A high-precision altitude control system based on aerodynamics

Abstract

The present invention discloses a high-precision altitude control system based on aerodynamics, which belongs to the field of unmanned aerial vehicle control systems, including an altitude control system and an aircraft having the altitude control system, wherein the aircraft includes an aircraft body and an altitude controller, an actuator, and a sensor element arranged on the aircraft body, wherein the altitude control system is integrated on the altitude controller; the altitude controller includes an altitude control loop and a pitch control loop connected in series with the altitude control loop, wherein the altitude control loop is used to process an altitude error signal and use it as an input control signal of the actuator, and during the process, the altitude controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control. The present invention ensures accurate altitude flight through a PID controller, thereby achieving stable, efficient flight, and high safety. The control system based on the PID controller has the advantage of strong robustness.

Description

Technical Field

[0001] The invention relates to an unmanned aerial vehicle control system, in particular to a high-precision altitude control system based on aerodynamic force. Background Art

[0002] Fixed-wing UAVs have advantages such as high flight speed and long range. At present, the longitudinal altitude control system of fixed-wing UAVs is mainly based on PID controller to control the elevator, and longitudinal altitude control is achieved by controlling the pitch of the UAV. This control system can achieve altitude control without static error in the absence of disturbance. However, in the presence of external disturbances, the longitudinal altitude control accuracy of this control system is not high. Therefore, in the actual flight of fixed-wing UAVs, the altitude fluctuation in the fixed altitude flight mode reaches 1-2m. In addition, since this control system controls the longitudinal altitude by controlling the pitch of the aircraft, the pitch attitude of the aircraft fluctuates greatly when maintaining the flight altitude. For high-precision altitude control, some people use the L1 adaptive control method. The simulation results show that the error is small and can meet the accuracy requirements. However, it is very difficult to select the optimal low-pass filter in the feedback loop. Others construct a nonlinear disturbance observer, observe the external disturbance and the error of the system, and design a controller based on this to improve the control accuracy, but the control has a jitter problem and poor robustness.

[0003] It can be seen that the shortcomings of the existing aircraft altitude control system include that the flight altitude control is not accurate and stable enough, the flight attitude of the aircraft is not stable enough when controlling the flight altitude, and the robustness is poor, resulting in low flight safety and efficiency. Therefore, those skilled in the art provide a high-precision altitude control system based on aerodynamics to solve the problems raised in the above background technology. Summary of the invention

[0004] The purpose of the present invention is to provide a high-precision altitude control system based on aerodynamics, which ensures precise altitude flight through a PID controller to achieve stable, efficient flight and high safety. The control system based on the PID controller has the advantage of strong robustness to solve the problems raised in the above background technology.

[0005] To achieve the above object, the present invention provides the following technical solutions:

[0006] A high-precision altitude control system based on aerodynamics comprises an altitude control system and an aircraft having the altitude control system, wherein the aircraft comprises an aircraft body and an altitude controller, an actuator and a sensor element arranged on the aircraft body, wherein the altitude control system is integrated on the altitude controller; the altitude controller comprises an altitude control loop and a pitch control loop connected in series with the altitude control loop, wherein the altitude control loop is used to process an altitude error signal and then use it as an input control signal of the actuator, during which the altitude controller converts the actual output signal through the sensor element and then compares it with the input signal to form a closed-loop control; the output signal of the altitude control loop is used as a reference input signal of the pitch control loop, and is compared with a signal from the sensor element to obtain an error signal, and the error signal is processed by the pitch control loop and becomes an input control signal of the actuator; the actuator is used to receive the control signal transmitted from the altitude controller, and adjust the altitude and pitch attitude of the aircraft according to the control signal; the sensor element is used to measure the flight status information of the aircraft and convert the flight status information into a usable output signal.

[0007] As a further solution of the present invention: the actuator specifically includes: an elevator, a flap and a motor integrated on the aircraft.

[0008] As a further solution of the present invention: the altitude control loop is a P+PID controller, that is, a proportional+proportional-integral-differential controller, which is used to control the flaps and the motor to achieve altitude control of the aircraft.

