFCS and AFCS are specifically as follows: Step 1c: Receive the control command. After the control command passes through the PFCS command mode, a first expected response is obtained. After the control command passes through the AFCS command mode, a second expected response is obtained. Step 2c: Compare the second desired response with the response fed back by the sensor to form an error signal; Step 3c: Perform stability compensation on the error signal, and after limiting it through the filter and AFCS terminal, combine it with the first expected response and the state information fed back by the sensor to form a total control command. The total control command includes a longitudinal command. Step 4c: Generate rotor, control surface, and full authority digital engine control commands based on the overall control commands; Step 5c: Each control command enters the aircraft dynamics module to complete the flight dynamics response calculation and is transmitted to the sensors.

[0010] The process of the adaptive sliding membrane allocation module allocating instructions is as follows: (1) Calculate the longitudinal motion attitude error vector For longitudinal motion control, the formula for calculating the longitudinal motion attitude error vector is:

[0011] in, Forward flight speed command; This is the pitch angle command; This is a vertical speed command; For command signals; (2) Design the representation of the synovial surface of the flight dynamics system model: , Design parameters Both are diagonal matrices, with t0 being the time point. (3) Design the control input expression: , (4) Design the boundary layer expression: , The thickness of the boundary layer, , The overall control input expression is: , The saturation function is defined as: , (5) Design an adaptive sliding diaphragm control scheme and complete command allocation. The sliding mode control law is:

[0012] Define a new variable based on the SAT function and boundary layer expression. It is the distance between the sliding variable and the boundary layer: , When the sliding variable remains outside the boundary layer When the sliding variable remains within the boundary layer , Design Estimation Parameters Real-time update expression: , Using estimated parameters The real-time update expression enables the system to slide on the expected sliding surface, completes instruction allocation, and ensures the system's tracking performance.

[0013] The control input expression is designed as follows:

[0014] In the formula, It is the continuous control part of a stable, ideal system. It is the part of discontinuous control used to compensate for disturbances and ensure that the system moves according to the designed control law, maintaining its trajectory. That is, get , Solution:

[0015] The discontinuous control section is designed as follows:

[0016] in, It is a diagonal matrix. Combining the continuous control section and the discontinuous control section, the overall control input expression is: .

[0017] Optionally, the correction process for stability compensation in step 3b is as follows: (1) Establish a stability compensation model When the actual sensor response does not match the expected response of the command model, the diagonal matrix becomes a non-identity matrix, resulting in the equation: , I represents the identity matrix. make The stability compensation model is as follows: , use replace ,get: , (2) Establish a modified parameter control allocation model The corrected parameter control allocation model expression is as follows:

[0018] in, It is the actual control input; This is the required steady-state control input; B is the virtual control command generated from the high-level controller; B is the control performance matrix. For weighted matrices, Describing the Euclidean 2-norm, (3) The stability compensation model adaptively changes the parameters. This generates more virtual control signals, which are then fed into the correction parameter control allocation model to compensate for the adverse effects of uncertain actuator operating performance and maintain the original tracking performance of the system.

[0019] In a second aspect, a rotor flapping suppression device for a multi-rotor tiltrotor aircraft is provided, which performs any of the methods described in the first aspect, the device comprising: The first processing module is used to adjust the pitch moment characteristics of the entire aircraft by feeding back the rotor flapping state to the elevator in the transition flight mode, and to track the control input commands by designing an adaptive sliding mode controller and control allocation scheme to form a rotor flapping suppression control law in the transition flight mode.

[0020] Furthermore, the device also includes: The second processing module is used to reduce the airframe pitch rate and rotor flapping instantaneous peak value in fixed-wing flight mode by limiting the rate and longitudinal commands, and to modify the AFCS command model and stability compensation mechanism, thereby reducing the sensitivity of control and increasing the damping of the closed-loop system. At the same time, it limits rotor flapping by feeding back the airframe pitch rate, forming a rotor flapping suppression control law for fixed-wing flight mode.

