CN110531777B - Four-rotor aircraft attitude control method and system based on active disturbance rejection control technology - Google Patents
Four-rotor aircraft attitude control method and system based on active disturbance rejection control technology Download PDFInfo
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Abstract
The invention discloses a method and a system for controlling the attitude of a four-rotor aircraft based on an active disturbance rejection control technology, which can effectively improve the robustness of attitude control of the four-rotor aircraft. The method adopts active disturbance rejection control to realize angular velocity loop control in attitude control; in an active disturbance rejection controller, a tracking differentiator is adopted to carry out smooth noise reduction processing on a given angular velocity signal to obtain a given angular velocity and an angular acceleration; the observer adopts a second-order extended state observer, and adopts the feedback angular velocity of the controlled object and the control quantity u (t-tau) added with the motor response delay to observe and estimate the sum of each order state of the angular velocity control model, internal disturbance and external disturbance acting on the model and unmodeled dynamics of the system; and calculating by utilizing the difference between the output quantity and the given quantity of the observer to obtain a primary control quantity through a nonlinear control rate, and compensating on the basis of the primary control quantity by utilizing the disturbance estimation quantity obtained by the observer to obtain a final output control quantity u (t).
Description
Technical Field
The invention belongs to the technical field of control of a four-rotor aircraft, and relates to a method and a system for controlling the attitude of the four-rotor aircraft based on an active disturbance rejection control technology.
Background
The four-rotor wing is an aircraft capable of taking off and landing Vertically (VTOL), belongs to a non-coaxial disc aircraft in the overall layout, has a more compact structure and generates more lift force compared with the conventional rotor wing aircraft, and two pairs of rotor wings with opposite rotation directions can mutually offset the reaction torque, so that the reaction torque paddle is not needed. Compared with a fixed-wing aircraft, the quad-rotor unmanned aerial vehicle can take off and land vertically and hover freely, has strong maneuvering capability and is particularly suitable for executing tasks under complex conditions. With the wide application of the quad-rotor unmanned aerial vehicle in the civil and military fields in recent years, the productivity level is improved, and meanwhile, the fighting mode of modern war is changed.
A quad-rotor aircraft has three translational degrees of freedom and three rotational degrees of freedom, for a total of six degrees of freedom, but only four control inputs. A quad-rotor aircraft is therefore a typical under-actuated system. Meanwhile, the four-rotor aircraft is a static unstable system with characteristics of strong coupling, nonlinearity, multivariable and the like, and can be interfered by external environments such as various physical effects and wind disturbance in low-altitude flight, so that the control effect is influenced and even instability is caused, and therefore a stable and reliable position control algorithm needs to be designed to ensure a series of requirements such as stability and precision of control under the condition of disturbance outside.
The active disturbance rejection control technology is developed by Korean Jingqing researchers and leading research groups thereof in the institute of mathematics and system science of Chinese academy of sciences, inherits and develops the concept of classical control, and absorbs the thought of modern control theory. The active disturbance rejection means: the unmodeled dynamic disturbance and the unknown external disturbance are both attributed to the unknown disturbance of the object, the input and output data are used for estimation and compensation, thereby realizing the dynamic feedback linearization of the dynamic system, and then the nonlinear configuration is used for forming a nonlinear feedback control law to improve the control performance of the closed-loop system.
Active disturbance rejection control technology is applied to the relevant control of the four-rotor aircraft by relevant scholars at present, but the effect of the active disturbance rejection control technology is yet to be further enhanced.
Disclosure of Invention
In view of the above, the invention provides a method and a system for controlling a posture of a quad-rotor aircraft based on an active disturbance rejection control technology, aiming at the defect of poor disturbance rejection capability in the posture control of the existing quad-rotor aircraft, and the method and the system can effectively improve the disturbance rejection capability and the tracking precision of the whole quad-rotor aircraft, so that the robustness of the posture control of the quad-rotor aircraft can be effectively improved.
