Aircraft attitude control method, device, equipment and storage medium

Abstract

The present application relates to the field of aircraft technology, and discloses an attitude control method, device, equipment and storage medium for an aircraft, the method comprising: collecting flight status parameters of the aircraft; comparing pitch angle data in the flight status parameters with a preset angle threshold, executing corresponding flight mode switching, and calculating attitude deviation; performing power distribution calculations on three ducts respectively to obtain target rotation speeds of blades of each duct; performing position optimization calculations on guide plates below the ducts to obtain position control quantities of each guide plate; configuring the three guide plates as triangular control points, determining position compensation quantities of ailerons, yaw rudders and elevators through vector synthesis calculations to obtain angle control quantities of each rudder surface; converting the results into execution signals, and obtaining execution feedback data to perform real-time correction of control instructions to achieve closed-loop attitude control of the aircraft, thereby improving the accuracy of attitude control of the aircraft and ensuring the stability of the flight process.

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 (VTOL) aircraft primarily utilize tilt-rotor or composite wing configurations, but these solutions suffer from complex mechanical structures and low reliability. Tilt-rotors require complex tilting mechanisms to transition between vertical take-off and landing and horizontal flight, increasing system weight and the risk of failure. Composite wing configurations, on the other hand, require an additional lift system, resulting in redundant powertrains and reduced energy efficiency.

[0003] Existing vertical take-off and landing (VTOL) aircraft also face significant challenges in attitude control. During the transition between vertical take-off and landing (VTOL) and horizontal flight modes, the aircraft must simultaneously coordinate multiple actuators, including the power system, control surfaces, and attitude control devices. This creates a complex control system and places stringent demands on control precision. Furthermore, when a partial powertrain failure occurs, existing control solutions often struggle to maintain the aircraft's attitude stability, posing a significant safety hazard. Summary of the Invention

[0004] The present application provides an attitude control method, apparatus, device 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 aircraft's attitude angle data, acceleration data, angular velocity data, altitude data, speed data, and ducted speed data to obtain flight status parameters;

[0007] Comparing the pitch angle data in the flight state parameters with a preset angle threshold to determine whether a flight mode switching condition is met, and when the switching condition is met and a mode switching instruction is received from the autopilot module or the external control module, executing the corresponding flight mode switching, and calculating 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 for each of the three ducts to obtain a target rotational speed for each ducted propeller blade. If any duct fails, the power demand of the duct is distributed to the remaining two ducts.

[0009] performing position optimization calculations on the deflector plates below the duct according to the current flight mode and the attitude deviation, and obtaining position control values ​​for each deflector plate;

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

[0011] The target rotation speed of each ducted blade, the position control quantity of each guide plate and the angle control quantity 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 aircraft's attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data to obtain flight status parameters;

[0014] a difference calculation module, configured 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 satisfied; and when the switching condition is satisfied and a mode switching instruction is received from the autopilot module or the external control module, execute the corresponding flight mode switching and calculate the difference between the current attitude and the target attitude to obtain an attitude deviation;

[0015] a power distribution calculation module, configured to perform power distribution calculations on each of the three ducts based on the current flight mode and the attitude deviation, to obtain a target rotational speed for each ducted propeller blade, and to distribute the power demand of any duct to the remaining two ducts if a duct failure occurs;

[0016] a position optimization calculation module, configured to perform position optimization calculation on the guide plates below the duct according to the current flight mode and the attitude deviation, and obtain a position control value for each guide plate;

[0017] The 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 value of each control surface;

[0018] The execution module is used to convert the target rotation speed of each ducted blade, the position control quantity of each guide plate and the angle control quantity of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control instructions to achieve 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 existing technology, the present application has the following beneficial effects: the design scheme of a three-duct layout combined with a guide plate eliminates the tilting mechanism, simplifies the mechanical structure, improves the reliability of the system, and achieves more efficient power output through the coordinated control of the three ducts. By adopting an innovative control strategy in which the three guide plates are configured 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 distributed to the remaining normally operating ducts, ensuring the continuous controllability of the aircraft. A complete closed-loop control system is established, and control instructions are corrected through real-time feedback data, thereby improving 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. 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 following briefly introduces the drawings required for use in the embodiments or the description of the prior art. 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 any creative work.

[0023] The structures, proportions, sizes, etc. depicted 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 intended to limit the conditions under which the present invention can be implemented and therefore have no substantive technical significance. Any structural modifications, changes in proportional relationships, or adjustments in size should still fall within the scope of the technical contents disclosed in the present invention without affecting the effects and objectives that can be achieved by the present invention.

[0024] Figure 1 1 is a flow chart of a method for controlling an aircraft's attitude according to 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 clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of them. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0028] The flowcharts shown in the accompanying drawings are for illustrative purposes only and do not necessarily include all contents and operations / steps, nor must they be executed in the order described. For example, some operations / steps may be decomposed, combined, or partially merged, so the actual execution order may vary depending on the actual situation.

