Aircraft attitude control method and device, equipment and storage medium

Abstract

The invention relates to the technical field of aircrafts, and discloses an aircraft attitude control method, device and equipment and a storage medium, and the method comprises the steps: collecting flight state parameters of an aircraft; the pitch angle data in the flight state parameters are compared with a preset angle threshold value, corresponding flight mode switching is executed, and the attitude deviation value is calculated; power distribution calculation is conducted on the three ducts, and the target rotating speed of each duct paddle is obtained; performing position optimization calculation on the guide plates below the duct to obtain the position control quantity of each guide plate; the three flow guide plates are configured as triangular control points, the position compensation amount of an aileron, a yaw rudder and an elevator is calculated and determined through vector synthesis, and the angle control amount of each control surface is obtained; according to the method, the attitude of the aircraft is converted into an execution signal, execution feedback data is acquired to correct a control instruction in real time, closed-loop attitude control of the aircraft is realized, the attitude control precision of the aircraft is improved, and the stability of the flight process is ensured.

Description

Technical Field

[0001] The present application relates to the field of aircraft control technology, and in particular to an aircraft attitude control method, device, equipment and storage medium. Background Art

[0002] Traditional vertical take-off and landing aircraft mainly use tilt-rotor or composite wing layouts, but these solutions have complex mechanical structures and low reliability. Tilt-rotor requires complex tilting mechanisms to achieve the conversion between vertical take-off and landing and horizontal flight, which increases the weight and failure risk of the system; while the composite wing layout requires an additional lift system, resulting in redundancy of the power system and reduced energy efficiency.

[0003] Existing vertical take-off and landing aircraft also face severe challenges in attitude control. During the transition between vertical take-off and landing and horizontal flight modes, the aircraft needs to coordinate multiple actuators at the same time, including the power system, rudders, and attitude control devices. The control system is highly complex and has strict requirements on control accuracy. In addition, when a partial power system failure occurs in the aircraft, the existing control scheme often cannot guarantee the attitude stability of the aircraft, which poses a major safety hazard. Summary of the invention

[0004] The present application provides an attitude control method, device, equipment and storage medium for an aircraft, thereby improving the accuracy of attitude control of the aircraft and ensuring the stability of the flight process.

[0005] In a first aspect, the present application provides a method for controlling the attitude of an aircraft, the method comprising:

[0006] Collect the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight status parameters;

[0007] Compare the pitch angle data in the flight state parameter with a preset angle threshold to determine whether a flight mode switching condition is met; when the switching condition is met and a mode switching instruction issued by an autopilot module or an external control module is received, execute the corresponding flight mode switching, and calculate the difference between the current attitude and the target attitude to obtain an attitude deviation;

[0008] Based on the current flight mode and the attitude deviation, power distribution calculation is performed on the three ducts respectively to obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts;

[0009] According to the current flight mode and the attitude deviation, the position of the guide plate below the duct is optimized and calculated to obtain the position control amount of each guide plate;

[0010] Based on the current flight mode, the three deflectors are configured as triangular control points, and the position compensation of the aileron, yaw rudder and elevator is determined by vector synthesis calculation to obtain the angle control value of each rudder surface;

[0011] The target rotation speed of each ducted blade, the position control amount of each guide plate and the angle control amount of each control surface are converted into execution signals, and execution feedback data is obtained to perform real-time correction of control instructions to achieve closed-loop attitude control of the aircraft.

[0012] A second aspect of the present application provides an attitude control device for an aircraft, the attitude control device for an aircraft comprising:

[0013] The acquisition module is used to collect the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight status parameters;

[0014] A difference calculation module is used to compare the pitch angle data in the flight state parameters with a preset angle threshold to determine whether a flight mode switching condition is met. When the switching condition is met and a mode switching instruction issued by the autopilot module or the external control module is received, the corresponding flight mode switching is executed, and a difference calculation is performed between the current attitude and the target attitude to obtain an attitude deviation;

[0015] A power distribution calculation module is used to perform power distribution calculations on the three ducts respectively based on the current flight mode and the attitude deviation, and obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts;

[0016] A position optimization calculation module is used to perform position optimization calculation on the guide plates below the duct according to the current flight mode and the attitude deviation, so as to obtain the position control amount of each guide plate;

[0017] A vector synthesis calculation module is used to configure the three deflectors as triangular control points based on the current flight mode, determine the position compensation of the aileron, yaw rudder and elevator through vector synthesis calculation, and obtain the angle control amount of each rudder surface;

[0018] The execution module is used to convert the target rotation speed of each ducted blade, the position control value of each guide plate and the angle control value of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control command to realize the closed-loop attitude control of the aircraft.

[0019] The third aspect of the present application provides an electronic device, comprising: a memory and at least one processor, wherein the memory stores instructions; the at least one processor calls the instructions in the memory so that the electronic device executes the above-mentioned aircraft attitude control method.

[0020] A fourth aspect of the present application provides a computer-readable storage medium, wherein the computer-readable storage medium stores instructions, which, when executed on a computer, enable the computer to execute the above-mentioned aircraft attitude control method.

[0021] Compared with the prior art, the present application has the following beneficial effects: the design scheme of the combination of three duct layout and guide plate eliminates the tilting mechanism, simplifies the mechanical structure, improves the reliability of the system, and realizes more efficient power output through the coordinated control of the three ducts. By adopting the innovative control strategy of configuring the three guide plates as triangular control points, the precise control of the horizontal attitude of the aircraft is realized, and the complexity of the control system is reduced. A fault-tolerant control scheme based on power redundancy is designed. When a single duct fails, the power demand can be automatically allocated to the remaining normally working ducts to ensure the continuous controllability of the aircraft. A complete closed-loop control system is established, and the control command correction is performed through real-time feedback data, which improves the accuracy of attitude control and the robustness of the system. Smooth switching between vertical take-off and landing and horizontal flight modes is realized, and the stability of the flight process is ensured by real-time monitoring of the pitch angle and mode judgment. The multi-sensor data fusion technology and the extended Kalman filter algorithm are adopted to improve the accuracy of state estimation and provide reliable data support for attitude control. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative labor.

[0023] The structures, proportions, sizes, etc. illustrated in the drawings of this specification are only used to match the contents disclosed in the specification so as to facilitate understanding and reading by persons familiar with this technology. They are not used to limit the conditions under which the present invention can be implemented, and therefore have no substantive technical significance. Any structural modification, change in proportion or adjustment of size, without affecting the effects and purposes that can be achieved by the present invention, should still fall within the scope of the technical contents disclosed by the present invention.

[0024] Figure 1 is a schematic flow chart of a method for controlling an attitude of an aircraft provided by an embodiment of the present invention;

[0025] Figure 2 is a schematic block diagram of the structure of an attitude control device for an aircraft provided by an embodiment of the present invention;

[0026] Figure 3 It is a schematic block diagram of the structure of an electronic device provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0027] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0028] The flowcharts shown in the accompanying drawings are only examples and do not necessarily include all the contents and operations / steps, nor must they be executed in the order described. For example, some operations / steps may also be decomposed, combined or partially merged, so the actual execution order may change according to actual conditions.

[0029] It should also be understood that the terms used in this application specification are only for the purpose of describing specific embodiments and are not intended to limit the application. As used in this application specification and the appended claims, unless the context clearly indicates otherwise, the singular forms "a", "an" and "the" are intended to include plural forms.

[0030] It should be further understood that the term "and / or" used in the specification and appended claims of this application refers to any combination and all possible combinations of one or more of the associated listed items, and includes these combinations. Figure 1 , an embodiment of the attitude control method of the aircraft in the embodiment of the present application includes:

[0031] Step 100, collecting attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight state parameters;

[0032] It is understandable that the execution subject of the present application may be an attitude control device of an aircraft, or a terminal or a server, which is not limited here. The present application embodiment is described by taking a server as the execution subject as an example.

[0033] Specifically, the original signal output by the inertial measurement unit in the aircraft is low-pass filtered to eliminate the influence of high-frequency noise on the attitude angle data, and the attitude angle data of the pitch angle, roll angle and yaw angle of the aircraft relative to the ground are obtained. The acceleration sensor is used to collect data in the three directions of the x-axis, y-axis and z-axis. The output signal of the acceleration sensor is processed by the Kalman filter algorithm to effectively filter out the noise and obtain accurate acceleration data in the three orthogonal axes. The Kalman filter algorithm can adjust the weight of the filtering process in real time through the combination of dynamic model and observation model, so that the processed data is more consistent with the actual acceleration of the aircraft. At the same time, the zero bias compensation calculation is performed on the original angular velocity signal output by the gyroscope to eliminate the systematic error caused by sensor offset, and the angular velocity data is obtained to describe the rotation state of the aircraft around the three orthogonal axes. The output signal of the barometric altimeter is fused with the GPS altitude signal to obtain the altitude data. The dynamic pressure signal output by the pitot tube and the GPS ground speed signal are speed resolved. The pitot tube provides speed information relative to the air, while the GPS module provides speed information of the aircraft relative to the ground. Accurate speed data is obtained by jointly calculating these two signals. The output signals of the Hall sensors of the three ducts are sampled and processed with median filtering to remove occasional spike noise, so that the rotation speed data of the duct is smoother and more stable, and the duct rotation speed data is obtained. The output signals of the deflector position sensor, the aileron position sensor, the yaw rudder position sensor and the elevator position sensor are sampled and filtered to obtain the real-time position data of the rudder surface, and effectively remove the noise in the measurement to ensure the accuracy and stability of the rudder surface control signal. Based on attitude angle data, acceleration data, angular velocity data, altitude data and speed data, and duct rotation speed data, the present invention can also collect rudder surface position data for processing. The extended Kalman filter algorithm is used for state estimation calculation to obtain flight state parameters. The extended Kalman filter is a filtering method suitable for nonlinear systems, which can estimate and correct the state of the system in real time. The extended Kalman filter algorithm combines the aircraft's motion model and observation data to calculate flight state parameters such as pitch angle parameters, current position parameters, current speed parameters, and current acceleration parameters.