In order to solve the technical problem, the invention is realized as follows:
a four-rotor aircraft attitude control method based on an active disturbance rejection control technology adopts an active disturbance rejection angular speed controller to realize angular speed loop control in attitude control; in the auto-disturbance-rejection angular velocity controller, a tracking differentiator is adopted to carry out smooth noise reduction processing on a given angular velocity signal v, and the input signal v subjected to smooth noise reduction processing is used1And its derivative v2As given angular velocity and angular acceleration; the observer adopts a second-order extended state observer, and adopts the feedback angular velocity y of a controlled object and a control quantity u (t-tau) added with the motor response delay tau to observe and estimate the sum of each order state of the angular velocity control model, internal disturbance and external disturbance acting on the model and unmodeled dynamic state of the system; estimating the angular velocity z output by an observer1With a given amount v of angular velocity1Making a difference, and estimating the angular velocityQuantity z1Differentiating to obtain angular acceleration estimation quantity and angular acceleration given quantity v2Making a difference, and calculating the two difference values through a nonlinear control rate to obtain a primary control quantity u0Using the disturbance estimator obtained by the observer to estimate the initial control quantity u0And compensating to obtain the final control quantity output u (t).
Preferably, the attitude control adopts a double-ring control structure, the inner ring angular velocity ring is realized by adopting the active disturbance rejection angular velocity controller, and the outer ring angular velocity ring is realized by adopting a PD angle controller.
Preferably, the method further comprises: establishing a PID attitude controller; adopting a PID attitude controller to control a manual takeoff process, starting to simultaneously operate the PID attitude controller and an active disturbance rejection attitude controller consisting of an active disturbance rejection angular velocity controller and a PD angle controller after the aircraft enters an automatic mode and is stably hovered, and still adopting the output quantity of the PID attitude controller as an actual output control quantity at the moment; and judging whether the difference between the control quantities calculated by the active disturbance rejection attitude controller and the PID attitude controller is within a certain range, and if the output quantity of the active disturbance rejection attitude controller is not dispersed and the difference between the control quantity values output by the two controllers is within a set range, switching to the control of the active disturbance rejection attitude controller.
Preferably, the tracking differentiator is constructed by adopting a steepest synthesis function fhan; in the steepest synthesis function, the fast factor r is 104Magnitude data, filter factor h1By using 10-2Magnitude data.
The invention also discloses a four-rotor aircraft attitude control system based on the active disturbance rejection control technology, which comprises an active disturbance rejection attitude controller, wherein the active disturbance rejection attitude controller comprises an inner ring angular velocity controller and an outer ring angular velocity controller; the angular velocity controller is an active disturbance rejection angular velocity controller and comprises a tracking differentiator, an observer, a nonlinear control rate calculation module, a time delay module and a control quantity compensation module;
a tracking differentiator for performing smoothing noise reduction processing on the given angular velocity signal v, and smoothing the input signal v subjected to the noise reduction processing1And its derivative v2As given angular velocity and angular acceleration for subsequent calculations;
the time delay module is used for adding the output control quantity u (t) into the motor response time delay tau to form u (t-tau) which is input into the observer;
the observer adopts a second-order extended state observer, and adopts the feedback angular velocity y of the controlled object and the control quantity u (t-tau) added with time delay to observe and estimate the state of each order of the angular velocity control model and the sum of internal and external disturbance acting on the model and unmodeled dynamics of the system to obtain the angular velocity estimation quantity z1And a disturbance estimator, and an angular velocity estimator z1Differentiating to obtain angular acceleration estimator
A non-linear control rate calculation module for estimating the angular velocity1And angular acceleration estimator
With a given quantity v from the tracking differentiator1And v2Correspondingly making difference, and calculating the two difference values through the nonlinear control rate to obtain a primary control quantity u0；
A control quantity compensation module for utilizing the disturbance estimation quantity obtained by the observer to perform the control on the preliminary control quantity u0And compensating to obtain the final control quantity output u (t).
Preferably, the system further comprises a PID attitude controller, a switching control module and a switch;
the switching control module is used for controlling the switching switch to enable the PID attitude controller to work in a manual takeoff stage; after the aircraft enters an automatic mode and is stably suspended, the PID attitude controller and the active disturbance rejection attitude controller work simultaneously by controlling the selector switch, but the output quantity of the PID attitude controller is still adopted as the actual output control quantity at the moment; and judging whether the difference between the control quantities calculated by the PID attitude controller and the active disturbance rejection attitude controller is within a certain range, and if the output quantity of the active disturbance rejection attitude controller is not dispersed and the difference between the two control quantity values is within a set range, switching to the active disturbance rejection attitude controller to work by controlling a selector switch.