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

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

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

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

[0033] Specifically, the raw signal output by the aircraft's inertial measurement unit is low-pass filtered to eliminate the impact of high-frequency noise on the attitude angle data, thereby obtaining the aircraft's pitch, roll, and yaw angles relative to the ground. Accelerometers are used to collect data in the x, y, and z axes. The accelerometer output signals are processed using a Kalman filter algorithm to effectively filter out noise and obtain accurate acceleration data along the three orthogonal axes. By combining a dynamic model with an observation model, the Kalman filter algorithm adjusts the weights of the filtering process in real time, ensuring that the processed data more accurately reflects the aircraft's actual acceleration. Furthermore, bias compensation is performed on the raw angular velocity signal output by the gyroscope to eliminate systematic errors caused by sensor offset. This results in angular velocity data describing the aircraft's rotational state about the three orthogonal axes. The barometric altimeter output signal is fused with the GPS altitude signal to obtain altitude data. Velocity is calculated by combining the dynamic pressure signal output by the pitot tube with the GPS ground speed signal. The pitot tube provides information on the speed relative to the air, while the GPS module provides information on the speed of the aircraft relative to the ground. By jointly calculating these two signals, accurate speed data is obtained. The output signals of the Hall sensors of the three ducts are sampled and median filtered to remove occasional spike noise, making the duct speed data smoother and more stable, and obtaining duct speed data. The output signals of the deflector position sensor, aileron position sensor, yaw rudder position sensor, and elevator position sensor are sampled and filtered to obtain real-time position data of the rudder surface, and effectively remove 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, speed data, and duct 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 that 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 status parameters with a preset angle threshold to determine whether a flight mode switching condition is met. If the switching condition is met and a mode switching instruction is received from the autopilot module or the external control module, execute the corresponding flight mode switching and calculate the difference between the current attitude and the target attitude to obtain an attitude deviation.

[0035] Specifically, the pitch angle data in the flight status parameters is compared with a preset angle threshold to determine whether the flight mode switching conditions are met. If the switching conditions are met and a mode switch command is received from the autopilot module or external control module, the corresponding flight mode switch is executed. A logical operation is performed on the vertical take-off and landing mode determination data and the horizontal flight mode determination data to determine 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 and avoid misjudgments due to interference between different modes. Different flight modes have different control requirements and control objectives. For example, in vertical take-off and landing mode, the aircraft requires higher thrust and stability, while in horizontal flight mode, the focus is on flight direction control and speed stability. Based on the current flight mode, the coordinate difference between the current position parameter in the flight status parameters and the target position along three orthogonal axes is calculated to obtain position deviation data, which reflects the spatial displacement of the aircraft. Based on the current flight mode, the speed difference between the current velocity parameter (a flight state parameter) and the target velocity along three orthogonal axes is calculated to generate velocity deviation data. This provides the control system with information on the velocity deviation during the aircraft's motion, enabling it to adjust thrust magnitude and direction to achieve the target velocity. Simultaneously, based on the current flight mode, the acceleration parameter (a flight state parameter) and the target acceleration along three orthogonal axes are calculated to generate acceleration deviation data. Acceleration deviation data describes the difference between the aircraft's acceleration change and the target acceleration. Especially during takeoff, acceleration, deceleration, and landing, acceleration deviations directly reflect the aircraft's force conditions and motion state, providing a basis for power system adjustments. The position deviation data, velocity deviation data, and acceleration deviation data are combined to generate a three-axis motion state deviation matrix, which comprehensively reflects the overall difference between the aircraft's current motion state 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 analyzes the deviation matrix and, in conjunction with the system's dynamic model, calculates the required attitude adjustments for the aircraft, achieving closed-loop feedback for attitude control. Through state-space computation, the control system comprehensively considers the differences between the aircraft's motion state and its target state, and based on this, formulates a reasonable attitude adjustment strategy to gradually approach the target attitude and achieve stable flight control. The calculated attitude deviation is used to adjust the power distribution of the ducted blades, the position control of the deflectors, and the angle compensation of the control surfaces, ensuring that the aircraft maintains a stable attitude in various flight modes, such as vertical takeoff and landing, cruising, and horizontal flight, thereby improving flight safety and accuracy.

[0036] Step 300: Based on the current flight mode and attitude deviation, power distribution calculation is performed for each of the three ducts to obtain a target rotational speed for each ducted propeller blade. If any duct fails, the power demand of that duct is distributed to the remaining two ducts.