[0034] Step 200: Compare the pitch angle data in the flight state parameters with the preset angle threshold to determine whether the flight mode switching condition is met. When the switching condition is met and a mode switching instruction is received from the autopilot module or the external control module, the corresponding flight mode switching is executed, and the difference between the current attitude and the target attitude is calculated to obtain the attitude deviation;

[0035] Specifically, the pitch angle data in the flight state parameter is compared with the preset angle threshold to determine whether the flight mode switching condition is met. When the switching condition is met and the mode switching instruction issued by the autopilot module or the external control module is received, the corresponding flight mode switching is executed. The determination data of the vertical take-off and landing mode and the determination data of the horizontal flight mode are processed by logical operations to obtain the current flight mode. The logical operation includes a comprehensive analysis of the determination results of each mode, such as through logical operations such as AND or NOT, to ensure that the aircraft is in a clear flight mode without misjudgment due to interference between different modes. The control requirements and control targets are different in different flight modes. For example, in the vertical take-off and landing mode, the aircraft requires higher thrust and stability, while in the horizontal flight mode, it pays more attention to the direction control and speed stability of the flight. Based on the current flight mode, the coordinate difference between the current position parameter in the flight state parameter and the target position in three orthogonal axes is calculated to obtain position deviation data, reflecting the displacement of the aircraft in space. Based on the current flight mode, the speed difference between the current speed parameter in the flight state parameter and the target speed in three orthogonal axes is calculated to obtain speed deviation data, which provides the control system with speed deviation information of the aircraft during the movement, so that the control system can adjust the thrust size and direction to achieve the target speed. At the same time, based on the current flight mode, the acceleration difference between the current acceleration parameter in the flight state parameter and the target acceleration in three orthogonal axes is calculated to obtain acceleration deviation data. Acceleration deviation data is used to describe the difference between the acceleration change of the aircraft and the target acceleration, especially in the stages of take-off, acceleration, deceleration and landing. The acceleration deviation can directly reflect the force condition and motion state of the aircraft, providing a basis for the adjustment of the power system. The position deviation data, speed deviation data and acceleration deviation data are combined to obtain the motion state deviation matrix in the three-axis direction, which comprehensively reflects the overall difference between the current motion state of the aircraft and the target state. The motion state deviation matrix is ​​input into the dynamic feedback controller for state space solution to obtain the attitude deviation. The dynamic feedback controller calculates the attitude adjustment required by the aircraft by analyzing the deviation matrix and combining it with the dynamic model of the system, thus realizing the closed-loop feedback of attitude control. Through the state space solution method, the control system can comprehensively consider the difference between the motion state and the target state of the aircraft, and on this basis formulate a reasonable attitude adjustment strategy to gradually approach the target attitude of the aircraft and realize stable flight control. The calculation results of the attitude deviation are used to adjust the power distribution of the ducted blades, the position control of the guide plate and the angle compensation of the rudder surface, so as to ensure that the aircraft can maintain a stable attitude in different flight modes such as vertical take-off and landing, cruising and horizontal flight, thereby improving the safety and accuracy of the flight.

[0036] Step 300: Based on the current flight mode and attitude deviation, power distribution calculation is performed on the three ducts respectively to obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts.

[0037] It should be noted that the current flight mode and attitude deviation are subjected to state analysis calculation to obtain the overall power demand in the Z-axis direction and the attitude torque demand in the three orthogonal axes. The overall power demand in the Z-axis direction reflects the thrust required for the aircraft to maintain hovering or stable flight in the vertical direction, while the attitude torque demand describes the torque demand of the aircraft in the three axes, thereby ensuring precise attitude control. According to the overall power demand in the Z-axis direction, the three ducts are evenly distributed with power to obtain the reference power distribution coefficient of each duct, ensuring that under normal circumstances, the three ducts of the aircraft can work with balanced power output to maintain the stability and attitude control effect of the aircraft. In the process of even power distribution, the mass distribution and structural characteristics of the aircraft are considered to ensure that the power distribution of each duct is reasonable and effective. The real-time speed data of the duct is compared with the rated speed, and a fault detection operation is performed through the comparison to obtain the working state identification information of the duct. Through the duct working state identification information, the working state of each duct is monitored in real time, and faults are judged. During the flight, if any duct is detected to be faulty, the remaining ducts need to be compensated for power to ensure that the attitude control of the aircraft will not be significantly affected. Based on the duct working status identification information, when a duct is detected to be abnormal, the power demand of the abnormal duct is proportionally distributed to calculate the compensation power coefficients of the remaining normal working ducts. The compensation operation can ensure that the aircraft can still achieve overall attitude stability through the power adjustment of the remaining ducts when a duct fails, thereby improving the fault tolerance and reliability of the system. Based on the reference power distribution coefficient and the compensation power coefficient, the power output of the three ducts is dynamically adjusted and calculated to obtain the target power output of each duct, ensuring that each duct can flexibly adjust the power output according to the current flight state and fault conditions to maintain the overall stability and flight control effect of the aircraft. The target power output is mapped to the duct blade characteristics to obtain the corresponding duct speed control. The power demand is converted into a specific duct blade speed control. This process needs to consider multiple factors such as the aerodynamic characteristics, drag characteristics, and interaction between the duct and the air of the duct blade to ensure that the speed control can effectively reflect the required power output. Based on the duct speed control amount, the torque of each duct motor is calculated by PID control to obtain the control compensation amount of the motor. PID control can adjust the output torque of the motor in real time according to the speed error, ensuring that the duct can quickly respond to changes in the target speed, while reducing steady-state errors and overshoots, and achieving precise control of the duct motor. After completing the PID control calculation, the duct speed control amount and the motor control compensation amount are superimposed to obtain the target speed of each duct blade. Comprehensively considering the reference speed requirements and control compensation, the actual output of the duct motor is more in line with the current attitude control requirements of the aircraft.

[0038] Step 400: Optimize the position of the guide plates below the duct according to the current flight mode and attitude deviation, and obtain the position control value of each guide plate;

[0039] Specifically, according to the current flight mode, the attitude deviation is decomposed and calculated in the three axes of U, V, and W to obtain the expected correction force in the three directions. The overall attitude adjustment requirements are refined into specific axial correction forces to ensure that the aircraft can obtain reasonable correction effects in all directions. The expected correction force is calculated for the duct airflow guidance to obtain the flow angle data of the duct down-swirling airflow. The data of the down-swirling airflow direction angle is directly related to the guidance of the duct airflow, and the main function of the guide plate is to achieve the attitude adjustment of the aircraft by changing the direction of the down-swirling airflow. Based on the current flight mode, the action area of ​​the guide plate below the duct and the down-swirling airflow is dynamically calculated to obtain the airflow action coefficient. The action area between the guide plate and the down-swirling airflow will change with the change of the angle of the guide plate, so it is necessary to calculate the airflow action coefficient in real time to ensure the effectiveness and accuracy of the airflow action. The airflow action coefficient and the down-swirling airflow direction angle data are subjected to aerodynamic analysis to obtain the initial angle value of the guide plate. Through this process, an initial position of the guide plate relative to the duct is obtained. At the same time, the flow field characteristics of the downward swirling airflow direction angle data are analyzed, and the mapping relationship between the deflection angle of the guide plate and the generated axial force component is established to obtain the aerodynamic characteristic parameters. The deflection angle of the guide plate will directly affect the attitude correction effect of the aircraft. The aerodynamic characteristics of the guide plate at different deflection angles can be accurately understood through the flow field characteristics analysis, so as to optimize and adjust the guide plate angle in the future. Aerodynamic characteristic parameters are an important basis for achieving precise attitude control, which can help the control system predict the effect of the guide plate and adjust the angle of the guide plate based on actual needs. According to the current flight mode and aerodynamic characteristic parameters, the initial angle value of the guide plate is subjected to attitude decoupling calculation to obtain the independent control amount of each guide plate. The overall attitude control demand is decomposed into each guide plate, so that each guide plate can be controlled independently to achieve more flexible attitude adjustment. The calculation of the independent control amount takes into account the specific role of each guide plate in different attitude adjustments, so that the system can more accurately control the deflection angle of each guide plate to adapt to different flight requirements. The independent control amount of each deflector is collaboratively optimized to obtain the target angle control amount of the deflector. The coordination between the deflectors is ensured, so that the overall attitude adjustment of the aircraft is more coordinated, and the attitude imbalance will not be caused by the independent control of each deflector. Through collaborative optimization, mutual compensation and coordination between the deflectors are achieved, and the accuracy and stability of attitude control are improved. Based on the target angle control amount, the angular displacement of the deflector servo is servo-controlled to obtain the position control amount of each deflector. The servo control calculation process includes real-time feedback and adjustment of the steering angle of the servo to ensure that the deflector can reach the target angle quickly and accurately.

