Ornithopter attitude control method

Abstract

The application discloses a flapping-wing aircraft attitude control method, characterized in that the flapping-wing aircraft comprises a main control board, a roll attitude controller, a pitch attitude controller, an attitude estimator and an actuator connected with the main control board, the pitch attitude controller comprises an angle angular velocity controller and a low-pass filter angle controller, and the method comprises the following steps: in the take-off and landing stage of the aircraft, the output of the angle angular velocity controller is completely adopted as a final control instruction; and in the flight stage, a double-channel combined control mode is adopted. The application adopts the double-channel switching in the pitch attitude control strategy, mainly adopts the angle control of the low-pass filter pitch attitude angle in the stable level flight state, and improves the control efficiency and stability. The roll angle and the pitch angle are decoupled and controlled as a whole.

Description

Technical Field

[0001] This invention relates to an aircraft, and more particularly to an attitude control method for an ornithopter, belonging to the field of ornithopter technology. Background Technology

[0002] A flapping-wing aerial vehicle (FWAV) is a biomimetic aircraft that mimics the flight of birds or insects, generating lift and thrust through the active flapping of its wings. Its core principle is based on aerodynamics: during wing flapping, the downward flapping phase generates upward lift and forward thrust, while the upward flapping phase reduces drag, creating the force required for sustained flight.

[0003] Currently, most flapping-wing aircraft use a single main gear to drive symmetrical flapping of both wings. This flapping motion generates a large periodic pitch moment, causing severe oscillations in the pitch attitude. This affects real-time Kalman filtering of pitch attitude estimation, leading to an inability to accurately reflect the overall flight state of the aircraft. Relevant literature can be found in Chinese Utility Model Patent No. ZL 202321606077.4, entitled "Installation Structure of Avionics System for Flapping-Wing Aircraft" (Authorization Announcement No. CN 220595200U).

[0004] The inherent dynamic characteristics of flapping-wing aircraft mean that the high-frequency pitch attitude oscillations caused by flapping during flight cannot be resolved by horizontal stabilizer control. Therefore, cascaded PID pitch attitude angle control based on real-time attitude calculation is unreasonable. It cannot alleviate the pitch attitude oscillation problem during flight through horizontal stabilizer control, and it also causes oscillations in the horizontal stabilizer's own control angle, wasting energy and increasing flight drag. Furthermore, due to the coupling between the horizontal and vertical stabilizers, an inappropriate pitch attitude control strategy can also affect control performance in other dimensions. Therefore, it is necessary to improve the attitude control of flapping-wing aircraft. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to provide a flapping-wing aircraft attitude control method that can alleviate the high-frequency pitch attitude angle vibration caused by the symmetrical flapping of the two wings during flight, in view of the above-mentioned technical status.

[0006] The technical solution adopted by the present invention to solve the above-mentioned technical problems is as follows: a flapping-wing aircraft attitude control method, characterized in that the flapping-wing aircraft includes a main control board and a roll attitude controller, a pitch attitude controller, an attitude estimator and an actuator connected to the main control board, wherein the aforementioned pitch attitude controller includes an angle and angular velocity controller and a low-pass filter angle controller, and includes the following steps:

[0007] The aforementioned roll attitude controller obtains the roll angle difference based on the roll angle setpoint and the current roll angle value. This roll angle difference is divided by a preset time constant to obtain the expected roll angle angular velocity. Then, the expected roll angle angular velocity in the inertial frame is converted to the machine system through the Jacobian matrix to obtain the target value of the front axle angular velocity of the machine system. Based on the target value of the front axle angular velocity of the machine system and the actual front axle angular velocity value of the machine system, the front axle angular velocity difference is obtained. The final control output is generated by superimposing the three factors, and the result is output to the actuator to control the channel.

[0008] The aforementioned angle and angular velocity controller obtains the first pitch angle difference based on the current pitch angle value and the first pitch angle setpoint. This first pitch angle difference is divided by a preset time constant to obtain the first pitch angular velocity setpoint. Then, using the Jacobian matrix, the first pitch angular velocity setpoint in the inertial frame is converted to the machine system to obtain the target value of the right axis angular velocity of the first machine system. Based on the difference between the target value of the right axis angular velocity of the first machine system and the actual right axis angular velocity value of the machine system, the right axis angular velocity difference of the first machine system is obtained. The final control output is generated by superimposing three factors, and the result is output to the actuator to control this channel.