[0037] It should be noted that a state analysis calculation is performed based on the current flight mode and attitude deviation to obtain the total power demand in the Z-axis direction and the attitude torque requirements in three orthogonal axes. The total power demand in the Z-axis reflects the thrust required to maintain vertical hovering or stable flight, while the attitude torque requirements describe the torque requirements in the three axes, thereby ensuring precise attitude control. Based on the total power demand in the Z-axis, power is evenly distributed among the three ducts to obtain a baseline power distribution coefficient for each duct. This ensures that under normal conditions, the three ducts operate with balanced power output to maintain aircraft stability and attitude control. During this power distribution process, the aircraft's mass distribution and structural characteristics are taken into account to ensure reasonable and effective power distribution for each duct. The real-time duct speed data is compared with the rated speed, and fault detection calculations are performed based on this comparison to obtain duct operating status information. This duct operating status information allows real-time monitoring of the operating status of each duct and fault diagnosis. During flight, if a duct failure is detected in any one duct, power compensation is applied to the remaining ducts to ensure that the aircraft's attitude control is not significantly impacted. Based on the duct operating status information, when a duct failure is detected, the power demand of the faulty duct is proportionally distributed to calculate the compensation power coefficients for the remaining functioning ducts. This compensation ensures that even if a duct failure occurs, the aircraft's overall attitude can still be stabilized through power adjustments in the remaining ducts, improving the system's fault tolerance and reliability. Based on the baseline power distribution coefficient and the compensation power coefficient, the power output of the three ducts is dynamically adjusted to determine the target power output for each duct. This ensures that each duct can flexibly adjust its power output based on the current flight state and fault conditions, maintaining overall aircraft stability and flight control. The target power output is mapped to the duct blade characteristics to obtain the corresponding duct speed control variable. Converting the power demand to a specific duct blade speed control requires considering multiple factors, including the aerodynamic characteristics of the duct blades, drag characteristics, and the interaction between the duct and the air, to ensure that the speed control variable effectively reflects the required power output. Based on the ducted speed control variable, PID control calculations are performed on the torque of each ducted motor to determine the motor's control compensation. PID control adjusts the motor's output torque in real time based on the speed error, ensuring the ducted motor can quickly respond to changes in the target speed while minimizing steady-state error and overshoot, achieving precise control of the ducted motor. After the PID control calculations are completed, the ducted speed control variable is superimposed with the motor control compensation to determine the target speed for each ducted blade. By comprehensively considering the baseline speed requirement and control compensation, the actual output of the ducted motor is more closely aligned with the aircraft's current attitude control requirements.

[0038] Step 400: Optimize the position of the deflectors below the duct according to the current flight mode and attitude deviation to obtain the position control value of each deflector.

[0039] Specifically, based on the current flight mode, the attitude deviation is decomposed and calculated along the three axes (U, V, and W) to obtain the desired correction forces in these three directions. The overall attitude adjustment requirement is broken down into specific axial correction forces to ensure that the aircraft achieves reasonable correction results in all directions. The desired correction forces are then used to calculate the duct airflow guidance to obtain the flow angle data for the ducted downswirling airflow. The downswirling airflow direction angle data is directly related to the ducted airflow's guidance performance. The primary function of the deflector is to achieve aircraft attitude adjustment by changing the direction of the downswirling airflow. Based on the current flight mode, the interaction area between the deflector below the duct and the downswirling airflow is dynamically calculated to obtain the airflow action coefficient. The interaction area between the deflector and the downswirling airflow changes with the deflector angle, so the airflow action coefficient must be calculated in real time to ensure the effectiveness and accuracy of the airflow action. Aerodynamic analysis is performed on the airflow action coefficient and the downswirling airflow direction angle data to determine the initial angle value of the deflector. This process determines the initial position of the deflector relative to the duct. Simultaneously, the downswirling airflow angle data is analyzed for flow field characteristics, establishing a mapping relationship between the deflector deflection angle and the generated axial force component to obtain aerodynamic parameters. The deflector deflection angle directly affects the aircraft's attitude correction. This flow field analysis accurately understands the aerodynamic characteristics of the deflector at different deflection angles, enabling subsequent optimization of the deflector angle. Aerodynamic parameters are crucial for achieving precise attitude control, helping the control system predict the effectiveness of the deflector and adjust the deflector angle based on actual requirements. Based on the current flight mode and aerodynamic parameters, an attitude decoupling calculation is performed on the initial deflector angle values ​​to determine the independent control variables for each deflector. This decomposition of the overall attitude control requirements onto each deflector allows for independent control of each deflector, enabling more flexible attitude adjustment. The calculation of the independent control variables takes into account the specific role of each deflector in different attitude adjustments, enabling the system to more precisely control the deflection angle of each deflector to accommodate diverse flight requirements. The independent control variables of each deflector are collaboratively optimized to determine the target angle control variable for each deflector. This ensures coordination between the individual deflectors, making the overall attitude adjustment of the aircraft more coordinated and preventing attitude imbalance caused by the independent control of each deflector. Through collaborative optimization, mutual compensation and coordination between the deflectors are achieved, improving the accuracy and stability of attitude control. Based on the target angle control variable, servo control calculations are performed on the angular displacement of the deflector servos to determine the position control variable for each deflector. The servo control calculation process includes real-time feedback and adjustment of the servo angle to ensure that the deflectors can quickly and accurately reach the target angle.