[0021] The beneficial effects of this invention are at least as follows: Compared to the state without rotor flapping suppression, after applying the corresponding control law, the multi-rotor tiltrotor aircraft can achieve rotor flapping suppression in both transitional flight mode and fixed-wing flight mode. Attached Figure Description

[0022] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. The drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0023] Figure 1 A block diagram of the control law structure for a multi-rotor tiltrotor aircraft; Figure 2 Block diagram of the control law for suppressing rotor flapping in transition mode; Figure 3 Block diagram of the control law for suppressing rotor flapping in fixed-wing mode; Figure 4 The figure shows the variation of rotor flapping angle under different nacelle tilt angles (without rotor flapping suppression).

[0024] Figure 5 A comparison chart showing the changes in rotor flapping angle under different nacelle tilt angles (with rotor flapping suppression).

[0025] Figure 6 This is a comparison chart of longitudinal control response in fixed-wing flight mode. Detailed Implementation

[0026] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0027] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0028] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited from each other.

[0029] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments.

[0030] The control law structure of a multi-rotor tiltrotor aircraft is divided into a Basic Flight Control System (PFCS) and an Automatic Flight Control System (AFCS). The PFCS meets the basic control requirements of the aircraft, while the AFCS enhances performance in complex missions by increasing stability and ensuring a flight quality level of 1. This invention proposes a rotor flapping suppression method applicable to different flight modes of multi-rotor tiltrotor aircraft based on the PFCS and AFCS.

[0031] See Figure 1 The control law is calculated using a model-tracking structure by comparing expected and measured values. Error signals are compensated and combined with PFCS commands, then the overall control commands are transmitted to the cabin attitude and airspeed data to form rotor, control surface, and full-authority digital engine control commands. The control law structure includes PFCS filters to counteract the dynamic behavior of the object, filters from the AFCS command model to generate the desired response, and stabilization filters to generate feedback compensation. All filters included in the control law are structured filters used to enhance flight stability.

[0032] The specific steps of the method in this embodiment of the invention are as follows: Step 11: Receive the control command. After the control command passes through the PFCS command mode, the first expected response is obtained. After the control command passes through the AFCS command mode, the second expected response is obtained. Step 12: Compare the second desired response with the response obtained from the sensor to form an error signal; Step 13: Perform stability compensation on the error signal, and after limiting it through the filter and AFCS terminal, combine it with the first expected response and the state information fed back by the sensor to form a total control command. The total control command includes the longitudinal command. Step 14: Generate rotor, control surface, and full authority digital engine control commands based on the overall control commands; Step 15: Each control command enters the aircraft dynamics module to complete the flight dynamics response calculation and is transmitted to the sensors.

[0033] In transitional flight mode, based on the aforementioned control architecture, a rotor flapping suppression control law is superimposed on the longitudinal control commands. An adaptive sliding mode controller and control allocation technology are designed to track optimal control. The flapping limitation control law feeds back the rotor flapping state to the elevator, adjusting the overall aircraft pitch moment characteristics to reduce rotor flapping and rotor hub moment.

[0034] Control law block diagram as follows Figure 2 As shown, the specific steps are as follows: Step 21: Use the longitudinal command obtained in Step 13 based on PFCS and AFCS as the control law input; Step 22: Input the longitudinal command to the adaptive slicker distribution module, and redistribute the longitudinal command to the elevator and rotor control through the adaptive slicker distribution module; Step 23: Use the rotor flapping state quantity in the aircraft dynamics as feedback, and obtain a feedback quantity within a suitable range after passing through the dead zone and gain circuit; Step 24: The feedback quantity is further processed through filtering, rate limiting, and saturation circuits to obtain a control input command adapted to the control mechanism; Step 25: Input the control input commands to the adaptive gliding membrane allocation module, and then further allocate the control input commands and longitudinal commands to the elevator controller and rotor controller. The elevator controller and rotor controller output commands to the aircraft dynamics module to complete the flight dynamics response calculation, and then adjust the pitch moment characteristics of the entire aircraft.