[0009] The aforementioned low-pass filter angle controller filters the original three-axis Euler angle attitude of the aircraft to obtain the filtered pitch angle value. Based on the preset second pitch angle setpoint and the pitch angle filtered value, the second pitch angle difference is obtained. This second pitch angle difference is divided by a preset time constant to obtain the second pitch angle angular velocity setpoint. The aforementioned second pitch angle angular velocity setpoint is transformed into the machine system coordinate system through a Jacobian matrix to obtain the target value of the second machine system right axis angular velocity. This target value of the second machine system right axis angular velocity is the second machine system right axis angular velocity difference. The final control output is generated by superimposing three factors, and the result is output to the actuator for control of this channel.

[0010] The aforementioned angle and angular velocity controller and low-pass filter angle controller achieve dynamic switching through the following safety strategy: during the takeoff and landing phases of the aircraft, the output of the angle and angular velocity controller is used as the final control command; during the flight phase, a dual-channel joint control mode is adopted.

[0011] The aforementioned three factors include: ① the product of the system angular velocity setpoint and the feedforward coefficient and scaling factor; ② the product of the system shaft angular velocity difference and the square of the proportional coefficient and scaling factor; ③ the updated integral term.

[0012] The aforementioned current roll angle value, front axle angular velocity value of the actual aircraft system, current pitch angle value, and right axle angular velocity value of the actual aircraft system are all obtained through the attitude estimator.

[0013] Preferably, the desired roll angle angular velocity is obtained by the following steps to obtain the target value of the front axle angular velocity of the engine system:

[0014] The expected value of the roll angle angular velocity in the inertial coordinate system is subtracted from the product of the sine of the current pitch angle and the target yaw angular velocity in the inertial coordinate system, and then the airframe angular velocity feedforward compensation is added.

[0015] As a preferred option, the integral term in roll angle control is subjected to anti-saturation treatment, and the steps are as follows:

[0016] If the output value is less than -1.0, the integral accumulation term is restricted to a value not less than 0; if the output value of the previous cycle is greater than 1.0, the integral accumulation term is restricted to a value not greater than 0. After completing the anti-saturation adjustment, the product of the adjusted integral accumulation term and the integral coefficient is added to the integral term of the previous cycle to obtain the updated integral term.

[0017] Preferably, the control output of the roll attitude controller is processed by the mixer before driving the actuator.

[0018] Preferably, the first pitch angular velocity setting value is obtained by obtaining the target value of the right axis angular velocity of the first machine system through the following steps:

[0019] Multiply the first pitch angular velocity setting value in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw angular velocity setting value in the inertial coordinate system, and then add the airframe angular velocity feedforward compensation.

[0020] Preferably, the second pitch angular velocity setpoint is obtained by the following steps to obtain the target value of the right axis angular velocity of the second machine system:

[0021] Multiply the second pitch angular velocity setting value in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw angular velocity setting value in the inertial coordinate system, and then add the airframe angular velocity feedforward compensation.

[0022] Preferably, the dynamic switching between the angle velocity controller and the low-pass filter angle controller includes the following steps:

[0023] During takeoff, the system records the takeoff point altitude Hhome in real time. When the current altitude Hhome of the aircraft is detected... current≥Hhome+20m, and after continuously meeting this height condition for a 1s delay detection time, the system switches to dual-channel joint control mode. By introducing the control quantity ratio coefficient α (α∈[0,1]), the weighted fusion of the outputs of the two types of controllers is achieved. The formula for calculating the joint control quantity Utotal is: Utotal=α×Uangle+(1-α)×Urate, where Uangle is the output of the low-pass filter angle controller and Urate is the output of the angular velocity controller.

[0024] If the throttle value is detected to drop below the set throttle threshold during flight, regardless of whether the current altitude is within the safe range, the system will immediately switch back to the full angular velocity controller to respond in real time to any attitude disturbances that may occur during the descent. Once the throttle value is detected to rise above the safe range above the throttle threshold, a delay condition of 1 second must be met before switching back to the dual-channel joint control mode.