[0009] As a further solution of the present invention: the specific control process of the height control loop is:

[0010] Step 1: Generate height error signal e h , e h =h c -h, where h c is the height control input signal, h is the height signal from the sensor element;

[0011] Step 2: e h After being processed by the height proportional controller, it becomes the height change rate control signal w c , w c =P h ×e h , where P h is the height proportional gain of the proportional controller;

[0012] Step 3: w c Compared with the height change rate signal w from the sensor element, a height change rate error signal e is generated w , e w =w c -w;

[0013] Step 4: After the speed proportional-integral-differential controller is processed, the coefficient K T , K F , K e After processing, they become the input signals of the aircraft's flap, throttle and pitch control loops. Δθ C ,in, Among them, P w is the height change rate proportional gain of the proportional-integral-derivative controller, I w is the height change rate integral gain of the proportional-integral-derivative controller, D w is the rate-of-change derivative gain of the proportional-integral-derivative controller.

[0014] As a further solution of the present invention: the pitch control loop is a P+PI controller, that is, a proportional+proportional-integral controller, which is used to control the elevator in series with the altitude control loop. The specific control process is:

[0015] S1: Convert the pitch control input signal Δθ from the altitude control loop C Compared with the pitch angle signal θ from the sensor element, a pitch angle error signal e is generated θ , e θ =Δθ C -θ;

[0016] S2:e θ After being processed by the proportional controller of the pitch control loop, it becomes the pitch angle rate control signal Δq C , Δq C =P θ ×e θ , where P θ is the pitch angle proportional gain of the proportional controller;

[0017] S3: Δq C Compared with the angular rate signal q from the sensor element, an angular rate error signal e is generated q , e q =Δq C -q;

[0018] S4: After being processed by the proportional-integral controller of the angular rate, it becomes the control signal of the elevator Among them, P q is the angular rate proportional gain of the proportional-integral controller, I q is the angular rate integral gain of the proportional-integral controller.

[0019] As a further solution of the present invention: the aircraft is subjected to external disturbance signals during flight, and the disturbance signals include but are not limited to gusts.

[0020] As a further solution of the present invention: the sensor element includes but is not limited to an accelerometer, a gyroscope, a barometer, and a distance measuring sensor.

[0021] The present application also discloses a high-precision altitude control method based on aerodynamic force, which is applied to a high-precision altitude control system based on aerodynamic force. The control method comprises the following steps:

[0022] (1): The sensor element measures the flight status information of the aircraft and converts the flight status information into a usable output signal;

[0023] (2): The altitude control loop processes the altitude error signal and uses it as the input control signal of the actuator;

[0024] (3): The output signal of the altitude control loop is used as the reference input signal of the pitch control loop, and an error signal is obtained after being compared with the signal from the sensor element;

[0025] (4): The pitch control loop processes the error signal in step 3 and uses it as the input control signal of the actuator;

[0026] (5): The actuator receives the control signal and adjusts the altitude and pitch attitude of the aircraft according to the control signal;

[0027] (6): During the flight of the aircraft, the altitude controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control.

[0028] Compared with the prior art, the present invention has the following beneficial effects:

[0029] 1. The present invention combines the advantages of pitch control altitude and flap control altitude, ensuring accurate control of flight altitude while greatly reducing the pitch attitude fluctuation of the aircraft when maintaining the flight altitude, thereby improving the safety and stability of the aircraft.

[0030] 2. The altitude control loop works in conjunction with the pitch control loop to effectively control the flaps, motor-propellers, and elevators and achieve precise altitude control of the aircraft.

[0031] 3. The PID controller is used to ensure accurate altitude flight, thereby achieving stable and efficient flight. The control system based on the PID controller greatly improves the robustness of altitude control.