[0035] The process of the adaptive sliding membrane allocation module allocating instructions is as follows: (1) Establish the longitudinal motion attitude error vector For longitudinal motion control, the corresponding attitude error vector is established as follows: The process of the adaptive sliding membrane allocation module allocating instructions is as follows: (1) Calculate the longitudinal motion attitude error vector For longitudinal motion control, the formula for calculating the longitudinal motion attitude error vector is:

[0036] Forward flight speed command; This is the pitch angle command; This is a vertical speed command; For command signals; (2) Establishing the slippery surface of the flight dynamics system model: Considering the tracking error vector in the above equation, the sliding surface of the system model is constructed as follows:

[0037] in,

[0038] The first part is a linear combination of system states similar to a traditional sliding mode design, and the second part is used to introduce the integral term, let:

[0039] This includes the design parameters of the integral sliding surface. and polynomial It must satisfy the conditions of the Hurwitz polynomial, where p is a Laplace operator. Therefore, The eigenvalues ​​should have one negative real root, for numbers containing positive numbers. equation ,Right now Design parameters and Each satisfies and .

[0040] In the integral sliding mode method, because the order of the equations of motion is equal to the order of the original system, the system's robustness is maintained throughout the entire response from the initial time. The system trajectory under the integral sliding mode method starts from the designed sliding surface at the initial moment, without a stage of reaching the sliding surface, effectively avoiding control chattering. The sliding surface is designed as follows: , Among them, design parameters Both are diagonal matrices, with t0 being the time point. (3) Design the control input expression: After defining the sliding surface, it is necessary to keep the sliding variables on the sliding plane. Therefore, given... When the actuator does not malfunction, The state symbolizing the ideal system needs to be in Ensure The control input is designed as follows:

[0041] It is the continuous control part of a stable, ideal system. It is the part of discontinuous control used to compensate for disturbances and ensure that the system moves according to the designed control law, maintaining its trajectory. That is, get ,

[0042] The solution is:

[0043] The discontinuous control section is designed as follows: , Using a diagonal matrix makes the sliding surface more attractive.

[0044] Combining the continuous control section and the discontinuous control section, the overall control input expression is: .

[0045] (4) Design the boundary layer expression: To mitigate the effects of model uncertainties and actuator malfunctions, discontinuous control inputs may exhibit control chatter due to overuse. This chatter is smoothed by adding a thin boundary layer, as shown below: , in, The thickness of the boundary layer, , Therefore, the overall control input expression is: , The saturation function is defined as follows: , When the discontinuous control gain is At this time, the system can achieve the expected sliding motion and keep it within the boundary layer.

[0046] (5) Design an adaptive sliding diaphragm control scheme and complete command allocation. Introducing adaptive control and combining it with sliding mode control increases the robustness of the control system, ensuring that the controller still has good tracking performance during configuration changes.

[0047] The sliding mode control law is:

[0048] Define a new variable based on the SAT function and boundary layer expression. It is the distance between the sliding variable and the boundary layer: , When the sliding variable remains outside the boundary layer When the sliding variable remains within the boundary layer , Design Estimation Parameters Real-time update expression: .

[0049] Using the above-mentioned adaptive sliding mode control method, when At that time, the system achieves sliding motion on the expected sliding surface, completes control allocation, and ensures the tracking performance of the system.

[0050] In fixed-wing flight mode, based on the above control architecture, three special modifications to the flight control system are considered to suppress rotor flapping, such as... Figure 3As shown. First, the instantaneous peak values ​​of airframe pitch rate and rotor flapping are reduced through rate limiting and longitudinal commands. Second, the AFCS command model and stability compensation mechanism are modified to reduce control sensitivity and increase the damping of the closed-loop system. Third, rotor flapping is directly limited by feedback of airframe pitch rate.