[0040] Step 500: Based on the current flight mode, the three deflectors are configured as triangular control points, and the position compensation amounts of the ailerons, yaw rudders, and elevators are determined through vector synthesis calculation to obtain the angle control amount of each control surface;

[0041] Specifically, a triangular control plane is constructed and calculated based on the spatial position coordinates of the three deflectors. The three deflectors are treated as independent control points on the control plane, and the plane attitude parameters are obtained. These plane attitude parameters accurately reflect the aircraft's attitude control requirements and, with the three deflectors as control points, provide more precise attitude adjustment. Based on the plane attitude parameters, a three-axis vector decomposition operation is performed on the horizontal plane control requirements to obtain the roll, yaw, and pitch axis moment components. The overall control requirements of the aircraft are refined into three torque components, corresponding to roll, yaw, and pitch, respectively. Based on the current flight mode, the rudder response characteristics of the roll, yaw, and pitch axis moment components are analyzed. The aileron rudder response characteristics of the roll axis moment component are analyzed to obtain the initial aileron deflection. Based on the current flight mode, the yaw rudder rudder response characteristics of the yaw axis moment component are analyzed to obtain the initial rudder deflection. The yaw rudder is used to adjust the aircraft's yaw motion. By analyzing the yaw axis torque components, the required initial yaw rudder deflection angle is calculated to ensure optimal yaw control performance. The elevator control surface response characteristics of the pitch axis torque components are analyzed to determine the initial elevator deflection. The elevator controls the aircraft's pitch motion. Analysis of the pitch axis torque components helps determine the appropriate elevator deflection angle to ensure stable pitch control. Surface coupling effect compensation is calculated for the initial deflections of the ailerons, yaw rudder, and elevator. Multiple aircraft control surfaces interact with each other during actual control. Coupling effect compensation corrects for the cross-influence between these surfaces and produces more accurate surface deflection angles. This compensation ensures coordination between the control surfaces when controlling the aircraft's attitude, preventing a decrease in control accuracy due to surface coupling effects and improving aircraft stability and accuracy. Based on the target rudder surface deflection angle, the control parameters of each rudder surface actuator are linearly transformed, and the target rudder surface deflection angle is converted into a specific execution signal to obtain the angle control value 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, a motor drive signal conversion calculation is performed for the target speed of each ducted blade, converting the target speed into a PWM control signal for the ducted motor. For the deflector position control variable, a servo drive signal conversion calculation is performed, converting the target position into a PWM control signal for the deflector servo. Similarly, the rudder control signal conversion is performed through a servo drive signal conversion calculation, converting the target angle into PWM control signals for the aileron, yaw, and elevator servos. These signals directly drive the corresponding servos for precise adjustment during flight. Real-time execution feedback data is collected to facilitate closed-loop control. For ducted motor speed control, Hall sensor and encoder signals are collected and filtered in real time to obtain actual ducted speed data. Similarly, for position control of the deflector and rudder servos, servo sensor signals are collected and filtered in real time to obtain actual servo angle data. The actual ducted speed data is compared with the target speed to obtain a speed control error, and the current servo angle data is compared with the target angle to obtain an angle control error. Based on the speed and angle control errors, a preset control algorithm is used for dynamic parameter adjustment. By weighting the proportional, integral, and differential effects of the errors, the preset control algorithm dynamically adjusts control commands to better meet actual flight requirements. The PID controller adjusts the control signal in real time based on changes in speed and angle errors to minimize errors and ensure stability and accuracy in complex flight environments. This control algorithm effectively minimizes the difference between the actual state of the aircraft and the target attitude, achieving 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, which eliminates the tilting mechanism, simplifies the mechanical structure, improves the reliability of the system, and achieves more efficient power output through the coordinated control of the three ducts. By adopting an innovative control strategy in which the three guide plates are configured 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 distributed to the remaining normally operating ducts, ensuring the continuous controllability of the aircraft. A complete closed-loop control system is established, and control command correction is performed through real-time feedback data, thereby improving 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 raw signal output by the inertial measurement unit in the aircraft is low-pass filtered to obtain the attitude angle data of the aircraft relative to the ground, including pitch, roll and yaw angles;

[0047] Perform Kalman filtering calculation on the output signals of the acceleration sensor in the x-axis, y-axis, and z-axis directions to obtain acceleration data in the 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 the three orthogonal axes, and the output signal of the barometric altimeter is fused 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 to calculate the speed data. The output signals of the Hall sensors of the three ducts are sampled and median filtered 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 are the acceleration components of the aircraft on the 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, it is assumed 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 aircraft's attitude angles (pitch, roll, and yaw) are derived through integration and rotation matrices. The aircraft's pitch angle (θ), roll angle (φ), and yaw angle (ψ) are calculated using 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 from the acceleration sensor. Kalman filter is a recursive data processing algorithm that can extract valuable information from noise. Kalman filter 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 at the previous moment. Through Kalman filtering, more accurate acceleration data is 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 pressure data P(t), and the altitude data provided by the 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 velocity is resolved by the dynamic pressure signal of the pitot tube and the GPS ground speed signal. The pitot tube measures the dynamic pressure q(t) of the aircraft, and the ground speed data provided by the GPS is the velocity V GPS (t). The dynamic pressure and ground speed signals are solved by physical formulas to get 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 median filtered by the Hall sensor 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 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 pitch angle, current position, current speed and acceleration, etc. These parameters are continuously corrected through Kalman filtering 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 by 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] Comparing the pitch angle data in the flight state parameters with a first preset angle threshold to determine whether a switching condition to a vertical take-off and landing mode is satisfied, and comparing the pitch angle data in the flight state parameters with a second preset angle threshold to determine whether a switching condition to a horizontal flight mode is satisfied;