[0040] Step 500: Based on the current flight mode, three guide plates are configured as triangular control points, and the position compensation amounts of the aileron, yaw rudder and elevator are determined by vector synthesis calculation to obtain the angle control amount of each rudder surface;

[0041] Specifically, the spatial position coordinates of the three deflectors are calculated to construct a triangular control plane. The three deflectors are regarded as independent control points on the control plane, and the plane attitude parameters are obtained. The plane attitude parameters can accurately reflect the attitude control requirements of the aircraft, and can provide more accurate attitude adjustment when the three deflectors are used as control points. Based on the plane attitude parameters, the horizontal plane control requirements are subjected to three-axis vector decomposition operations to obtain the roll axis moment component, the yaw axis moment component, and the pitch axis moment component. The overall control requirements of the aircraft are refined into the moment components in three directions, corresponding to roll, yaw, and pitch, respectively. According to the current flight mode, the rudder response characteristics of the moment components of the roll axis, yaw axis, and pitch axis are analyzed respectively. The aileron rudder response characteristics of the roll axis moment component are analyzed to obtain the initial deflection of the aileron. Based on the current flight mode, the yaw rudder rudder response characteristics of the yaw axis moment component are analyzed to obtain the initial deflection of the yaw rudder. The yaw rudder is used to adjust the yaw motion of the aircraft. By analyzing the yaw axis torque component, the initial deflection angle required for the yaw rudder is calculated to ensure that the aircraft can achieve ideal performance in yaw control. The elevator control surface response characteristics of the pitch axis torque component are analyzed to obtain the initial deflection of the elevator. The elevator is used to control the pitch motion of the aircraft. The analysis of the pitch axis torque component helps to determine the appropriate deflection angle of the elevator to ensure that the aircraft can remain stable during pitch control. The control surface coupling effect compensation calculation is performed on the initial deflection of the aileron, the initial deflection of the yaw rudder, and the initial deflection of the elevator. Multiple control surfaces of the aircraft will interact with each other during the actual control process. The cross-influence between these control surfaces is corrected by coupling effect compensation to obtain a more accurate control surface deflection angle. Through compensation calculation, it is ensured that each control surface coordinates with each other when controlling the aircraft attitude, avoiding the decrease in control accuracy due to the control surface coupling effect, and improving the stability and accuracy of the aircraft. Based on the target rudder surface deflection angle, the control parameters of each rudder surface actuator are linearly transformed, and the rudder surface target deflection angle is converted into a specific execution signal to obtain the angle control amount of each rudder surface to drive the rudder surface to make corresponding angle adjustments.

[0042] Step 600: convert the target rotation speed of each ducted blade, the position control value of each guide plate and the angle control value of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control command to achieve closed-loop attitude control of the aircraft.

[0043] Specifically, the target speed of each ducted blade is converted into a PWM control signal of the ducted motor by a motor drive signal conversion calculation. For the position control amount of the deflector, the steering gear drive signal conversion calculation is performed to convert the target position into a PWM control signal of the deflector steering gear. Similarly, the conversion of the rudder control signal is converted into a PWM control signal of the aileron, yaw rudder and elevator rudder by a steering gear drive signal conversion calculation, and the target angle is converted into the PWM control signal of the aileron, yaw rudder and elevator rudder, which will directly drive the corresponding rudder for precise adjustment during the flight of the aircraft. Real-time collection and execution feedback data are collected for closed-loop control. For the speed control of the ducted motor, the Hall sensor signal and the encoder signal are collected and filtered in real time to obtain the actual speed data of the duct. Similarly, for the position control of the deflector and the rudder rudder, the steering gear sensor signal is collected and filtered in real time to obtain the actual angle data of the rudder. The actual speed data of the duct is compared with the target speed to obtain the speed control error, and the current angle data of the rudder is compared with the target angle to obtain the angle control error. Based on the speed control error and angle control error, a preset control algorithm is used for dynamic parameter adjustment. The preset control algorithm can dynamically adjust the control instructions to make it more in line with actual flight requirements by weighting the proportional, integral and differential effects of the error. The PID controller adjusts the control signal in real time according to the changes in the speed error and angle error to minimize the error and ensure the stability and accuracy of the aircraft in complex flight environments. This control algorithm can effectively minimize the difference between the actual state of the aircraft and the target attitude, and realize closed-loop attitude control.

[0044] In the embodiment of the present application, a design scheme combining a three-duct layout and a guide plate is adopted to eliminate the tilting mechanism, simplify the mechanical structure, improve the reliability of the system, and achieve more efficient power output through the coordinated control of the three ducts. By adopting an innovative control strategy that configures the three guide plates as triangular control points, precise control of the horizontal attitude of the aircraft is achieved, reducing the complexity of the control system. A fault-tolerant control scheme based on power redundancy is designed. When a single duct fails, the power demand can be automatically allocated to the remaining normally operating ducts to ensure the continuous controllability of the aircraft. A complete closed-loop control system is established, and control instructions are corrected through real-time feedback data, which improves the accuracy of attitude control and the robustness of the system. Smooth switching between vertical take-off and landing and horizontal flight modes is achieved, and the stability of the flight process is ensured through real-time monitoring of the pitch angle and mode judgment. The use of multi-sensor data fusion technology and extended Kalman filter algorithm improves the accuracy of state estimation and provides reliable data support for attitude control.

[0045] In a specific embodiment, the process of executing step 100 may specifically include the following steps:

[0046] The original signal output by the inertial measurement unit in the aircraft is processed by low-pass filtering to obtain the attitude angle data of the pitch angle, roll angle and yaw angle of the aircraft relative to the ground;

[0047] Perform Kalman filter calculation on the output signals of the acceleration sensor in the x-axis, y-axis and z-axis directions to obtain acceleration data in three orthogonal axes;

[0048] The original angular velocity signal output by the gyroscope is subjected to zero bias compensation calculation to obtain the angular velocity data of the aircraft around three orthogonal axes, and the output signal of the barometric altimeter is subjected to data fusion calculation with the GPS altitude signal to obtain the altitude data;

[0049] The dynamic pressure signal output by the pitot tube and the GPS ground speed signal are used for speed calculation to obtain speed data. The output signals of the Hall sensors of the three ducts are sampled and processed by median filtering to obtain the duct speed data.

[0050] Based on the attitude angle data, the acceleration data, the angular velocity data, the altitude data, the speed data and the duct speed data, an extended Kalman filter algorithm is used to perform state estimation calculation to obtain flight state parameters, which include pitch angle parameters, current position parameters, current speed parameters and current acceleration parameters.

[0051] Specifically, the attitude angle data of the aircraft is low-pass filtered. The original signal output by the inertial measurement unit contains noise and high-frequency interference, and low-pass filtering is performed on it to extract effective attitude information. The output of the inertial measurement unit includes acceleration data and angular velocity data, which are combined to calculate the pitch angle, roll angle and yaw angle of the aircraft. Assume that the acceleration signal output by the inertial measurement unit is a(t)=(a x (t),a y (t),a z (t)), where a x 、a y 、a z They are the acceleration components of the aircraft on three orthogonal axes. By low-pass filtering, high-frequency noise is removed to obtain a relatively stable acceleration signal, which is used to calculate the attitude angle of the aircraft. For the angular velocity signal, assume that the angular velocity signal output by the IMU is ω(t) = (ω x (t),ω y (t),ω z(t)), and after low-pass filtering, relatively smooth angular velocity data is obtained. Using the acceleration data and angular velocity data, the attitude angle (pitch angle, roll angle and yaw angle) of the aircraft is derived through integration and rotation matrix. The pitch angle (θ), roll angle (φ) and yaw angle (ψ) of the aircraft are calculated by the following formula:

[0052]

[0053]

[0054]

[0055] These formulas use the three components of acceleration data to derive the attitude angle of the aircraft through certain geometric relationships. Kalman filter calculations are performed on the data of the acceleration sensor. Kalman filtering is a recursive data processing algorithm that can extract valuable information from noise. Kalman filtering estimates the state of a system by fusing multiple sensor data. The output signal of the acceleration sensor is the acceleration component a of three orthogonal axes. x (t), a y (t0 and a z (t). Assume that the output of the acceleration sensor is a(t)=(a x (t),a y (t),a z (t)), Kalman filter is used to estimate the acceleration state of the system Right now:

[0056]

[0057] Where K(t) is the Kalman gain, which represents the weighting factor between the system prediction and the actual measurement, a(t) is the actual measurement value, is the estimated value of the previous moment. Through Kalman filtering, more accurate acceleration data can be obtained Perform zero bias compensation on the angular velocity output by the gyroscope. Assume that the output of the gyroscope is The angular velocity after zero bias compensation is ω(t), which is calculated by the following formula:

[0058]

[0059]

[0060]

[0061] Among them, b x 、b y 、b zis the zero bias value of the gyroscope, obtained through calibration. The attitude of the aircraft is integrated according to the compensated angular velocity data to obtain its current attitude angle. In order to obtain the altitude data of the aircraft, the output signal of the barometric altimeter and the GPS signal are fused. The signal output by the barometric altimeter is the air pressure data P(t), and the altitude data provided by GPS is the position altitude H GPS (t). The fusion of the two uses the Kalman filter algorithm, and the fused height data is calculated using the following formula:

[0062]

[0063] Among them, K f (t) is the fusion gain, which represents the weight of the air pressure data and GPS data. The two data are fused through Kalman filtering to obtain the aircraft altitude data. The dynamic pressure signal of the pitot tube and the ground speed signal of GPS are used to calculate the speed. The pitot tube measures the dynamic pressure q(t) of the aircraft, and the ground speed data provided by GPS is the speed V GPS (t). The dynamic pressure and ground speed signals are solved by physical formulas to obtain the true speed of the aircraft. Assume that the dynamic pressure and ground speed data are q(t) and V GPS (t), and is solved using the following formula:

[0064]

[0065] The duct speed data is sampled and processed by the Hall sensor and median filter to obtain the actual duct speed data. Assume that the output signal of the duct speed sensor is Ω raw (t), the speed after median filtering is Ω(t), which can be obtained by calculating the median filter:

[0066] Ω(t)=median(Ω raw (t-1),Ω raw (t),Ω raw (t+1));

[0067] The rudder position data is sampled and filtered by the rudder position sensor to obtain the angle data of each rudder. Through the above processing and filtering, the flight state parameters of the aircraft are obtained, including the pitch angle, current position, current speed and acceleration parameters. Based on the sensor data and filtering processing, the aircraft uses the extended Kalman filter algorithm for state estimation. The extended Kalman filter is a recursive algorithm that can combine multiple sensor data to estimate the state parameters of the aircraft. During the state estimation process, the state variables of the aircraft include the pitch angle, current position, current speed and acceleration, etc. These parameters are continuously corrected through the Kalman filter to obtain accurate flight status. The basic formula of the extended Kalman filter is:

[0068]

[0069] in, is the state estimate at the current moment, z(t) is the actual measurement value, is the predicted value obtained through the system model, and K(t) is the Kalman gain. Through multiple iterative calculations, accurate flight state parameters are obtained.