[0032] 4. The control system based on PID controller greatly improves flight reliability and safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 It is a control flow diagram of a high-precision height control system based on aerodynamic force;

[0034] Figure 2 This is a control flow chart of a height controller in a high-precision height control system based on aerodynamic force;

[0035] Figure 3 Schematic diagram of a precise altitude control algorithm in a high-precision altitude control system based on aerodynamic forces. DETAILED DESCRIPTION

[0036] In order to better understand the above technical solution, the above technical solution will be described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0037] See also Figures 1 to 3 In an embodiment of the present invention, a high-precision altitude control system based on aerodynamic force includes an altitude control system and an aircraft having the altitude control system, wherein the aircraft includes an aircraft body and an altitude controller, an actuator, and a sensor element arranged on the aircraft body, wherein the altitude control system is integrated on the altitude controller; the altitude controller includes an altitude control loop and a pitch control loop connected in series with the altitude control loop, wherein the altitude control loop is used to process an altitude error signal and use it as an input control signal of the actuator, during which the altitude controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control; the output signal of the altitude control loop is used as a reference input signal of the pitch control loop, and is compared with the signal from the sensor element to obtain an error signal, which is processed by the pitch control loop and becomes the input control signal of the actuator; the actuator is used to receive the control signal from the altitude controller, and adjust the altitude and pitch attitude of the aircraft according to the control signal; the sensor element is used to measure the flight status information of the aircraft and convert the flight status information into a usable output signal. The present application ensures precise altitude flight through a PID controller, thereby achieving stable, efficient flight and high safety. The control system based on the PID controller has the advantage of strong robustness.

[0038] In this embodiment, the actuator specifically includes: an elevator, a flap, and a motor integrated on the aircraft. The elevator adopts a V-tail elevator.

[0039] In this embodiment: the altitude control loop is a P+PID controller, that is, a proportional+proportional-integral-differential controller, which is used to control the flaps and the motor to achieve altitude control of the aircraft. The motor is connected to the propeller of the aircraft, and controlling the speed of the motor can control the thrust of the propeller. It should be noted that in order to avoid the altitude control loop being unable to respond in time due to the integral saturation of the PID controller, it is proposed to use an integral clamp to deal with the impact of the integral saturation of the PID controller. Specifically, the PID controller is an anti-saturation controller, and it is determined whether the PID controller meets the integral saturation condition; if so, the integral term is clamped by the integral clamp, wherein this embodiment provides an implementation method for determining whether the PID controller meets the integral saturation condition, the altitude change rate parameter value of the aircraft is input into the saturation function corresponding to the PID controller, and a control parameter value is obtained; it is determined whether the control parameter value is equal to the altitude change rate parameter value of the aircraft, if so, the PID controller is not saturated; if not, the PID controller is saturated.

[0040] In this embodiment: Figure 3 As shown in the figure, the specific control process of the height control loop is:

[0041] Step 1: Generate height error signal e h , e h =h c -h, where h c is the height control input signal, h is the height signal from the sensor element;

[0042] Step 2: e h After being processed by the height proportional controller, it becomes the height change rate control signal w c , w c =P h ×e h , where P h is the height proportional gain of the proportional controller;

[0043] Step 3: w c Compared with the height change rate signal w from the sensor element, a height change rate error signal e is generated w , e w =w c -w;

[0044] Step 4: After the speed proportional-integral-differential controller is processed, the coefficient K T , K F , K e After processing, they become the input signals of the aircraft's flap, throttle and pitch control loops. Δθ C ,in, Among them, P w is the height change rate proportional gain of the proportional-integral-derivative controller, I w is the height change rate integral gain of the proportional-integral-derivative controller, D w is the differential gain of the height change rate of the proportional-integral-differential controller. It should be noted that the running state of the motor can be controlled by controlling the throttle of the motor.

[0045] In this embodiment: Figure 3 As shown, the pitch control loop is a P+PI controller, i.e., a proportional + proportional-integral controller, which is used to control the elevator in series with the altitude control loop. The specific control process is:

[0046] S1: Convert the pitch control input signal Δθ from the altitude control loop C Compared with the pitch angle signal θ from the sensor element, a pitch angle error signal e is generated θ , e θ =Δθ C -θ;

[0047] S2:e θ After being processed by the proportional controller of the pitch control loop, it becomes the pitch angle rate control signal Δq C , Δq C =P θ ×e θ , where P θ is the pitch angle proportional gain of the proportional controller;

[0048] S3: Δq C Compared with the angular rate signal q from the sensor element, an angular rate error signal e is generated q , e q =Δq C -q;

[0049] S4: After being processed by the proportional-integral controller of the angular rate, it becomes the control signal of the elevator Among them, P q is the angular rate proportional gain of the proportional-integral controller, I q is the angular rate integral gain of the proportional-integral controller.