[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 parameters with the preset angle threshold, different flight mode judgments 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 judgment 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. Assume that the second preset threshold is θ H (Threshold value used to determine horizontal flight mode), the judgment data is 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 vertical take-off and landing mode VTOL and the judgment data M of the horizontal flight mode H Perform logical operations to determine 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 whether it is in vertical take-off and landing mode or horizontal flight mode. For example, the following logical expression is used to determine whether it is in vertical take-off and landing mode:

[0082]

[0083] Among them, M flight Indicates the current flight mode of the aircraft, ∧ represents the logical AND operation, Indicates logical NOT operation. 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 vertical take-off and landing mode, otherwise it is not in 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 and the target position in the three orthogonal axes in the flight state parameter 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 the 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 using 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 axes. 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. Assuming that the dynamic feedback controller uses the state space model for solution, the state vector is defined 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, a dynamic feedback controller provides feedback on the motion state deviation, correcting the aircraft's state in real time and generating the corresponding attitude deviation ΔΘ. This solution generates an output signal for aircraft attitude adjustment, enabling the command and control surfaces to make corresponding adjustments, thereby reducing the attitude deviation and achieving stable attitude control.

[0096] Among them, the position deviation data, velocity deviation data and acceleration deviation data are combined to obtain a 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 attitude deviation, including: setting a first weight coefficient, a second weight coefficient and a third weight coefficient for the position deviation data, the velocity deviation data and the acceleration deviation data, and combining the weighted deviation data in the three directions of the x-axis, the y-axis and the z-axis to obtain a motion state deviation matrix in the three-axis direction; constructing a position loop control layer of the dynamic feedback controller, the position loop control layer includes a first PID controller and a 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: proportional gain, integral gain and differential gain, and the first state observer includes a state prediction module and a state correction module; constructing a 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 contains 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 contains a nonlinear state prediction module and a measurement update module; an acceleration loop control layer of a dynamic feedback controller is constructed, and 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 cascade control; the output of the position loop control layer is used as the given value of the velocity loop control layer, and the output of the velocity loop control layer is used as the 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 requirement in the Z-axis direction and the attitude torque requirements 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 benchmark power distribution coefficient of each duct;

[0100] A fault detection calculation is performed on the deviation between the real-time speed data of the three ducts and the rated speed to obtain duct operating status identification information. The operating status of the three ducts is judged based on the duct operating status identification information. When a fault is detected in any duct, the power demand of the abnormal duct is proportionally distributed based on the duct operating status identification information to obtain the compensation power coefficient of the remaining normally operating ducts;

[0101] Based on the baseline 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 then mapped to the ducted propeller blade characteristics to obtain the corresponding ducted speed control value.

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

[0103] Specifically, the current flight mode and attitude deviation are analyzed and calculated to obtain the total 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 aircraft's roll angle, pitch angle and yaw angle 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 ). For the power requirement F in the Z-axis direction z , calculated using 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, which represent the response strength of the system to each attitude deviation. z , evenly distribute the power to the three ducts. Assume that the three ducts are duct A, duct B and duct C, and the total power demand F z The power distribution coefficient is expressed as:

[0108]

[0109] Among them, k A 、k B 、k C Represents the base 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 maintains a stable hover or flight state. Fault detection calculation is performed on the real-time speed data of the ducts and the rated speed to obtain the working status identification information of the ducts. 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] Where ΔΩ A represents the speed deviation of duct A, ∈ is the fault detection threshold, S A is the working status identification information of duct A, 1 means normal, 0 means fault. Similarly, for ducts B and C, the corresponding status identification S is obtained. B and SC When any duct is detected to be faulty, 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 for ducts B and C, respectively. By distributing the power demand of the abnormal duct equally to the remaining two ducts, the aircraft can maintain sufficient power output to maintain its attitude stability. Based on the baseline 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 for ducts A, B, and C, respectively. For a normally functioning duct, its target power output is calculated based on the baseline 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 variable. 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 Indicates 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 quantity of each duct is obtained. In order to control the torque output of the duct motor, PID control calculation is performed on the speed of each duct. 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 the proportional, integral, and differential gain coefficients, respectively. The speed control error is processed by the PID controller to obtain the motor control compensation, which enables the motor to reach the target speed more accurately. The ducted speed control value and the motor control compensation are superimposed to obtain the target speed of each ducted 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 real-time speed correction, thereby ensuring that the aircraft's power system can be dynamically adjusted according to actual conditions to achieve precise attitude control.

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

[0126] Based on the current flight mode, the attitude deviation is decomposed and calculated along the three axes of U, V, and W to obtain the desired correction forces in the three directions. The duct airflow guidance calculation is then performed on the desired correction forces to obtain the down-swirl airflow direction angle data for the three ducts.