[0025] Compared with existing technologies, the advantages of this invention are as follows: Angle and angular velocity controllers are more sensitive and can quickly achieve control effects (however, they do not consider the dynamic characteristics of flapping-wing aircraft, resulting in greater drag and energy consumption); low-pass filter angle controllers have relatively smooth control responses, but their control effects are significantly delayed. This invention addresses these shortcomings by employing a dual-channel switching strategy for pitch attitude control. In stable level flight, it primarily uses low-pass filter pitch attitude angle control, improving control efficiency and stability. Overall, it employs decoupled control of roll and pitch angles. The attitude control algorithm during flight employs a switching strategy to optimize flight control modes during takeoff, landing, and stable flight, improving safety during takeoff and landing and achieving energy savings during level flight. Attached Figure Description

[0026] Figure 1 This is a schematic diagram of the structure of an embodiment.

[0027] Figure 2 This is a scaled-down schematic diagram from another perspective of the embodiment.

[0028] Figure 3 This is a control principle diagram for an example. Detailed Implementation

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

[0030] like Figure 1 and Figure 2 As shown, Figure 1 and Figure 2As shown, the flapping-wing aircraft in this embodiment includes a frame 4 and a flight control module 5, a wing 1, a front servo 11, a gear set 12, a tail 3, and a rear servo 21 mounted on the frame 4. The power output end of the front servo 11 is connected to the gear set 12, and the output end of the gear set 12 is connected to the front connecting rod 13. After the front servo 11 is started, it can drive the gear set 12 to rotate, which in turn drives the front connecting rod 13 to flap the wing 1 up and down.

[0031] The flight control module 5 has a built-in main controller, roll attitude controller, pitch attitude controller and attitude estimator.

[0032] The tail fin 3 includes a vertical stabilizer 32 and a horizontal stabilizer 31, which are orthogonally integrated. The horizontal stabilizer 31 is used to control the pitch moment and adjust the pitch angle, while the vertical stabilizer 32 is used to control the yaw moment and control the yaw angle and roll angle. The tail fin 3 is connected to the rear end of the frame 4 via a rotating component 24.

[0033] The power output end of the rear servo 21 is connected to the rear linkage 22, which is connected to the tail fin 3 via an L-shaped connector. In this embodiment, the rear servo 21 and the rear linkage 22 are a pair, symmetrically arranged. The rear servo 21 drives the tail fin 3 to control its vertical angle via the rear linkage 22. The rotating component 24 is used to control the left-right angle of the tail fin.

[0034] In this embodiment, when the two rear servos 21 output in the same direction, the rear linkage 22 controls in opposite directions, which means that the tail fin 3 rotates to the left / right and the flapping-wing aircraft turns. When the two rear servos 21 output in opposite directions, the rear linkage 22 controls in the same direction, which means that the tail fin 3 rotates to the up / down direction and controls the pitch direction of the flapping-wing aircraft to tilt up or down.

[0035] In this embodiment, the rear servo 21 and the rear linkage 22 constitute the actuator of the tail wing 3.

[0036] In this embodiment, due to the symmetrical flapping of the flapping wing, the flapping wing will inevitably cause pitch attitude oscillations due to the pitch moment in one cycle, which manifests as periodic fluctuations in the pitch angle. Moreover, such fluctuations are difficult to resolve by the tail control.

[0037] Combination Figure 3 As shown, the attitude control method for a flapping-wing aircraft includes a main control board and a roll attitude controller, a pitch attitude controller, an attitude estimator, and an actuator connected to the main control board. The pitch attitude controller includes an angle and angular velocity controller and a low-pass filter angle controller, and includes the following steps:

[0038] The roll attitude controller obtains the roll angle difference based on the roll angle setpoint and the current roll angle value. This roll angle difference is divided by a preset time constant to obtain the expected roll angle angular velocity. Then, using the Jacobian matrix, the expected roll angle angular velocity in the inertial frame is converted to the aircraft system to obtain the target value of the aircraft system's front axle angular velocity. Based on the target value and the actual value of the aircraft system's front axle angular velocity, the difference in front axle angular velocity is obtained. The final control output is generated by superimposing these three factors, and the result is output to the actuator for control of this channel. Specifically, the control output of the roll attitude controller is processed by the mixer before driving the actuator. After parameter tuning, it has good performance in roll angle control, which can give the flapping-wing aircraft good maneuverability.