[0070] In a specific embodiment, the process of executing step 200 may specifically include the following steps:

[0071] Compare and calculate the pitch angle data in the flight state parameter with the first preset angle threshold to determine whether the switching condition of the vertical take-off and landing mode is met, and compare and calculate the pitch angle data in the flight state parameter with the second preset angle threshold to determine whether the switching condition of the horizontal flight mode is met;

[0072] When the corresponding switching conditions are met and a mode switching instruction is received from the autopilot module or the external control module, the corresponding flight mode switching is executed;

[0073] Based on the current flight mode, the coordinate difference between the current position parameter in the flight state parameter and the target position in three orthogonal axes is calculated to obtain position deviation data;

[0074] Based on the current flight mode, the speed difference between the current speed parameter in the flight state parameter and the target speed in three orthogonal axes is calculated to obtain speed deviation data;

[0075] Based on the current flight mode, the acceleration difference between the current acceleration parameter in the flight state parameter and the target acceleration in three orthogonal axes is calculated to obtain acceleration deviation data;

[0076] The position deviation data, the velocity deviation data and the acceleration deviation data are combined to obtain a motion state deviation matrix in three-axis directions, and the motion state deviation matrix is ​​input into a dynamic feedback controller for state space solution to obtain a posture deviation amount.

[0077] Specifically, by comparing the pitch angle data in the flight state parameter with the preset angle threshold, different flight mode determinations can be obtained. Assume that the pitch angle data of the aircraft is θ, and the first preset angle threshold is θ VTOL (Threshold for determining vertical take-off and landing mode), determined by the following conditions:

[0078]

[0079] Among them, M VTOLIt is the vertical take-off and landing mode determination data. If the pitch angle θ is greater than the preset threshold θ VTOL , the system enters the vertical take-off and landing mode, and determines the data M VTOL The value of is 1, otherwise it is 0. Similarly, for the determination of the horizontal flight mode, the pitch angle data is compared with the second preset angle threshold. Assuming that the second preset threshold is θ H (Threshold for determining horizontal flight mode), the determination data are as follows:

[0080]

[0081] Among them, M H is the judgment data of the horizontal flight mode. When the pitch angle is less than the preset threshold θ H When the system enters the horizontal flight mode, the judgment data M H The value of is 1, otherwise it is 0. The judgment data M for the vertical take-off and landing mode VTOL and the judgment data M of the horizontal flight mode H Perform logical operations to obtain the current flight mode of the aircraft. The flight mode determination result is comprehensively determined by logical operators (such as AND, OR, NOT, etc.) to determine the vertical take-off and landing mode and the horizontal flight mode. For example, the following logical expression is used to determine whether it is in the vertical take-off and landing mode:

[0082]

[0083] Among them, M flight Indicates the current flight mode of the aircraft, ∧ indicates logical AND operation, Indicates logical negation. When the vertical take-off and landing mode determination data is 1 and the horizontal flight mode determination data is 0, the aircraft is in the vertical take-off and landing mode, otherwise it is not in the vertical take-off and landing mode. Based on the current flight mode, the attitude deviation is calculated. The coordinate difference between the current position parameter in the flight state parameter and the target position in three orthogonal axes is calculated to obtain the position deviation data. Assume that the current position parameter is P current =(x current ,y current ,z current ), and the target position is P target =

[0084] x target ,y target ,z target ), the position deviation data ΔP is calculated by the following formula:

[0085] ΔP=P target -P current =(x target -x current ,ytarget -y current ,z target -z current );

[0086] Among them, ΔP = (Δx, Δy, Δz) represents the position deviation of the aircraft on the three orthogonal axes, representing the position error in the x, y, and z directions respectively. At the same time, based on the current flight mode, the speed difference between the current speed parameter in the flight state parameter and the target speed in the three orthogonal axes is calculated to obtain the speed deviation data. Assume that the current speed parameter is V current =

[0087] The target speed is The speed deviation data ΔV is calculated by the following formula:

[0088]

[0089] Where, ΔV=(Δv x ,Δv y ,Δv z ) represents the velocity deviation of the aircraft on three orthogonal axes. Calculate the acceleration deviation. Assume that the current acceleration parameter is The target acceleration is The acceleration deviation data ΔA is calculated by the following formula:

[0090]

[0091] Among them, ΔA=*Δa x ,Δa y ,Δa z ) represents the acceleration deviation of the aircraft on the three orthogonal axes. The position deviation, velocity deviation and acceleration deviation are combined to obtain the motion state deviation matrix in the three-axis direction. Assuming that the position deviation, velocity deviation and acceleration deviation are ΔP, ΔV, and ΔA respectively, the motion state deviation matrix E is expressed as:

[0092]

[0093] The motion state deviation matrix E contains the deviation information of the position, velocity and acceleration of the aircraft on three orthogonal axes. By inputting the motion state deviation matrix into the dynamic feedback controller, the state space is solved to obtain the attitude deviation. Assume that the dynamic feedback controller uses the state space model for solution and defines the state vector as x(t) = [Δx, Δv x ,Δa x ,Δy,Δv y ,Δa y ,Δz,Δvz ,Δa z ] T , the control input is u(t), the system matrix is ​​A, and the input matrix is ​​B, then the state equation of the system is expressed as:

[0094]

[0095] In the state space model, the motion state deviation is fed back through the dynamic feedback controller, the state of the aircraft is corrected in real time, and the corresponding attitude deviation ΔΘ is obtained. Through the solution, the output signal of the aircraft attitude adjustment is obtained so that the command control surface can be adjusted accordingly, thereby reducing the attitude deviation and achieving attitude stability control of the aircraft.

[0096] Among them, the position deviation data, velocity deviation data and acceleration deviation data are combined to obtain the motion state deviation matrix in the three-axis direction, and the motion state deviation matrix is ​​input into the dynamic feedback controller for state space solution to obtain the posture deviation, including: setting the first weight coefficient, the second weight coefficient and the third weight coefficient for the position deviation data, the velocity deviation data and the acceleration deviation data, combining the weighted deviation data in the three directions of the x-axis, the y-axis and the z-axis to obtain the motion state deviation matrix in the three-axis direction; constructing the position loop control layer of the dynamic feedback controller, the position loop control layer includes the first PID controller and the first state observer, the first PID controller adopts a cascade structure, and the first state observer adopts a Kalman filter structure; the first PID controller includes three parallel independent loops, each loop corresponds to an axial position control, each loop includes three parameters of proportional gain, integral gain and differential gain, and the first state observer includes a state prediction module and a state correction module; constructing the speed loop control layer of the dynamic feedback controller, the speed loop control The layer includes a second PID controller and a second state observer, wherein the second PID controller adopts a feedforward compensation structure, and the second state observer adopts an extended Kalman filter structure; the second PID controller includes three adaptive gain modules, each module dynamically adjusts the control parameters according to the rate of change of the velocity deviation, and the second state observer includes a nonlinear state prediction module and a measurement update module; an acceleration loop control layer of a dynamic feedback controller is constructed, the acceleration loop control layer includes a third PID controller and a feedback linearization module, the third PID controller adopts a gain scheduling structure, and the feedback linearization module adopts a dynamic inverse system structure; the motion state deviation matrix is ​​input into the position loop control layer, the velocity loop control layer and the acceleration loop control layer in sequence, and the state space solution is realized through a cascade control method; the output of the position loop control layer is used as a given value of the velocity loop control layer, and the output of the velocity loop control layer is used as a given value of the acceleration loop control layer; the optimal control calculation is performed based on the output result of the acceleration loop control layer, and the control signal is optimized by a linear quadratic regulator to obtain the attitude deviation.

[0097] In a specific embodiment, the process of executing step 300 may specifically include the following steps:

[0098] Perform state analysis on the current flight mode and attitude deviation to obtain the overall power demand in the Z-axis direction and the attitude torque demand in the three orthogonal axes;

[0099] According to the overall power demand, the power distribution of the three ducts is calculated evenly to obtain the reference power distribution coefficient of each duct;

[0100] Perform fault detection calculation on the deviation between the real-time speed data of the three ducts and the rated speed to obtain the duct working state identification information, and judge the working state of the three ducts according to the duct working state identification information. When any duct is detected to have a fault, perform proportional distribution calculation on the power demand of the abnormal duct based on the duct working state identification information to obtain the compensation power coefficient of the remaining normal working ducts;

[0101] According to the reference power distribution coefficient and the compensation power coefficient, the power output of the three ducts is dynamically adjusted and calculated to obtain the target power output of each duct, and the duct blade characteristic mapping operation is performed on the target power output to obtain the corresponding duct speed control amount;

[0102] Based on the duct speed control quantity, PID control calculation is performed on the torque of each duct motor to obtain the motor control compensation quantity, and the duct speed control quantity and the motor control compensation quantity are superimposed to obtain the target speed of each duct blade.