[0127] Based on the current flight mode, the interaction area between the deflector and the downswirling airflow is dynamically calculated to obtain the airflow action coefficient. Aerodynamic analysis is then performed on the airflow action coefficient and the downswirling airflow direction angle data to determine the initial angle value of the deflector.

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

[0129] Based on the current flight mode and aerodynamic parameters, the initial angle value of the deflector is subjected to attitude decoupling calculation to obtain the independent control value of each deflector. The independent control value of each deflector is then subjected to collaborative optimization calculation to obtain the target angle control value of the deflector.

[0130] Based on the target angle control value, the angular displacement of the deflector servo is calculated by servo control to obtain the position control value 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 ). Based on the relationship between attitude deviation and flight mode, the expected correction force is decomposed and calculated using 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 desired 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, the airflow direction angle is α i Where i represents the duct number (e.g., 1, 2, and 3 correspond to three ducts). The relationship between the down-swirl flow angle 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 adjust in order to affect the direction of the airflow. Based on the judgment results of the 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 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 velocity of the downward swirling airflow is Vair , the angle of the guide plate is β i , then the airflow coefficient C i 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 effect of the guide plate angle on the airflow direction. By performing aerodynamic analysis on the airflow action coefficient and the down-swirl 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 generated axial force component is established to obtain the aerodynamic characteristic parameters. Assuming 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 degree of 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 quantity 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 collaboratively optimized to obtain the target angle control amount of the deflector, ensuring the coordination between all deflectors and avoiding the incoordination caused by independent control, which affects 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 collaborative optimization operation is performed by weighted average method:

[0142]

[0143] Among them, w1, w2, and w3 are the weight coefficients of each deflector, indicating the contribution of each deflector to the overall attitude adjustment. Through collaborative optimization, the target angle control value 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 value, the angular displacement of the deflector servo is servo-controlled to obtain the position control value of each deflector. The goal of the servo control is to enable the deflector to reach the target angle quickly and accurately. Assume that the target angle control value 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] Among them, u(t) is the control compensation of the servo, K p , K i and K d The PID controller adjusts the angular displacement of the servo in real time to ensure that the deflector can quickly and stably reach the target angle control value.

[0148] Among them, 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 are provided, 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 the two decomposition results are orthogonalized to obtain a eigenvector group of the deflector attitude control; a first control component and a second control component are set for the eigenvector 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 torque, 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 weighted 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 the aerodynamic torque 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 torque meets the balance requirements, The baseline 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. 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, the first response prediction function and the 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. The calculation results of the two prediction functions are combined and optimized to obtain the target angle control value 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 based on 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, a 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 axis moment component is analyzed for the yaw rudder surface response characteristics to obtain the initial yaw rudder deflection;

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

[0154] The rudder 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 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 calculated 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 deflector relative to the center of gravity on the aircraft. 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] Where × represents the cross product of vectors, 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-axis vectors to obtain the roll axis moment component, yaw axis moment component and pitch axis moment component. Let the control demand vector of the aircraft be T = T x ,T y ,T z ), which represents the total torque requirement of the aircraft in the three axes. By projecting it with the plane normal vector, the components on 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 aileron response coefficient 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 a coupling effect in actual operation, the deflection of one rudder will affect the effect of other rudders, so coupling compensation is required. Let the coupling effect coefficients of the aileron, yaw rudder and elevator be C respectively. φψ 、C φθ 、C φθ 、C ψθ , then the target rudder deflection angle is calculated by the following formula:

[0167]

[0168]

[0169]

[0170] in, are the target deflection angles of the aileron, yaw rudder, and elevator, respectively. By considering the coupling effect, the attitude of the aircraft is controlled more accurately. After obtaining the target deflection angle of the rudder surface, the control parameters of each rudder surface actuator are linearly transformed to obtain the angle control value 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 the 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 quantity of each deflector is converted into a servo drive signal to obtain the PWM control signal of the deflector servo; the angle control quantity 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 ducted motor are collected and filtered in real time to obtain the actual ducted speed data. The sensor signals of the deflector servo and the control surface servo are also 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 angle control error, a preset control algorithm is used to adjust dynamic parameters and perform real-time correction on 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 for the ducted motor so that the duct of the aircraft can be controlled by the motor. The drive of the ducted motor is realized using a pulse width modulation (PWM) signal, the goal of which 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 used to drive the ducted motor. Similarly, for the position control variable of the deflector, it is converted into the PWM control signal of the servo. Assume that the position control variable 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 values ​​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] By calculating these PWM signals, the target angles of each rudder are converted into actual control signals, thereby driving the corresponding actuators to adjust the attitude. The Hall sensor signals and encoder signals of the ducted motor are collected in real time and filtered to obtain the actual speed data of the ducted motor. 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 with 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] Where LPF represents a low-pass filter, which is used to remove high-frequency noise from 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 control surface 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. To achieve closed-loop attitude control, a preset control algorithm is used to dynamically adjust parameters based on the speed control error and the angle control error, thereby correcting the control command in real time. The preset control algorithm corrects the error through proportional, integral, and differential actions. 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, representing the system's response strength to errors. Through a preset control algorithm, errors are corrected in real time to ensure that the ducted motor and rudder servo can quickly follow the target control variable, achieving stable attitude control of the aircraft. The output of the PID controller is superimposed on the original control signal to obtain the PWM control signal actually used for the actuator. For example, the final PWM control signal for the ducted motor is expressed as:

[0200]

[0201] The final PWM control signal of the rudder 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 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;

[0206] The difference calculation module 12 is used to compare the pitch angle data in the flight state parameters 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 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 an attitude deviation.