[0039] The angle and angular velocity controller obtains the first pitch angle difference based on the current pitch angle value and the first pitch angle setpoint. This first pitch angle difference is divided by a preset time constant to obtain the first pitch angular velocity setpoint. Then, using the Jacobian matrix, the first pitch angular velocity setpoint in the inertial frame is converted to the machine system to obtain the target value of the right axis angular velocity of the first machine system. Based on the difference between the target value of the right axis angular velocity of the first machine system and the actual right axis angular velocity value of the machine system, the right axis angular velocity difference of the first machine system is obtained. The final control output is generated by superimposing three factors, and the result is output to the actuator to control this channel.

[0040] The low-pass filter angle controller filters the original three-axis Euler angle attitude of the aircraft to obtain the filtered pitch angle value. Based on the preset second pitch angle setpoint and the pitch angle filtered value, the second pitch angle difference is obtained. This second pitch angle difference is divided by a preset time constant to obtain the second pitch angle angular velocity setpoint. The second pitch angle angular velocity setpoint is transformed into the machine system coordinate system through a Jacobian matrix to obtain the target value of the second machine system right axis angular velocity. This target value of the second machine system right axis angular velocity is the second machine system right axis angular velocity difference. The final control output is generated by superimposing the three factors, and the result is output to the actuator to control this channel.

[0041] The angle and angular velocity controller and the low-pass filter angle controller achieve dynamic switching through the following safety strategy: during the takeoff and landing phases of the aircraft, the output of the angle and angular velocity controller is used as the final control command; during the flight phase, a dual-channel joint control mode is adopted.

[0042] The three factors include: ① the product of the system angular velocity setpoint and the feedforward coefficient and scaling coefficient; ② the product of the system axis angular velocity difference and the square of the proportional coefficient and scaling coefficient; ③ the updated integral term; the system angular velocity setpoint is the front axis angular velocity difference for the roll attitude controller and the right axis angular velocity difference for the angle angular velocity controller.

[0043] The current roll angle, the front axle angular velocity of the actual aircraft system, the current pitch angle, and the right axle angular velocity of the actual aircraft system are all obtained through the attitude estimator.

[0044] The expected value of the roll angle velocity is obtained by the following steps to obtain the target value of the front axle angular velocity of the airframe: subtract the product of the sine of the current pitch angle and the target yaw rate in the inertial coordinate system from the expected value of the roll angle velocity in the inertial coordinate system, and then add the airframe angular velocity feedforward compensation.

[0045] The steps for anti-saturation treatment of the integral term in roll angle control are as follows:

[0046] If the output value is less than -1.0, the integral accumulation term is restricted to a value not less than 0; if the output value of the previous cycle is greater than 1.0, the integral accumulation term is restricted to a value not greater than 0. After completing the anti-saturation adjustment, the product of the adjusted integral accumulation term and the integral coefficient is added to the integral term of the previous cycle to obtain the updated integral term.

[0047] The first pitch rate setpoint is obtained by the following steps to obtain the target value of the right axis angular velocity of the first aircraft system: multiply the first pitch rate setpoint in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw rate setpoint in the inertial frame, and then add the airframe angular velocity feedforward compensation.

[0048] The integral term of the angle and angular velocity controller undergoes anti-saturation processing. To avoid control saturation caused by the accumulation of integral terms, the integral term needs to be adjusted for anti-saturation first. Specifically, if the output value of the control channel in the previous cycle is less than -1.0, the integral accumulation term is limited to a value not less than 0; if the output value in the previous cycle is greater than 1.0, the integral accumulation term is limited to a value not greater than 0. After completing the anti-saturation adjustment, the product of the adjusted integral accumulation term and the integral coefficient is added to the integral term of the previous cycle to obtain the updated integral term.

[0049] The second pitch rate setpoint is obtained by the following steps to obtain the target value of the right axis angular velocity of the second aircraft system: multiply the second pitch rate setpoint in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw rate setpoint in the inertial frame, and then add the airframe angular velocity feedforward compensation.

[0050] The low-pass attitude filter in this embodiment is a second-order low-pass filter with a cutoff frequency of 2Hz. It filters the original three-axis Euler angle attitude of the aircraft to obtain the filtered Euler angles of the aircraft. For flapping-wing aircraft in level flight and hovering state with flapping frequency above 5Hz, the low-pass filter with this setting performs well in terms of delay and filtering effect.