[0103] Specifically, the current flight mode and attitude deviation are analyzed and calculated to obtain the overall power demand in the Z-axis direction and the attitude torque demand in the three orthogonal axes. Assume that the attitude deviation of the aircraft is ΔΘ=(Δφ,Δθ,Δψ), where Δφ, Δθ and Δψ are the deviations of the roll angle, pitch angle and yaw angle of the aircraft respectively. Assume that the current flight mode is M flight , its value is 1 for vertical take-off and landing mode, and 0 for horizontal flight mode. Through state analysis calculation, the total power demand F in the Z-axis direction is obtained z And the posture torque requirements M of the three orthogonal axes x ,M y ,M z ). The power requirement F in the Z-axis direction z , calculated by the following formula:

[0104] F z =F gravity +K p Δθ;

[0105] Among them, F gravity represents the gravity of the aircraft, K p is a proportional gain used to adjust the deviation of the aircraft in the pitch direction. x 、M y and M z , then we get it through the following formula:

[0106] M x =K φ ·Δφ,M y =K θ ·Δθ,M z =K ψ Δψ;

[0107] Among them, K φ , K θ and K ψ are the proportional gain coefficients for roll, pitch and yaw, respectively, indicating the response strength of the system to each attitude deviation. z , evenly distribute the power to the three ducts. Assuming the three ducts are duct A, duct B and duct C, the total power demand F z The power distribution coefficient is expressed as:

[0108]

[0109] Among them, k A , k B , k C Represent the reference power distribution coefficients of ducts A, B and C respectively. Under normal circumstances, the power output of the three ducts is equal, so that the aircraft can maintain a stable hovering or flying state. Perform fault detection calculation on the real-time speed data of the duct and the rated speed to obtain the working status identification information of the duct. Assume that the real-time speed of duct A is Rated speed is Ω rated The goal of fault detection is to compare the deviation between the real-time speed and the rated speed. The fault judgment conditions are defined as follows:

[0110]

[0111]

[0112] Among them, ΔΩ A represents the speed deviation of duct A, ∈ is the threshold of fault detection, S A is the working status identification information of duct A, 1 indicates normal, 0 indicates fault. Similarly, for ducts B and C, the corresponding status identification S is obtained. B and SC When any duct failure is detected, the power demand of the abnormal duct needs to be distributed in equal proportion to maintain the stability of the aircraft. Assume that duct A fails (i.e. S A =0), at this time, ducts B and C need to compensate for the power output of duct A. Based on the working status identification information of the duct, the compensation power coefficient is expressed as:

[0113]

[0114] Among them, k′ B and k′ C Represent the compensation power coefficients of duct B and duct C respectively. By distributing the power demand of the abnormal duct to the other two ducts in equal proportion, the aircraft can maintain sufficient power output to maintain its attitude stability. According to the reference power distribution coefficient and the compensation power coefficient, the power output of the three ducts is dynamically adjusted and calculated to obtain the target power output of each duct. The target power output is expressed as:

[0115] F A =k A ·S A ,F B = k′ B ·S B ,F C = k′ C ·S C ;

[0116] Among them, F A 、F B 、F C are the target power outputs of ducts A, B and C respectively. For a normally functioning duct, its target power output is calculated based on the reference power distribution coefficient and the compensation power coefficient, while for a faulty duct, its power output is zero. The target power output is mapped to the duct blade characteristics to obtain the corresponding duct speed control amount. Assume that the relationship between the duct power output and the blade speed is expressed by the following formula:

[0117]

[0118] Among them, Ω i represents the speed control quantity of duct i, K T is the speed-power conversion coefficient of the duct, F i is the target power output. Through this formula, the speed control amount of each duct is obtained. In order to control the torque output of the duct motor, the speed of each duct is calculated by PID control. Assume that the target speed of the duct is Ω target , the actual speed is Ω actual , the speed control error is:

[0119] e(t)=Ω target -Ω actual ;

[0120] The control quantity of the PID controller is expressed as:

[0121]

[0122] Among them, u(t) is the motor control compensation, K p , K i and K d are proportional, integral and differential gain coefficients respectively. The speed control error is processed by the PID controller to obtain the control compensation of the motor, so that the motor can reach the target speed more accurately. The duct speed control amount and the motor control compensation amount are superimposed to obtain the target speed of each duct blade:

[0123] Ω final =Ω i +u(t);

[0124] Through this formula, the target speed of the duct can take into account both the power demand and the correction amount of the real-time speed, thereby ensuring that the power system of the aircraft can be dynamically adjusted according to actual conditions and achieve precise attitude control.

[0125] In a specific embodiment, the process of executing step 400 may specifically include the following steps:

[0126] According to the current flight mode, the attitude deviation is decomposed and calculated in the three axes of U, V, and W to obtain the expected correction forces in the three directions, and the duct airflow guidance calculation is performed on the expected correction forces to obtain the down-swirl airflow direction angle data of the three ducts;

[0127] Based on the current flight mode, the effective area of ​​the guide plate below the duct and the downswirling airflow is dynamically calculated to obtain the airflow action coefficient. The aerodynamic analysis of the airflow action coefficient and the downswirling airflow direction angle data is performed to obtain the initial angle value of the guide plate.

[0128] The flow field characteristics of the downward swirling airflow direction angle data are analyzed, and the mapping relationship between the deflection angle of the guide plate and the generated axial force component is established to obtain the aerodynamic characteristic parameters;

[0129] According to the current flight mode and aerodynamic characteristic parameters, the initial angle value of the guide plate is subjected to attitude decoupling calculation to obtain the independent control amount of each guide plate, and the independent control amount of each guide plate is subjected to collaborative optimization calculation to obtain the target angle control amount of the guide plate;

[0130] Based on the target angle control quantity, the angular displacement of the deflector servo is calculated by servo control to obtain the position control quantity of each deflector.

[0131] Specifically, according to the current flight mode, the attitude deviation is decomposed and calculated in the three axes of U, V, and W to obtain the expected correction force in the three directions. Assume that the attitude deviation of the aircraft is ΔΘ = (Δφ, Δθ, Δψ), where Δφ, Δθ and Δψ are the attitude deviations of the aircraft in the roll, pitch and yaw directions respectively. The three axes of U, V, and W represent the lateral, longitudinal and vertical directions in the aircraft's own coordinate system respectively. The expected correction force is expressed as F desired =(F U ,F V ,F W ). According to the relationship between attitude deviation and flight mode, the expected correction force is decomposed and calculated by the following formula:

[0132] F U =K φ ·Δφ,F V =K θ ·Δθ,F W =K ψ Δψ;

[0133] Among them, K φ , K θ , K ψ are the proportional gain coefficients in the roll, pitch and yaw directions, respectively, which are used to convert the attitude deviation into a specific correction force. The duct airflow guidance calculation is performed on the expected correction force to obtain the down-swirl airflow direction angle data of the three ducts. Assuming that the airflow under the duct is controlled by the guide plate, its airflow direction angle is expressed as α i Where i represents the number of the duct (for example, 1, 2, and 3 correspond to three ducts). The relationship between the flow angle of the downward swirling airflow and the desired correction force is expressed by the following formula:

[0134]

[0135] Among them, F Ui 、F Vi 、F Wi are the force components of the i-th duct in the U, V, and W directions respectively. This formula decomposes the correction force in the direction of the downswirling airflow to obtain the angle that the guide plate needs to be adjusted to affect the direction of the airflow. Based on the results of the flight mode judgment, the effective area of ​​the guide plate and the downswirling airflow below the duct is dynamically calculated to obtain the airflow action coefficient. The airflow action coefficient is used to describe the control effect of the guide plate on the downswirling airflow, and its size is related to the effective area of ​​the guide plate, the speed of the downswirling airflow, and the angle of the guide plate. Assume that the airflow action area is A i , the speed of the downward swirling airflow is Vair , the angle of the guide plate is β i , then the airflow coefficient C i It is expressed as:

[0136] C i =A i ·V air ·cos(β i );

[0137] Among them, A i Represents the effective area of ​​the guide plate, V air is the velocity of the downward swirling airflow, cos(β i ) represents the influence of the guide plate angle on the airflow direction. By performing aerodynamic analysis on the airflow action coefficient and the downward swirling airflow direction angle data, the initial angle value β of the guide plate is obtained. i , so that the deflector generates appropriate aerodynamic force in the current flight mode, thereby achieving attitude correction. The flow field characteristics of the downward swirling airflow direction angle data are analyzed, and the mapping relationship between the deflection angle of the deflector and the axial force generated is established to obtain the aerodynamic characteristic parameters. Assume that the deflection angle of the deflector is γ i , the aerodynamic characteristic parameter is k f , then the relationship between the deflection angle of the guide plate and the generated axial force is expressed as:

[0138] F axial =k f ·sin(γ i );

[0139] Among them, F axial It represents the component force generated by the guide plate in the axial direction, sin(γ i ) represents the influence of the deflection angle of the guide plate on the aerodynamic force. Through the analysis of the flow field characteristics, the aerodynamic characteristic parameters are obtained to reflect the mechanical performance of the guide plate at different angles. According to the current flight mode and aerodynamic characteristic parameters, the initial angle value of the guide plate is subjected to attitude decoupling calculation to obtain the independent control amount of each guide plate. The overall attitude control requirements of the aircraft are decomposed into each guide plate, so that each guide plate can independently adjust the downward swirling airflow to achieve more flexible attitude adjustment. Assume that the initial angle of the guide plate is β i , the aerodynamic characteristic parameter is k fi , then the independent control quantity of the guide plate is expressed by the following formula:

[0140] δ i =k fi β i ;

[0141] Among them, δ iRepresents the independent control amount of the i-th deflector. By calculating the independent control amount of each deflector, it is ensured that the angle adjustment of each deflector meets the attitude adjustment requirements of the aircraft. In order to coordinate the adjustment actions of each deflector, the independent control amount is coordinated and optimized to obtain the target angle control amount of the deflector, ensuring the coordination between all deflectors to avoid the incoordination caused by independent control, thereby affecting the overall attitude stability of the aircraft. Assume that the independent control amounts of all deflectors are δ1, δ2, δ3, and the target angle control amount is β target , then the weighted average method is used to perform collaborative optimization operations:

[0142]

[0143] Among them, w1, w2, w3 are the weight coefficients of each deflector, indicating the contribution of each deflector to the overall attitude adjustment. Through collaborative optimization calculation, the target angle control of the deflector is obtained, thereby ensuring that the attitude control of the aircraft is more stable and coordinated. Based on the target angle control, the angular displacement of the deflector servo is servo-controlled to obtain the position control of each deflector. The goal of the servo control of the servo is to enable the deflector to reach the target angle quickly and accurately. Assuming that the target angle control is β target , the actual angular displacement of the servo is θ servo , then the servo control error is expressed as:

[0144] e(t)=β target -θ servo ;

[0145] The PID controller is used to process the servo control error, and its control quantity is expressed as:

[0146]

[0147] Where u(t) is the control compensation of the servo, K p , K i and K d They are proportional, integral and differential gain coefficients respectively. The PID controller ensures that the deflector can quickly and stably reach the target angle control amount by adjusting the angular displacement of the servo in real time.