[0207] The power distribution calculation module 13 is used to calculate the power distribution for each of the three ducts based on the current flight mode and attitude deviation, and obtain the target rotational speed of each duct blade. If 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 deflector below the duct according to the current flight mode and attitude deviation, and obtain the position control value of each deflector;

[0209] A 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, yaw rudder, and elevator through vector synthesis calculation, and obtain the angle control value of each control surface;

[0210] The execution module 16 is used to convert the target 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 instructions to achieve closed-loop attitude control of the aircraft.

[0211] By integrating all of the aforementioned components, a three-duct layout combined with deflectors eliminates the tilt mechanism, simplifies the mechanical structure, and improves system reliability. The coordinated control of the three ducts also achieves more efficient power output. An innovative control strategy employing the three deflectors as triangular control points enables precise control of the aircraft's horizontal attitude and reduces control system complexity. A fault-tolerant control scheme based on power redundancy automatically distributes power to the remaining functioning ducts in the event of a single duct failure, ensuring continuous controllability of the aircraft. A complete closed-loop control system, utilizing real-time feedback data to correct control commands, improves attitude control accuracy and system robustness. Smooth transitions between vertical takeoff and landing and horizontal flight modes are achieved, ensuring flight stability through real-time pitch angle monitoring and mode determination. Multi-sensor data fusion and the extended Kalman filter algorithm improve state estimation accuracy, providing 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, which includes program instructions. When the program instructions are executed by the processor 301, the processor 301 may execute any of the above-mentioned aircraft attitude control methods.

[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 aircraft attitude control methods.

[0216] Those skilled in the art will understand that Figure 3 The structure shown in the figure is only a block diagram of a part of the structure related to the solution of the present application, and does not constitute a limitation on the electronic device 300 involved in the solution of the present application. 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), or 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. The general-purpose processor may be a microprocessor or any conventional processor.

[0218] It should be noted that those skilled in the art can clearly understand that, for the convenience and brevity 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 (SMC), a secure digital (SD) card, a flash memory card, etc., equipped with the electronic device 300.

[0221] Those skilled in the art will 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 for enabling an electronic device (which can be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: various media that can store program codes, such as a U disk, a mobile hard disk, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk.

[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 above embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the above 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 includes: Collect the aircraft's attitude angle data, acceleration data, angular velocity data, altitude data, speed data, and ducted speed data to obtain flight status parameters; Comparing the pitch angle data in the flight state parameters with a preset angle threshold to determine whether a flight mode switching condition is met, and when the switching condition is met and a mode switching instruction is received from the autopilot module or the external control module, executing the corresponding flight mode switching, and calculating 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 for each of the three ducts to obtain a target rotational speed for each ducted propeller blade. If any duct fails, the power demand of the duct is distributed to the remaining two ducts. performing position optimization calculations on the deflector plates below the duct according to the current flight mode and the attitude deviation, and obtaining position control values ​​for each deflector plate; Based on the current flight mode, the three deflectors are configured as triangular control points. The position compensation of the aileron, yaw rudder, and elevator is determined through vector synthesis calculation to obtain the angle control value of each control surface. The target rotation speed of each ducted blade, the position control quantity of each guide plate and the angle control quantity 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, wherein: 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 raw signal output by the inertial measurement unit in the aircraft is low-pass filtered to obtain the attitude angle data of the aircraft relative to the ground, including pitch, roll and yaw angles; Perform Kalman filtering calculation on the output signals of the acceleration sensor in the x-axis, y-axis, and z-axis directions to obtain acceleration data in the 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 the three orthogonal axes, and 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 used to calculate the speed data. The output signals of the Hall sensors of the three ducts are sampled and median filtered 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, wherein: The pitch angle data in the flight state parameters is compared 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 is received from an autopilot module or an 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 an attitude deviation, including: Comparing the pitch angle data in the flight state parameters with a first preset angle threshold to determine whether a switching condition to a vertical take-off and landing mode is satisfied, and comparing the pitch angle data in the flight state parameters with a second preset angle threshold to determine whether a switching condition to a horizontal flight mode is satisfied; 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, wherein: The power distribution calculation is performed on each of the three ducts based on the current flight mode and the attitude deviation to obtain a target rotational speed of each duct blade. When any duct fails, the power demand of the duct is distributed to the remaining two ducts, including: Performing a state analysis calculation on the current flight mode and the attitude deviation to obtain the overall power requirement in the Z-axis direction and the attitude torque requirements in the three orthogonal axes; Performing a power distribution calculation on the three ducts based on the overall power demand to obtain a reference power distribution coefficient for each duct; Performing a fault detection calculation on the deviation between the real-time speed data of the three ducts and the rated speed to obtain duct operating status identification information. The operating status of the three ducts is determined based on the duct operating status identification information. When a fault is detected in any duct, the power demand of the abnormal duct is proportionally distributed based on the duct operating status identification information to obtain the compensation power coefficients of the remaining normally operating ducts. Dynamically adjusting the power outputs of the three ducts based on the reference power distribution coefficient and the compensation power coefficient to obtain a target power output of each duct, and performing a ducted blade characteristic mapping operation on the target power output to obtain a corresponding ducted speed control variable; Based on the ducted speed control quantity, PID control calculation is performed on the torque of each ducted motor to obtain the motor control compensation quantity, and the ducted speed control quantity and the motor control compensation quantity are superimposed to obtain the target speed of each ducted blade.