[0051] The integral term of the low-pass attitude filter is subjected to anti-saturation processing. If the output value of the control channel in the previous cycle is less than -1.0, the integral accumulation term is restricted to a value not less than 0; if the output value in the previous cycle is greater than 1.0, the integral accumulation term is restricted to a value not greater than 0. After the anti-saturation adjustment is completed, the product of the adjusted integral accumulation term and the integral coefficient is added to the integral term of the previous cycle to obtain the updated integral term.

[0052] The dynamic switching between the angle and angular velocity controller and the low-pass filter angle controller includes the following steps:

[0053] During takeoff, the system records the takeoff point altitude Hhome in real time. When the current altitude Hhome of the aircraft is detected... current ≥Hhome+20m, and after continuously meeting this height condition for a 1s delay detection time, the system switches to dual-channel joint control mode. By introducing the control quantity ratio coefficient α (α∈[0,1]), the weighted fusion of the outputs of the two types of controllers is achieved. The formula for calculating the joint control quantity Utotal is: Utotal=α×Uangle+(1-α)×Urate, where Uangle is the output of the low-pass filter angle controller and Urate is the output of the angular velocity controller.

[0054] If the throttle value is detected to drop below the set throttle threshold during flight, regardless of whether the current altitude is within the safe range, the system will immediately switch back to the full angular velocity controller to respond in real time to any attitude disturbances that may occur during the descent. Once the throttle value is detected to rise above the safe range above the throttle threshold, a delay condition of 1 second must be met before switching back to the dual-channel joint control mode.

[0055] This embodiment proposes an attitude control algorithm for coupled pitch and roll attitude control. The roll angle and pitch angle each follow a separate control channel. It proposes a switching strategy for the attitude (pitch) control algorithm during the entire flight process, and optimizes the flight control logic during takeoff, landing and stable flight. The aim is to optimize the safety during takeoff and landing and achieve energy saving optimization during level flight.

[0056] For flapping-wing aircraft with single-segment wing and single-motor symmetrical control, there is an inherent pitch oscillation period. A dual-channel switching strategy is adopted for pitch attitude control. In stable level flight, low-pass filtering is mainly used for pitch attitude angle control to improve control efficiency and stability. Overall, roll and pitch attitude decoupling control is employed.

Claims

1. A method for attitude control of an ornithopter, characterized in that... The flapping-wing aircraft includes a main control board and a roll attitude controller, a pitch attitude controller, an attitude estimator, and an actuator connected to the main control board. The aforementioned pitch attitude controller includes an angle and angular velocity controller and a low-pass filter angle controller, and includes the following steps: The aforementioned roll attitude controller obtains the roll angle difference based on the roll angle setpoint and the current roll angle value. This roll angle difference is divided by a preset time constant to obtain the expected roll angle angular velocity. Then, through the Jacobian matrix, the expected roll angle angular velocity in the inertial frame is converted to the machine system to obtain the target value of the front axle angular velocity of the machine system. Based on the target value of the front axle angular velocity of the machine system and the actual front axle angular velocity value of the machine system, the front axle angular velocity difference is obtained. The final control output is generated by superimposing three factors, and the result is output to the actuator for control of the roll attitude controller channel. The aforementioned angle and angular velocity controller obtains the first pitch angle difference based on the current pitch angle value and the first pitch angle setpoint. This first pitch angle difference is divided by a preset time constant to obtain the first pitch angular velocity setpoint. Then, using the Jacobian matrix, the first pitch angular velocity setpoint in the inertial frame is converted to the machine system to obtain the target value of the right axis angular velocity of the first machine system. Based on the difference between the target value of the right axis angular velocity of the first machine system and the actual right axis angular velocity value of the machine system, the right axis angular velocity difference of the first machine system is obtained. The final control output is generated by superimposing three factors, and the result is output to the actuator for angle and angular velocity control channel control. The aforementioned low-pass filter angle controller filters the original three-axis Euler angle attitude of the aircraft to obtain the filtered pitch angle value. Based on the preset second pitch angle setpoint and the pitch angle filtered value, the second pitch angle difference is obtained. The second pitch angle difference is divided by the preset time constant to obtain the second pitch angle angular velocity setpoint. The aforementioned second pitch angle angular velocity setpoint is transformed into the machine system coordinate system through the Jacobian matrix to obtain the target value of the second machine system right axis angular velocity. The target value of the second machine system right axis angular velocity is the second machine system right axis angular velocity difference. The final control output is generated by superimposing the three factors, and the result is output to the actuator for control of the low-pass filter angle controller channel. The aforementioned angle and angular velocity controller and low-pass filter angle controller achieve dynamic switching through the following safety strategy: during the takeoff and landing phases of the aircraft, the output of the angle and angular velocity controller is used as the final control command; during the flight phase, a dual-channel joint control mode is adopted. The aforementioned three factors include The product of the system angular velocity setpoint and the feedforward coefficient and scaling coefficient; The product of the difference in the angular velocity of the machine system shaft and the square of the proportional coefficient and the scaling factor; Updated points system; The setpoint for the engine system angular velocity is the target value for the front axle angular velocity of the engine system in the roll attitude controller; The setpoint for the machine system angular velocity is the target value of the right axis angular velocity of the first machine system in the angle and angular velocity controller. The setpoint for the machine system angular velocity in the low-pass filter angle controller is the target value of the right axis angular velocity of the second machine system. The aforementioned current roll angle value, front axle angular velocity value of the actual aircraft system, current pitch angle value, and right axle angular velocity value of the actual aircraft system are all obtained through the attitude estimator.