[0051] The specific steps are as follows: Step 31: Use the longitudinal command obtained in step 13 based on PFCS and AFCS as the control law input; Step 32: Limit the rate of vertical commands to control the rate of change of manipulation commands; Step 33: The longitudinal command enters the forward channel and the command model respectively to obtain the first longitudinal command and the second longitudinal command. The second longitudinal command is compared with the response obtained by the sensor to form an error signal. The error signal is corrected by the stability compensation mechanism to form a new second longitudinal command. Step 34: The pitch rate of the aircraft fed back by the sensor is processed by frequency limiting filter and amplitude limiting filter, and then input to the rotor controller along with the new second longitudinal command and the first longitudinal command. The new second longitudinal command and the first longitudinal command are input to the elevator controller. Step 35: The commands output by the rotor controller and the elevator controller are entered into the aircraft dynamics module to complete the flight dynamics response calculation and transmit it to the sensors.

[0052] The correction process for stability compensation is as follows: (1) Establish a modified parameter control allocation model The control allocation expression is set as follows:

[0053] in, These are system state variables and output variables; These are the system's state matrix and control performance matrix; For manipulation quantity, These are the upper limit of the deflection of the i-th control surface, respectively.

[0054] Assumption Decompose the control effectiveness matrix into By introducing pseudo-control commands The original system will be transformed into the following form:

[0055] Configure control assignments This makes the output torque as close as possible to the desired torque v. d Set the control allocation problem as follows:

[0056] Without considering other constraints or desired performance indicators, we seek a control variable u that minimizes the p-norm of the tracking error. The dynamic control allocation method based on sequential quadratic programming describes it as follows:

[0057]

[0058] It is the actual control input; This is the required steady-state control input; B is the virtual control command generated from the high-level controller; B is the control performance matrix. For weighted matrices, This represents the Euclidean 2-norm.

[0059] (2) Establish a stability compensation model When the actual sensor response does not match the expected response of the command model, the diagonal matrix becomes a non-identity matrix, resulting in: I is the identity matrix. make The stability compensation model is as follows: , At this point, because the control allocation module lacks accurate uncertainty information, a virtual control error exists. This will reduce the overall tracking performance of the system and may even cause the system to lose stability. In this situation, to maintain the tracking performance of the closed-loop system, the advanced controller needs to be reconfigured. From the above equation, we can obtain the required virtual control signal v when an error exists. d To maintain high-level control tracking performance, parameter H needs to be adaptively adjusted to eliminate virtual control errors between the actual virtual control signal v and the actual virtual control signal v. In this case, using Replacing Hv, we get: , At this point, the actual virtual control signal will not be equal to the expected virtual control signal generated by the controller. In this situation, the parameters are adaptively changed. The controller will generate more virtual control signals to compensate for the adverse effects of uncertain actuator operating performance and maintain the original tracking performance of the system. For example, transitional flight mode: Figure 4 and Figure 5 This figure shows the variation of the flapping angle with flight speed during transitional flight modes under different flight control laws. As can be seen from the figure, the overall rotor flapping angle decreases significantly when the rotor flapping suppression law is increased. Fixed-wing flight mode: Figure 6 This describes the helicopter response history under the same longitudinal stick control in fixed-wing flight modes with different flight control laws, at a flight speed of V. D Rapid longitudinal dipole maneuvering is performed. As can be seen from the figure, when the rotor flapping suppression control law is increased, the change in helicopter pitch speed is small and the overall overload is reduced, but the sum of the helicopter pitch rate and flapping rate is reduced by 30% to 50%, which can effectively reduce the rotor hub load.

[0060] In summary, compared with the state without rotor flapping suppression, after applying the corresponding control law, the multi-rotor tiltrotor aircraft can achieve rotor flapping suppression in both transitional flight mode and fixed-wing flight mode, which shows that the flapping suppression method proposed in this invention is truly effective.