[0148] According to the current flight mode and aerodynamic characteristic parameters, the initial angle value of the guide plate is subjected to attitude decoupling calculation to obtain the independent control amount of each guide plate, and the independent control amount of each guide plate is subjected to collaborative optimization operation to obtain the target angle control amount of the guide plate, including: setting the first mode matrix and the second mode matrix for the vertical take-off and landing mode data and the horizontal flight mode data in the current flight mode, respectively, setting the first characteristic matrix and the second characteristic matrix for the aerodynamic coefficient and the airflow guide angle in the aerodynamic characteristic parameters, respectively, combining the first mode matrix, the second mode matrix, the first characteristic matrix and the second characteristic matrix to obtain the initial state matrix of the guide plate angle control; inputting the initial state matrix into the first decoupling control matrix. A controller and a second decoupling controller, wherein the first decoupling controller performs singular value decomposition on the row vectors of the matrix, and the second decoupling controller performs singular value decomposition on the column vectors of the matrix, and orthogonalizes the two decomposition results to obtain a characteristic vector group of the deflector attitude control; a first control component and a second control component are set for the characteristic vector group, wherein the first control component represents the angle control requirement in the vertical take-off and landing state, and the second control component represents the angle control requirement in the horizontal flight state, and a principal component analysis is performed based on the first control component and the second control component to obtain the independent action components of each deflector in three orthogonal axes; a first influence coefficient and a second influence coefficient are calculated for the independent action components, wherein the first influence coefficient represents the effect of the change in the deflector angle on the aerodynamics The sensitivity of the moment, the second influence coefficient characterizes the sensitivity of the change in the guide plate angle to the intensity of the downward swirling airflow, the first influence coefficient and the second influence coefficient are weightedly calculated to obtain the weight distribution coefficient of the guide plate angle; the first weight value and the second weight value are set according to the weight distribution coefficient, wherein the first weight value is used to adjust the priority of aerodynamic moment balance, and the second weight value is used to adjust the priority of the downward swirling airflow intensity, the first weight value and the second weight value are weighted with the initial angle value of the guide plate to obtain the reference control angle of each guide plate; the first constraint condition and the second constraint condition are set for the reference control angle, wherein the first constraint condition ensures that the deflection angle of the guide plate does not exceed the mechanical limit, and the second constraint condition ensures that the aerodynamic moment meets the balance requirements, The reference control angle is substituted into the multi-objective optimization solver for iterative calculation to obtain the optimized angle solution of the guide plate; the optimized angle solution is input into the first compensation model and the second compensation model, wherein the first compensation model corrects the airflow interference between the guide plates, and the second compensation model corrects the dynamic response characteristics of the guide plates, and the output results of the two compensation models are fused to obtain the corrected angle value of the guide plate; based on the corrected angle value, a first response prediction function and a second response prediction function are constructed, wherein the first response prediction function is used to evaluate the transient characteristics of the guide plate angle change, and the second response prediction function is used to evaluate the steady-state characteristics of the guide plate angle change, and the calculation results of the two prediction functions are combined and optimized to obtain the target angle control amount of the guide plate.

[0149] In a specific embodiment, the process of executing step 500 may specifically include the following steps:

[0150] The triangular control plane is constructed and calculated for the spatial position coordinates of the three deflectors. The three deflectors are used as three independent control points on the control plane to obtain the plane attitude parameters. Based on the plane attitude parameters, the three-axis vector decomposition operation is performed on the horizontal plane control requirements to obtain the roll axis moment component, the yaw axis moment component and the pitch axis moment component.

[0151] According to the current flight mode, the aileron control surface response characteristics of the roll axis moment component are analyzed to obtain the initial aileron deflection;

[0152] According to the current flight mode, the yaw rudder surface response characteristics of the yaw axis moment component are analyzed to obtain the initial deflection of the yaw rudder;

[0153] According to the current flight mode, the elevator control surface response characteristics of the pitch axis moment component are analyzed to obtain the initial deflection of the elevator;

[0154] The rudder surface coupling effect compensation calculation is performed on the initial deflection of the aileron, the initial deflection of the yaw rudder and the initial deflection of the elevator to obtain the target rudder surface deflection angle;

[0155] Based on the target rudder surface deflection angle, the control parameters of each rudder surface actuator are linearly transformed to obtain the angle control value of each rudder surface.

[0156] Specifically, the spatial position coordinates of the three deflectors are used to construct a triangular control plane. Assume that the spatial position coordinates of the three deflectors are P1 = (x1, y1, z1), P2 = (x2, y2, z2) and P3 = (x3, y3, z3). These coordinates describe the position of the deflectors on the aircraft relative to the center of gravity. In order to construct the control plane, the coordinates of the three deflectors are used as three independent control points on the plane to form a triangular control plane. The attitude parameters of the control plane are calculated through the position coordinates of these three control points. In order to describe the attitude of the plane, the normal vector n is constructed through the three vertices of the triangle. The normal vector represents the direction of the control plane. The normal vector is calculated using the following formula:

[0157] n=(P2-P1)×(P3-P1);

[0158] Among them, × represents the cross product operation of the vector, P2-P1 and P3-P1 are two edge vectors. The calculated n=(n x ,n y ,n z) is the normal vector of the control plane, which describes the direction of the plane in space. Based on the plane attitude parameters, the horizontal plane control demand is decomposed into three axes to obtain the rolling axis moment component, the yaw axis moment component and the pitch axis moment component. Assume that the control demand vector of the aircraft is T = T x ,T y ,T z ), which represents the total torque requirement of the aircraft in the three axes. By projecting with the plane normal vector, the components in the roll axis, yaw axis and pitch axis are obtained. The roll axis torque component M φ , yaw axis moment component M ψ and the pitch axis moment component M θ The calculation formula is:

[0159] M φ =T x ·n x ,M ψ =T y ·n y ,M θ =T z ·n z ;

[0160] Among them, T x 、T y 、T z are the components of the total control demand vector on the three axes, and n x 、n y 、n z are the components of the normal vector, thus decomposing the total control demand into three directions: roll, yaw and pitch. According to the current flight mode, the aileron control surface response characteristics of the roll axis torque component are analyzed to obtain the initial deflection of the aileron. Assuming the current flight mode is M flight , when the aircraft is in horizontal flight mode (i.e. M flight =1), rolling moment component M φ It will cause the aileron to deflect. Assume the response coefficient of the aileron is The initial deflection of the aileron is calculated by the following formula:

[0161]

[0162] Among them, δ φ is the deflection angle of the aileron, is the roll response characteristic coefficient, which describes the response degree of the aileron to the roll moment. Similarly, for the yaw axis moment component M ψ , it is necessary to analyze the response characteristics of the yaw rudder. Assume that the response coefficient of the yaw rudder is Then the initial deflection of the yaw rudder is expressed as:

[0163]

[0164] Among them, δ ψ is the deflection angle of the yaw rudder. Similarly, for the pitch axis moment component M θ , the initial deflection of the elevator is calculated by the following formula:

[0165]

[0166] Among them, δ θ is the deflection angle of the elevator, is the pitch response characteristic coefficient. After obtaining the initial deflection of the aileron, yaw rudder and elevator, the rudder coupling effect compensation calculation is performed on these initial deflections to obtain the target rudder deflection angle. Since the multiple rudders of the aircraft have coupling effects in actual operation, the deflection of one rudder will affect the effects of other rudders, so coupling compensation is required. Assume that the coupling effect coefficients of the aileron, yaw rudder and elevator are C φψ , C φθ , C φθ , C ψθ , then the target rudder surface deflection angle is calculated by the following formula:

[0167]

[0168]

[0169]

[0170] in, are the target deflection angles of aileron, yaw rudder and elevator respectively. The attitude of the aircraft is controlled more accurately by considering the coupling effect. After the target deflection angle of the rudder surface is obtained, the control parameters of each rudder surface actuator are linearly transformed to obtain the angle control amount of each rudder surface. Assume that the actuator of each rudder surface has a control gain K act , then the angle control amount of each rudder surface is calculated by the following formula:

[0171]

[0172]

[0173]

[0174] in, are the control angles of aileron, yaw rudder and elevator, respectively, K act It is the control gain coefficient of the actuator, which is used to convert the target deflection angle into an actual control signal.

[0175] In a specific embodiment, the process of executing step 600 may specifically include the following steps:

[0176] The target speed of each ducted blade is converted into a motor drive signal to obtain the PWM control signal of the ducted motor; the position control amount of each deflector is converted into a servo drive signal to obtain the PWM control signal of the deflector servo; the angle control amount of each rudder surface is converted into a servo drive signal to obtain the PWM control signal of the aileron, yaw rudder and elevator servo;

[0177] The Hall sensor signal and encoder signal of the duct motor are collected and filtered in real time to obtain the actual speed data of the duct, and the sensor signals of the deflector servo and the rudder servo are collected and filtered in real time to obtain the current angle data of the servo;

[0178] The actual duct speed data is compared with the duct target speed to obtain the speed control error, and the current angle data of the servo is compared with the servo target angle to obtain the angle control error;

[0179] Based on the speed control error and the angle control error, a preset control algorithm is used to adjust the dynamic parameters and make real-time corrections to the control instructions to achieve closed-loop attitude control of the aircraft.