[0050] In this embodiment, the aircraft is subjected to external disturbance signals during flight, and the disturbance signals include but are not limited to gusts. The presence of gusts affects the accuracy of the aircraft's flight control.

[0051] In this embodiment, the sensor elements include but are not limited to accelerometers, gyroscopes, barometers, and ranging sensors. The accelerometer can measure the acceleration of the aircraft, the gyroscope can be used as a horizontal, vertical, pitch, heading, and angular velocity sensor, the barometer can measure the atmospheric pressure, and the ranging sensor can measure the height of the aircraft from the ground.

[0052] In this embodiment, a high-precision altitude control method based on aerodynamic force is also disclosed, which is applied to a high-precision altitude control system based on aerodynamic force. The control method includes the following steps:

[0053] (1): The sensor element measures the flight status information of the aircraft and converts the flight status information into a usable output signal;

[0054] (2): The altitude control loop processes the altitude error signal and uses it as the input control signal of the actuator;

[0055] (3): The output signal of the altitude control loop is used as the reference input signal of the pitch control loop, and an error signal is obtained after being compared with the signal from the sensor element;

[0056] (4): The pitch control loop processes the error signal in step 3 and uses it as the input control signal of the actuator;

[0057] (5): The actuator receives the control signal and adjusts the altitude and pitch attitude of the aircraft according to the control signal;

[0058] (6): During the flight of the aircraft, the altitude controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control.

[0059] The working principle of the present invention is: when used, Figure 1 As shown, the altitude control process is specifically as follows: when there is an error between the altitude signal (real flight altitude) from the sensor element and the altitude control input signal (predetermined flight altitude), an error signal will be generated. This signal is processed by the altitude controller and becomes the input signal of the actuator (the input signal is divided into three paths, one path is used as the reference input signal of the pitch control loop, and the other two paths are used as the input signals of the flaps and motor of the aircraft respectively), which controls the actuator to adjust the pitch attitude and altitude of the aircraft. During the entire control process, the aircraft will also receive external disturbance signals (such as gusts of wind, etc.). The actual output signal is converted by the sensor elements (accelerometer, gyroscope, barometer, ranging sensor, etc.) and compared with the input signal to form a closed-loop control, thereby achieving altitude control of the aircraft. As shown Figure 2As shown, the output signal of the altitude control loop is used as the reference input signal of the pitch control loop. After being compared with the signal from the sensor element, an error signal is obtained. After being processed by the pitch control loop, the signal becomes the input signal of the aircraft elevator to control the pitch attitude of the aircraft. At the same time, the output signal of the altitude control loop is used as the input signal of the aircraft flaps and motors to control the motor speed and the angle of the flap control surface of the aircraft. By simultaneously controlling the flaps of the aircraft, the control surface angle of the elevator, and the thrust of the propeller, precise altitude control is achieved. In general, the altitude control loop controls the flaps and motors of the aircraft, and the altitude control loop and the pitch control loop control the elevator of the aircraft in series. The precise altitude control system achieves precise control of the altitude of the aircraft by connecting the altitude control loop and the pitch control loop in series to control the pitch angle of the aircraft, and the altitude control loop and the pitch control loop in parallel to control the altitude of the aircraft.

[0060] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include these modifications and variations.

[0061] The above are only preferred specific implementation modes of the present invention, but the protection scope of the present invention is not limited thereto. Any technician familiar with the technical field can make equivalent replacements or changes according to the technical solutions and inventive concepts of the present invention within the technical scope disclosed by the present invention, which should be covered by the protection scope of the present invention.