[0061] This invention provides a rotor flapping suppression device for a multi-rotor tiltrotor aircraft, which executes the method described in this invention. The device includes: The first processing module is used to adjust the pitch moment characteristics of the entire aircraft by feeding back the rotor flapping state to the elevator in the transition flight mode, and to track the control input commands by designing an adaptive sliding mode controller and control allocation scheme to form a rotor flapping suppression control law in the transition flight mode.

[0062] The device further includes: The second processing module is used to reduce the airframe pitch rate and rotor flapping instantaneous peak value in fixed-wing flight mode by limiting the rate and longitudinal commands, and to modify the AFCS command model and stability compensation mechanism, thereby reducing the sensitivity of control and increasing the damping of the closed-loop system. At the same time, it limits rotor flapping by feeding back the airframe pitch rate, forming a rotor flapping suppression control law for fixed-wing flight mode.

[0063] The above description merely illustrates embodiments of the present invention and is quite specific and detailed; however, it should not be construed as limiting the scope of the patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of the present invention, and these all fall within the scope of protection of the present invention. Furthermore, any parts of the present invention not described in detail are conventional techniques.

Claims

1. A method for suppressing rotor flapping in a multi-rotor tiltrotor aircraft, characterized in that, In transitional flight mode, the pitch moment characteristics of the entire aircraft are adjusted by feeding back the rotor flapping state to the elevator, and the control input commands are tracked by designing an adaptive sliding mode controller and control allocation scheme to form a rotor flapping suppression control law for transitional flight mode.

2. The method according to claim 1, characterized in that, The method further includes: In fixed-wing flight mode, the pitch rate and rotor flapping instantaneous peak values ​​are reduced by limiting the rate and using longitudinal commands. The AFCS command model and stability compensation mechanism are also modified to reduce the sensitivity of the control and increase the damping of the closed-loop system. At the same time, rotor flapping is limited by feedback of the pitch rate, forming a rotor flapping suppression control law for fixed-wing flight mode.

3. The method according to claim 1, characterized in that, In the transitional flight mode, the rotor flapping suppression process specifically includes: Step 1a: Obtain the vertical instructions from PFCS and AFCS; Step 2a: Input the longitudinal command to the adaptive slicker distribution module, and redistribute the longitudinal command to the elevator and rotor control through the adaptive slicker distribution module; Step 3a: Use the rotor flapping state quantity in the aircraft dynamics as feedback, and obtain a feedback quantity within a suitable range after passing through the dead zone and gain circuit; Step 4a: The feedback quantity is further processed through filtering, rate limiting, and saturation circuits to obtain a control input command adapted to the control mechanism; Step 5a: Input the control input commands to the adaptive gliding membrane allocation module, and then distribute the control input commands and longitudinal commands to the elevator controller and rotor controller. The elevator controller and rotor controller output commands to the aircraft dynamics module to complete the flight dynamics response calculation, and then adjust the pitch moment characteristics of the entire aircraft.

4. The method according to claim 2, characterized in that, In fixed-wing flight mode, the rotor flapping suppression process specifically includes: Step 1b: Obtain the vertical instructions from PFCS and AFCS; Step 2b: Limit the rate of vertical commands to control the rate of change of manipulation commands; Step 3b: The longitudinal command enters the forward channel and the command model respectively to obtain the first longitudinal command and the second longitudinal command. The second longitudinal command is compared with the response obtained by the sensor to form an error signal. The error signal is corrected by the stability compensation mechanism to form a new second longitudinal command. Step 4b: The pitch rate of the airframe fed back by the sensor is processed by frequency limiting filter and amplitude limiting filter, and then input to the rotor controller along with the new second longitudinal command and the first longitudinal command. The new second longitudinal command and the first longitudinal command are input to the elevator controller. Step 5b: The commands output by the rotor controller and the elevator controller are entered into the aircraft dynamics module to complete the flight dynamics response calculation and transmit it to the sensors.