[0180] Specifically, the target speed of each ducted blade is converted into a driving signal of the ducted motor so as to control the duct of the aircraft through the motor. The drive of the ducted motor is realized using a pulse width modulation (PWM) signal, and the goal is to convert the target speed of the ducted blade to Ω target Converted into PWM control signal. Assuming the target speed of the ducted blade is Ω target , the maximum speed of the motor is Ω max , then the PWM duty cycle of the motor is motor The conversion is calculated using the following formula:

[0181]

[0182] Among them, PWM motor Indicates the PWM duty cycle of the motor, Ω target is the target speed, Ω max is the maximum allowable speed of the motor. This formula converts the target speed to the maximum speed to obtain the PWM signal for driving the ducted motor. Similarly, for the position control of the deflector, it is converted into the PWM control signal of the servo. Assume that the position control of the deflector is θ target , the maximum deflection angle of the servo is θ max , then the PWM duty cycle of the servo is servo Calculated by the following formula:

[0183]

[0184] Among them, θ target is the target deflection angle of the deflector, θ max is the maximum deflection angle of the servo, PWM servo represents the PWM duty cycle of the servo. The PWM signal is used to control the deflection angle of the servo to achieve position control of the deflector. Similarly, for the angle control of the aileron, yaw rudder and elevator of the aircraft, the servo drive signal is converted and calculated to obtain the corresponding PWM control signal. Suppose the target deflection angles of the aileron, yaw rudder and elevator are and The maximum deflection angle of each rudder surface is δ max , then the PWM control signals of the aileron, yaw rudder and elevator are expressed as:

[0185]

[0186]

[0187]

[0188] Through the calculation of these PWM signals, the target angle of each rudder surface is converted into an actual control signal, thereby driving the corresponding actuator to adjust the attitude. The Hall sensor signal and encoder signal of the duct motor are collected in real time and filtered to obtain the actual speed data of the duct. Assuming that the Hall sensor signal is Ω hall (t), the encoder signal is Ω enc (t), the average value of the two is expressed as the real-time speed data of the duct Ω actual (t), is obtained by filtering through the following formula:

[0189]

[0190] The actual speed is denoised to obtain more accurate real-time speed data. The sensor signals of the deflector servo and the rudder servo are collected and filtered in real time to obtain the actual angle data of the servo. Assume that the output signal of the position sensor is θ sensor (t), after low-pass filtering, the actual deflection angle θ of the servo is obtained actual (t), the filtering process is expressed as:

[0191]

[0192] Among them, LPF represents a low-pass filter, which is used to remove high-frequency noise in the signal to ensure the smoothness and accuracy of the deflection angle data. The actual speed of the duct and the actual deflection angle of the rudder are compared with the target value to obtain the control error. The speed control error of the duct is expressed as:

[0193] e Ω (t) = Ω target -Ω actual (t);

[0194] Among them, e Ω (t) is the speed control error of the duct, Ω target is the target speed, Ω actual (t) is the actual speed collected in real time. Similarly, for the angle control of the rudder, the error is expressed as:

[0195] e θ (t) = θ target -θ actual (t);

[0196] Among them, e θ (t) is the angle control error of the rudder surface, θ target is the target deflection angle, θ actual (t) is the actual deflection angle. In order to achieve closed-loop attitude control, a preset control algorithm is used to adjust dynamic parameters based on the speed control error and the angle control error, so as to correct the control command in real time. The preset control algorithm corrects the error through proportional, integral and differential actions, and the output of the PID controller is expressed as:

[0197]

[0198]

[0199] Among them, u Ω (t) and u θ (t) are the control compensation of ducted motor and rudder, K p , K i and K d are the proportional, integral and differential gain coefficients, respectively, indicating the intensity of the system's response to the error. Through the preset control algorithm, the error is corrected in real time to ensure that the ducted motor and the rudder servo can quickly follow the target control quantity and achieve the attitude stability control of the aircraft. The output of the PID controller is superimposed with the original control signal to obtain the PWM control signal actually used for the actuator. For example, the final PWM control signal of the ducted motor is expressed as:

[0200]

[0201] The final PWM control signal of the rudder surface is expressed as:

[0202]

[0203] Through these control signals, the ducted motor and control surface servo can be adjusted in real time to meet the attitude control requirements of the aircraft in different flight modes.

[0204] The above describes the attitude control method of the aircraft in the embodiment of the present application. The following describes the attitude control device 10 of the aircraft in the embodiment of the present application. Figure 2 In one embodiment of the present application, an attitude control device 10 for an aircraft includes:

[0205] The acquisition module 11 is used to acquire the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight state parameters;

[0206] The difference calculation module 12 is used to compare the pitch angle data in the flight state parameters with the preset angle threshold to determine whether the flight mode switching condition is met. When the switching condition is met and a mode switching instruction issued by the autopilot module or the external control module is received, the corresponding flight mode switching is executed, and the difference between the current attitude and the target attitude is calculated to obtain the attitude deviation;

[0207] The power distribution calculation module 13 is used to perform power distribution calculations on the three ducts respectively based on the current flight mode and attitude deviation, and obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts;

[0208] The position optimization calculation module 14 is used to perform position optimization calculation on the guide plates below the duct according to the current flight mode and attitude deviation, and obtain the position control amount of each guide plate;

[0209] The vector synthesis calculation module 15 is used to configure the three deflectors as triangular control points based on the current flight mode, determine the position compensation of the aileron, the yaw rudder and the elevator through vector synthesis calculation, and obtain the angle control amount of each rudder surface;

[0210] The execution module 16 is used to convert the target rotation speed of each ducted blade, the position control value of each guide plate and the angle control value of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control command to realize the closed-loop attitude control of the aircraft.

[0211] Through the cooperation of the above components, the design scheme of three-duct layout and guide plate combination is adopted to eliminate the tilting mechanism, simplify the mechanical structure, improve the reliability of the system, and achieve more efficient power output through the coordinated control of the three ducts. Through the innovative control strategy of configuring the three guide plates as triangular control points, the horizontal attitude of the aircraft is precisely controlled, and the complexity of the control system is reduced. A fault-tolerant control scheme based on power redundancy is designed. When a single duct fails, the power demand can be automatically allocated to the remaining normally working ducts to ensure the continuous controllability of the aircraft. A complete closed-loop control system is established, and the control command correction is carried out through real-time feedback data, which improves the accuracy of attitude control and the robustness of the system. Smooth switching between vertical take-off and landing and horizontal flight modes is achieved, and the stability of the flight process is ensured through real-time monitoring of the pitch angle and mode judgment. The multi-sensor data fusion technology and extended Kalman filter algorithm are used to improve the accuracy of state estimation and provide reliable data support for attitude control.

[0212] See also Figure 3 , Figure 3 This is a schematic block diagram of the structure of an electronic device 300 provided in an embodiment of the present application. The electronic device 300 includes a processor 301 and a memory 302. The processor 301 and the memory 302 are connected via a device bus 303, wherein the memory 302 may include a non-volatile storage medium and an internal memory.

[0213] The non-volatile storage medium may store a computer program. The computer program includes program instructions, and when the program instructions are executed by the processor 301, the processor 301 may execute any of the above-mentioned attitude control methods for an aircraft.

[0214] The processor 301 is used to provide computing and control capabilities to support the operation of the entire electronic device 300 .

[0215] The internal memory provides an environment for the operation of the computer program in the non-volatile storage medium. When the computer program is executed by the processor 301, the processor 301 can execute any of the above-mentioned attitude control methods for the aircraft.

[0216] Those skilled in the art will understand that Figure 3 The structure shown in the figure is only a block diagram of a partial structure related to the present application scheme, and does not constitute a limitation on the electronic device 300 involved in the present application scheme. The specific electronic device 300 may include more or fewer components than shown in the figure, or combine certain components, or have a different component arrangement.

[0217] It should be understood that the processor 301 may be a central processing unit (CPU), and the processor 301 may also be other general-purpose processors, digital signal processors (DSP), application-specific integrated circuits (ASIC), field-programmable gate arrays (FPGA) or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. Among them, the general-purpose processor may be a microprocessor or the processor may also be any conventional processor, etc.

[0218] It should be noted that those skilled in the art can clearly understand that, for the convenience and simplicity of description, the specific working process of the electronic device 300 described above can refer to the corresponding process of the attitude control method of the aforementioned aircraft, and will not be repeated here.

[0219] An embodiment of the present application also provides a computer-readable storage medium, which stores a computer program. When the computer program is executed by one or more processors, the one or more processors implement the attitude control method of the aircraft provided in the embodiment of the present application.

[0220] The computer-readable storage medium may be an internal storage unit of the electronic device 300 in the aforementioned embodiment, such as a hard disk or memory of the electronic device 300. The computer-readable storage medium may also be an external storage device of the electronic device 300, such as a plug-in hard disk, a smart memory card (Smart Media Card, SMC), a secure digital (Secure Digital, SD) card, a flash card (Flash Card), etc. equipped with the electronic device 300.

[0221] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working processes of the systems, devices and units described above can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0222] If the integrated unit is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present application is essentially or the part that contributes to the prior art or all or part of the technical solution can be embodied in the form of a software product, and the computer software product is stored in a storage medium, including a number of instructions to enable an electronic device (which can be a personal computer, a server, or a network device, etc.) to perform all or part of the steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM), random access memory (RAM), disk or optical disk and other media that can store program code.

[0223] As described above, the above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit them. Although the present application has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A method for controlling the attitude of an aircraft, characterized in that: The attitude control method of the aircraft comprises: Collect the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight status parameters; Compare the pitch angle data in the flight state parameter with a preset angle threshold to determine whether a flight mode switching condition is met; when the switching condition is met and a mode switching instruction issued by an autopilot module or an external control module is received, execute the corresponding flight mode switching, and calculate the difference between the current attitude and the target attitude to obtain an attitude deviation; Based on the current flight mode and the attitude deviation, power distribution calculation is performed on the three ducts respectively to obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts; According to the current flight mode and the attitude deviation, the position of the guide plate below the duct is optimized and calculated to obtain the position control amount of each guide plate; Based on the current flight mode, the three deflectors are configured as triangular control points, and the position compensation of the aileron, yaw rudder and elevator is determined by vector synthesis calculation to obtain the angle control value of each rudder surface; The target rotation speed of each ducted blade, the position control amount of each guide plate and the angle control amount of each control surface are converted into execution signals, and execution feedback data is obtained to perform real-time correction of control instructions to achieve closed-loop attitude control of the aircraft.