5. The method for controlling the attitude of an aircraft according to claim 4, wherein: The position optimization calculation of the guide plates below the duct is performed according to the current flight mode and the attitude deviation to obtain the position control value of each guide plate, including: Based on the current flight mode, the attitude deviation is decomposed and calculated along the three axes of U, V, and W to obtain the desired correction forces in the three directions. The desired correction forces are then used to perform duct airflow guidance calculations to obtain the downswirling airflow direction angle data for the three ducts. Based on the current flight mode, dynamically calculate the interaction area between the deflector below the duct and the downswirling airflow to obtain an airflow action coefficient. Perform aerodynamic analysis on the airflow action coefficient and the downswirling airflow direction angle data to obtain an initial angle value for the deflector. 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; performing attitude decoupling calculations on initial angle values ​​of the deflectors according to the current flight mode and the aerodynamic force characteristic parameters to obtain independent control variables of each deflector, and performing collaborative optimization calculations on the independent control variables of each deflector to obtain target angle control variables of the deflectors; Based on the target angle control amount, a servo control calculation is performed on the angular displacement of the deflector servo to obtain the position control amount of each deflector.

6. The method for controlling the attitude of an aircraft according to claim 5, wherein: 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, including: Perform triangular control plane construction calculation on the spatial position coordinates of the three deflectors, use the three deflectors as three independent control points on the control plane, and obtain plane attitude parameters. Based on the plane attitude parameters, perform three-axis vector decomposition calculation on the horizontal plane control requirements to obtain the roll axis moment component, the yaw axis moment component, and the 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 yaw rudder surface response characteristics of the yaw axis moment component to obtain the initial yaw rudder deflection; According to the current flight mode, analyzing the elevator surface response characteristics of the pitch axis moment component to obtain the initial elevator deflection; performing a 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 value of each rudder surface.

7. The method for controlling the attitude of an aircraft according to claim 6, wherein: The method converts the target rotation speed of each ducted blade, the position control value of each deflector, and the angle control value of each control surface into an execution signal, and obtains execution feedback data 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 and calculated to obtain a PWM control signal of the ducted motor; the position control amount of each deflector is converted into a servo drive signal and calculated to obtain a PWM control signal of the deflector servo; the angle control amount of each rudder surface is converted into a servo drive signal and calculated to obtain a PWM control signal of the aileron, yaw rudder and elevator servo; The Hall sensor signal and encoder signal of the ducted motor are collected and filtered in real time to obtain the actual ducted speed data. The sensor signals of the deflector servo and the control surface servo are also collected and filtered in real time to obtain the current angle data of the servo. Comparing the actual duct speed data with the target duct speed to obtain a speed control error, and comparing the current angle data of the steering gear with the target steering gear angle to obtain an angle control error; Based on the 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: A method for controlling the attitude of an aircraft according to any one of claims 1 to 7, wherein the attitude control device of the aircraft comprises: The acquisition module is used to collect the aircraft's attitude angle data, acceleration data, angular velocity data, altitude data, speed data and duct speed data to obtain flight status parameters; a difference calculation module, configured 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 satisfied; and when the switching condition is satisfied and a mode switching instruction is received from the autopilot module or the external control module, execute the corresponding flight mode switching and calculate the difference between the current attitude and the target attitude to obtain an attitude deviation; a power distribution calculation module, configured to perform power distribution calculations on each of the three ducts based on the current flight mode and the attitude deviation, to obtain a target rotational speed for each ducted propeller blade, and to distribute the power demand of any duct to the remaining two ducts if a duct failure occurs; a position optimization calculation module, configured to perform position optimization calculation on the guide plates below the duct according to the current flight mode and the attitude deviation, and obtain a position control value for each guide plate; The 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 value of each control surface; The execution module is used to convert the target rotation speed of each ducted blade, the position control quantity of each guide plate and the angle control quantity of each control surface into an execution signal, and obtain the execution feedback data to perform real-time correction of the control instructions to achieve 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 for an aircraft according to any one of claims 1 to 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 according to any one of claims 1 to 7 is implemented.

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