5. The method according to claim 3 or 4, characterized in that, The specific steps to obtain the vertical instructions from PFCS and AFCS are as follows: Step 1c: Receive the control command. After the control command passes through the PFCS command mode, a first expected response is obtained. After the control command passes through the AFCS command mode, a second expected response is obtained. Step 2c: Compare the second desired response with the response fed back by the sensor to form an error signal; Step 3c: Perform stability compensation on the error signal, and after limiting it through the filter and AFCS terminal, combine it with the first expected response and the state information fed back by the sensor to form a total control command. The total control command includes a longitudinal command. Step 4c: Generate rotor, control surface, and full authority digital engine control commands based on the overall control commands; Step 5c: Each control command enters the aircraft dynamics module to complete the flight dynamics response calculation and is transmitted to the sensors.

6. The method according to claim 3, characterized in that, The process of the adaptive sliding membrane allocation module allocating instructions is as follows: (1) Calculate the longitudinal motion attitude error vector For longitudinal motion control, the formula for calculating the longitudinal motion attitude error vector is: Forward speed command; This is the pitch angle command; This is a vertical speed command; For command signals; (2) Design the representation of the synovial surface of the flight dynamics system model: ， Design parameters Both are diagonal matrices, with t0 being the time point. (3) Design the control input expression: ， (4) Design the boundary layer expression: ， The thickness of the boundary layer, , The overall control input expression is: ， The saturation function is defined as: ， (5) Design an adaptive sliding mode control scheme, complete the command allocation, and the sliding mode control law is: Define a new variable based on the SAT function and boundary layer expression. It is the distance between the sliding variable and the boundary layer: ， When the sliding variable remains outside the boundary layer When the sliding variable remains within the boundary layer , Design Estimation Parameters Real-time update expression: 。 7. The method according to claim 6, characterized in that, The control input expression is designed as follows: u0 is the continuous control part of a stable, ideal system, while u1 is the discontinuous control part used to compensate for disturbances and ensure that the system moves according to the designed control law, maintaining its position on the system trajectory. That is, we get u0. Solution: The discontinuous control section is designed as follows: , It is a diagonal matrix. Combining the continuous control section and the discontinuous control section, the overall control input expression is: 。 8. The method according to claim 4, characterized in that, The correction process for stability compensation in step 3b is as follows: (1) Establish a stability compensation model When the actual sensor response does not match the expected response of the command model, the diagonal matrix becomes a non-identity matrix, resulting in: I is the identity matrix. make The stability compensation model is as follows: ， use Replacing Hv, we get: ， (2) Establish a modified parameter control allocation model The corrected parameter control allocation model expression is as follows: It is the actual control input; This is the required steady-state control input; These are virtual control commands generated from the high-level controller; B is the control effectiveness matrix. For weighted matrices, Describing the Euclidean 2-norm, (3) The stability compensation model adaptively changes the parameters. This generates more virtual control signals, which are then fed into the correction parameter control allocation model to compensate for the adverse effects of uncertain actuator operating performance and maintain the original tracking performance of the system.

9. A rotor flapping suppression device for a multi-rotor tiltrotor aircraft, characterized in that, The apparatus for performing the method according to any one of claims 1 to 8 comprises: The first processing module is used to adjust the pitch moment characteristics of the entire aircraft by feeding back the rotor flapping state to the elevator in the transition flight mode, and to track the control input commands by designing an adaptive sliding mode controller and control allocation scheme to form a rotor flapping suppression control law in the transition flight mode.

10. The apparatus according to claim 9, characterized in that, The device further includes: The second processing module is used to reduce the airframe pitch rate and rotor flapping instantaneous peak value in fixed-wing flight mode by limiting the rate and longitudinal commands, and to modify the AFCS command model and stability compensation mechanism, thereby reducing the sensitivity of control and increasing the damping of the closed-loop system. At the same time, it limits rotor flapping by feeding back the airframe pitch rate, forming a rotor flapping suppression control law for fixed-wing flight mode.

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