2. The method for controlling the attitude of an aircraft according to claim 1, characterized in that: The flight state parameters are obtained by collecting the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft, including: The original signal output by the inertial measurement unit in the aircraft is processed by low-pass filtering to obtain the attitude angle data of the pitch angle, roll angle and yaw angle of the aircraft relative to the ground; Perform Kalman filter calculation on the output signals of the acceleration sensor in the x-axis, y-axis and z-axis directions to obtain acceleration data in three orthogonal axes; The original angular velocity signal output by the gyroscope is subjected to zero bias compensation calculation to obtain the angular velocity data of the aircraft around three orthogonal axes, and the output signal of the barometric altimeter is subjected to data fusion calculation with the GPS altitude signal to obtain the altitude data; The dynamic pressure signal output by the pitot tube and the GPS ground speed signal are used for speed calculation to obtain speed data. The output signals of the Hall sensors of the three ducts are sampled and processed by median filtering to obtain the duct speed data. Based on the attitude angle data, the acceleration data, the angular velocity data, the altitude data, the speed data and the duct speed data, an extended Kalman filter algorithm is used to perform state estimation calculation to obtain flight state parameters, which include pitch angle parameters, current position parameters, current speed parameters and current acceleration parameters.

3. The method for controlling the attitude of an aircraft according to claim 2, characterized in that: The pitch angle data in the flight state parameter is compared with a preset angle threshold to determine whether the flight mode switching condition is met. When the switching condition is met and a mode switching instruction issued by an autopilot module or an external control module is received, the corresponding flight mode switching is executed, and the difference between the current attitude and the target attitude is calculated to obtain an attitude deviation, including: Compare and calculate the pitch angle data in the flight state parameter with the first preset angle threshold to determine whether the switching condition of the vertical take-off and landing mode is met, and compare and calculate the pitch angle data in the flight state parameter with the second preset angle threshold to determine whether the switching condition of the horizontal flight mode is met; When the corresponding switching conditions are met and a mode switching instruction is received from the autopilot module or the external control module, the corresponding flight mode switching is executed; Based on the current flight mode, the coordinate difference between the current position parameter in the flight state parameter and the target position in three orthogonal axes is calculated to obtain position deviation data; Based on the current flight mode, the speed difference between the current speed parameter in the flight state parameter and the target speed in three orthogonal axes is calculated to obtain speed deviation data; Based on the current flight mode, the acceleration difference between the current acceleration parameter in the flight state parameter and the target acceleration in three orthogonal axes is calculated to obtain acceleration deviation data; The position deviation data, the velocity deviation data and the acceleration deviation data are combined to obtain a motion state deviation matrix in three-axis directions, and the motion state deviation matrix is ​​input into a dynamic feedback controller for state space solution to obtain a posture deviation amount.

4. The method for controlling the attitude of an aircraft according to claim 3, characterized in that: The method of performing power distribution calculation on the three ducts based on the current flight mode and the attitude deviation to obtain the target rotation speed of each duct blade, and distributing the power demand of the duct to the remaining two ducts when any duct fails, includes: Perform state analysis calculation on the current flight mode and the attitude deviation to obtain the overall power demand in the Z-axis direction and the attitude torque demand in the three orthogonal axes; According to the overall power demand, a power uniform distribution calculation is performed on the three ducts to obtain a reference power distribution coefficient of each duct; Performing fault detection calculation on the deviation between the real-time speed data of the three ducts and the rated speed to obtain the duct working state identification information, judging the working state of the three ducts according to the duct working state identification information, and when any duct is detected to have a fault, performing proportional distribution calculation on the power demand of the abnormal duct based on the duct working state identification information to obtain the compensation power coefficient of the remaining normal working ducts; According to the reference power distribution coefficient and the compensation power coefficient, the power output of the three ducts is dynamically adjusted and calculated to obtain the target power output of each duct, and the duct blade characteristic mapping operation is performed on the target power output to obtain the corresponding duct speed control amount; Based on the duct speed control amount, PID control calculation is performed on the torque of each duct motor to obtain the motor control compensation amount, and the duct speed control amount and the motor control compensation amount are superimposed to obtain the target speed of each duct blade.

5. The method for controlling the attitude of an aircraft according to claim 4, characterized in that: According to the current flight mode and the attitude deviation, the guide plate below the duct is optimized and calculated to obtain the position control amount of each guide plate, including: According to the current flight mode, the attitude deviation is decomposed and calculated in the three axes of U, V, and W to obtain the expected correction forces in the three directions, and the duct airflow guidance calculation is performed on the expected correction forces to obtain the down-swirl airflow direction angle data of the three ducts; Based on the current flight mode, dynamically calculate the effective area of ​​the guide plate below the duct and the downswirling airflow to obtain the airflow effect coefficient, perform aerodynamic analysis on the airflow effect coefficient and the downswirling airflow direction angle data to obtain the initial angle value of the guide plate; Performing flow field characteristic analysis on the downward swirling airflow direction angle data, establishing a mapping relationship between the deflection angle of the guide plate and the generated axial force component, and obtaining aerodynamic characteristic parameters; According to the current flight mode and the aerodynamic force characteristic parameters, an attitude decoupling calculation is performed on the initial angle value of the guide plate to obtain an independent control amount of each guide plate, and a collaborative optimization calculation is performed on the independent control amount of each guide plate to obtain a target angle control amount of the guide plate; Based on the target angle control amount, the angular displacement of the deflector servo is servo-controlled and calculated to obtain the position control amount of each deflector.

6. The method for controlling the attitude of an aircraft according to claim 5, characterized in that: Based on the current flight mode, the three deflectors are configured as triangular control points, and the position compensation of the aileron, yaw rudder and elevator is determined by vector synthesis calculation to obtain the angle control amount of each rudder surface, including: The spatial position coordinates of the three deflectors are calculated to construct a triangular control plane, and the three deflectors are used as three independent control points on the control plane to obtain plane attitude parameters. Based on the plane attitude parameters, a three-axis vector decomposition operation is performed on the horizontal plane control requirements to obtain a roll axis moment component, a yaw axis moment component, and a pitch axis moment component; According to the current flight mode, analyzing the aileron control surface response characteristics of the roll axis moment component to obtain the aileron initial deflection; According to the current flight mode, analyzing the response characteristics of the yaw rudder surface of the yaw axis moment component to obtain the initial deflection of the yaw rudder; According to the current flight mode, analyzing the elevator control surface response characteristics of the pitch axis moment component to obtain the initial elevator deflection; Performing rudder surface coupling effect compensation calculation on the initial deflection of the aileron, the initial deflection of the yaw rudder and the initial deflection of the elevator to obtain a target rudder surface deflection angle; Based on the target rudder surface deflection angle, the control parameters of each rudder surface actuator are linearly transformed to obtain the angle control amount of each rudder surface.

7. The method for controlling the attitude of an aircraft according to claim 6, characterized in that: The target rotation speed of each ducted blade, the position control amount of each guide plate and the angle control amount of each control surface are converted into an execution signal, and execution feedback data is obtained to perform real-time correction of control instructions to achieve closed-loop attitude control of the aircraft, including: The target speed of each ducted blade is converted into a motor drive signal to obtain a PWM control signal of the ducted motor; the position control amount of each deflector is converted into a steering gear drive signal to obtain a PWM control signal of the deflector steering gear; the angle control amount of each rudder surface is converted into a steering gear drive signal to obtain a PWM control signal of the aileron, yaw rudder and elevator rudder; The Hall sensor signal and encoder signal of the duct motor are collected and filtered in real time to obtain the actual speed data of the duct, and the sensor signals of the deflector servo and the rudder servo are collected and filtered in real time to obtain the current angle data of the servo; Compare the actual duct speed data with the duct target speed to obtain a speed control error, and compare the current angle data of the servo with the servo target angle to obtain an angle control error; Based on the rotation speed control error and the angle control error, a preset control algorithm is used to perform dynamic parameter adjustment, and the control instructions are corrected in real time to achieve closed-loop attitude control of the aircraft.

8. An attitude control device for an aircraft, characterized in that: Used to execute the attitude control method of an aircraft according to any one of claims 1 to 7, the attitude control device of the aircraft comprises: The acquisition module is used to collect the attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data of the aircraft to obtain flight status parameters; A difference calculation module is used to compare the pitch angle data in the flight state parameters with a preset angle threshold to determine whether a flight mode switching condition is met. When the switching condition is met and a mode switching instruction issued by the autopilot module or the external control module is received, the corresponding flight mode switching is executed, and a difference calculation is performed between the current attitude and the target attitude to obtain an attitude deviation; A power distribution calculation module is used to perform power distribution calculations on the three ducts respectively based on the current flight mode and the attitude deviation, and obtain the target rotation speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts; A position optimization calculation module is used to perform position optimization calculation on the guide plates below the duct according to the current flight mode and the attitude deviation, so as to obtain the position control amount of each guide plate; A vector synthesis calculation module is used to configure the three deflectors as triangular control points based on the current flight mode, determine the position compensation of the aileron, yaw rudder and elevator through vector synthesis calculation, and obtain the angle control amount of each rudder surface; The execution module is used to convert the target rotation speed of each ducted blade, the position control value of each guide plate and the angle control value of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control command to realize the closed-loop attitude control of the aircraft.

9. An electronic device, characterized in that: The electronic device comprises: a memory and at least one processor, wherein instructions are stored in the memory; The at least one processor calls the instructions in the memory to enable the electronic device to execute the attitude control method of the aircraft according to any one of claims 1-7.

10. A computer-readable storage medium having instructions stored thereon, characterized in that: When the instructions are executed by a processor, the attitude control method of an aircraft as described in any one of claims 1 to 7 is implemented.

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