2. The attitude control method for an flapping-wing aircraft according to claim 1, characterized in that... The desired roll rate is obtained by the following steps to obtain the target front axle angular velocity of the engine system: The expected value of the roll angle angular velocity in the inertial coordinate system is subtracted from the product of the sine of the current pitch angle and the target yaw angular velocity in the inertial coordinate system, and then the airframe angular velocity feedforward compensation is added.

3. The attitude control method for an flapping-wing aircraft according to claim 1, characterized in that... The steps for anti-saturation treatment of the integral term in roll angle control are as follows: If the output value is less than -1.0, the integral accumulation term is restricted to a value not less than 0; if the output value of the previous cycle is greater than 1.0, the integral accumulation term is restricted to a value not greater than 0. After completing the anti-saturation adjustment, the product of the adjusted integral accumulation term and the integral coefficient is added to the integral term of the previous cycle to obtain the updated integral term.

4. The attitude control method for an ornithopter according to claim 1, characterized in that... The control output of the roll attitude controller is processed by the mixer before driving the actuator.

5. The attitude control method for an flapping-wing aircraft according to claim 1, characterized in that... The first pitch angular velocity setpoint is obtained by the following steps to obtain the target value of the right axis angular velocity of the first machine system: Multiply the first pitch angular velocity setting value in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw angular velocity setting value in the inertial coordinate system, and then add the airframe angular velocity feedforward compensation.

6. The attitude control method for an flapping-wing aircraft according to claim 1, characterized in that... The second pitch angular velocity setpoint is obtained by the following steps to obtain the target value of the right axis angular velocity of the second machine system: Multiply the second pitch angular velocity setting value in the inertial coordinate system by the cosine of the current roll angle, add the product of the cosine of the current pitch angle, the sine of the current roll angle, and the calculated yaw angular velocity setting value in the inertial coordinate system, and then add the airframe angular velocity feedforward compensation.

7. The attitude control method for an ornithopter according to claim 1, characterized in that... The dynamic switching between the angle velocity controller and the low-pass filter angle controller includes the following steps: During takeoff, the system records the takeoff point altitude Hhome in real time. When the current altitude of the aircraft Hcurrent ≥ Hhome + 20m is detected and this altitude condition is met for a delay of 1 second, the system switches to the dual-channel joint control mode. By introducing the control quantity ratio coefficient α (α∈[0,1]), the weighted fusion of the outputs of the two types of controllers is achieved. The formula for calculating the joint control quantity Utotal is: Utotal = α×Uangle + (1-α)×Urate, where Uangle is the output of the low-pass filter angle controller and Urate is the output of the angular velocity controller. If the throttle value is detected to drop below the set throttle threshold during flight, regardless of whether the current altitude is within the safe range, the system will immediately switch back to the full angular velocity controller to respond in real time to any attitude disturbances that may occur during the descent. After the throttle value is subsequently detected to rise above the safe range above the throttle threshold, a delay condition of 1 second for the safe throttle state must be met before switching back to the dual-channel joint control mode.

Images