Claims

1. A high-precision altitude control system based on aerodynamics, characterized in that: It includes an altitude control system and an aircraft having the altitude control system, wherein the aircraft includes an aircraft body and an altitude control controller, an actuator, and a sensor element arranged on the aircraft body, and the altitude control system is integrated on the altitude control controller; The height controller includes a height control loop and a pitch control loop connected in series with the height control loop. The height control loop is used to process the height error signal and use it as the input control signal of the actuator. During the process, the height controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control. The output signal of the height control loop is used as the reference input signal of the pitch control loop, and is compared with the signal from the sensor element to obtain an error signal. The error signal is processed by the pitch control loop and becomes the input control signal of the actuator. The actuator is used to receive the control signal from the altitude controller and adjust the altitude and pitch attitude of the aircraft according to the control signal; The sensor element is used to measure the flight status information of the aircraft and convert the flight status information into a usable output signal; The actuator specifically includes: elevators, flaps and motors integrated on the aircraft; The altitude control loop is a P+PID controller, i.e. a proportional+proportional-integral-differential controller, which is used to control the flaps and the motor to achieve altitude control of the aircraft; The specific control process of the height control loop is: Step 1: Generate height error signal e h , e h =h c -h, where h c is the height control input signal, h is the height signal from the sensor element; Step 2: e h After being processed by the height proportional controller, it becomes the height change rate control signal w c , w c =P h ×e h , where P h is the height proportional gain of the proportional controller; Step 3: w c Compared with the height change rate signal w from the sensor element, a height change rate error signal e is generated w , e w =w c -w; Step 4: After the speed proportional-integral-differential controller is processed, the coefficient K T , K F , K e After processing, they become the input signals of the aircraft's flap, throttle and pitch control loops. Δθ C ,in, Among them, P w is the height change rate proportional gain of the proportional-integral-derivative controller, I w is the height change rate integral gain of the proportional-integral-derivative controller, D w is the height change rate differential gain of the proportional-integral-derivative controller; The pitch control loop is a P+PI controller, i.e., a proportional + proportional-integral controller, which is used to control the elevator in series with the altitude control loop. The specific control process is: S1: Convert the pitch control input signal Δθ from the altitude control loop C Compared with the pitch angle signal θ from the sensor element, a pitch angle error signal e is generated θ , e θ =Δθ C -θ; S2:e θ After being processed by the proportional controller of the pitch control loop, it becomes the pitch angle rate control signal Δq C , Δq C =P θ ×e θ , where P θ is the pitch angle proportional gain of the proportional controller; S3: Δq C Compared with the angular rate signal q from the sensor element, an angular rate error signal e is generated q , e q =Δq C -q; S4: After being processed by the proportional-integral controller of the angular rate, it becomes the control signal of the elevator Among them, P q is the angular rate proportional gain of the proportional-integral controller, I q is the angular rate integral gain of the proportional-integral controller.

2. A high-precision altitude control system based on aerodynamic force according to claim 1, characterized in that: The aircraft is subjected to external disturbance signals during flight, and the disturbance signals include but are not limited to gusts.

3. The high-precision altitude control system based on aerodynamic force according to claim 1, characterized in that: The sensor elements include but are not limited to accelerometers, gyroscopes, barometers, and ranging sensors.

4. A high-precision altitude control method based on aerodynamic force, applied to the high-precision altitude control system based on aerodynamic force as claimed in any one of claims 1 to 3, characterized in that: The control method comprises the following steps: (1): The sensor element measures the flight status information of the aircraft and converts the flight status information into a usable output signal; (2): The altitude control loop processes the altitude error signal and uses it as the input control signal of the actuator; (3): The output signal of the altitude control loop is used as the reference input signal of the pitch control loop, and an error signal is obtained after being compared with the signal from the sensor element; (4): The pitch control loop processes the error signal in step 3 and uses it as the input control signal of the actuator; (5): The actuator receives the control signal and adjusts the altitude and pitch attitude of the aircraft according to the control signal; (6): During the flight of the aircraft, the altitude controller converts the actual output signal through the sensor element and compares it with the input signal to form a closed-loop control.

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