Preferably, the tracking differentiator is constructed by adopting a steepest synthesis function fhan; in the steepest synthesis function, the fast factor r is 104Magnitude data, filter factor h1By using 10-2Magnitude data.
Compared with the prior art, the invention has the advantages and positive effects that:
(1) in actual flight, the attitude channel of a quad-rotor aircraft is susceptible to various disturbance factors and becomes unstable. According to the method, the uncertain part of the real-time estimation position model of the extended state observer and the sum of internal disturbance and external disturbance are added into the angular velocity control loop, the response delay time of the system is further considered in the design of ESO (electronic stability and engineering optimization), the obtained estimated value is more accurate, and finally the estimated value is compensated on the output control quantity u, so that the robustness of attitude control of the four-rotor aircraft can be effectively improved.
(2) For fast factor r and filter factor h in active disturbance rejection controller1The value range is limited, the parameter selection is to meet the minimum delay as far as possible, and the filtering requirement can be properly reduced, so that the limitation is to ensure that the addition of a differentiator cannot cause overlarge delay in the whole control process and avoid oscillation and even instability of a control system.
(3) The preferred embodiment of the invention adopts a scheme of an active disturbance rejection attitude controller and a PID attitude controller, the two controllers are used in stages, the PID control scheme is adopted in the takeoff stage, the problem that an observer is not easy to converge in the takeoff stage is solved, and the active disturbance rejection control scheme is adopted in the normal flight stage, so that the disturbance rejection capability and the tracking precision of the flight of the aircraft are improved. And by designing a reasonable switching control strategy, the stable transition of the switching process of the PID attitude controller to the active disturbance rejection attitude controller is ensured.
Drawings
FIG. 1 is a composite quad-rotor aircraft attitude control model based on active disturbance rejection + PID in an embodiment of the invention;
FIG. 2 is a schematic diagram of the auto-disturbance rejection attitude controller of FIG. 1;
fig. 3 is a block diagram of the auto-disturbance rejection angular velocity controller of fig. 2;
FIG. 4 is a schematic view of an embodiment of an active-disturbance-rejection based attitude control system for a quad-rotor aircraft;
FIG. 5 is a diagram illustrating the tracking of an increased disturbance angle (roll axis) at a fixed position in actual flight;
FIG. 6 shows the tracking of the increased disturbance angle (pitch axis) at the actual in-flight position fix;
FIG. 7 is a diagram showing the output of the disturbance variable (roll axis) increased under the actual fixed position in flight;
FIG. 8 is a view showing an output condition (pitch axis) of an increased disturbance control amount in the case of an actual in-flight position fix point;
FIG. 9 shows the situation (in x and y axes) that the position of the composite controller is adopted to track the circular track in the actual flight attitude control;
fig. 10 shows the situation that the actual in-flight attitude control adopts the position of the composite controller to track the circular track.
Detailed Description
The invention provides a four-rotor aircraft attitude control scheme based on active disturbance rejection control technology, and because a dynamics model of the attitude of the four-rotor unmanned aircraft is a multivariable, strong coupling and nonlinear second-order model, a double-ring controller is designed for the convenience of controlling and adjusting parameters, namely an angle controller of an outer ring and an angular speed controller of an inner ring are included to control the attitude of the four-rotor unmanned aircraft. Through analysis, the model uncertainty and disturbance of the four-rotor aircraft can be considered to be completely applied to the inner ring angular velocity ring, so that a general PD controller is still adopted for the outer ring angular velocity ring, and an active disturbance rejection angular velocity controller is designed for the inner ring angular velocity ring. The active disturbance rejection angular speed controller and the PD angle controller form an active disturbance rejection attitude controller.
For the design of an inner ring auto-disturbance rejection angular velocity controller, firstly, a tracking differentiator is adopted to carry out smooth noise reduction on a control quantity output by an outer ring angular velocity controller to be used as an inner ring angular velocity given value, derivation is carried out to obtain an inner ring angular acceleration given value, and then a second-order extended state observer is designed; the second-order extended state observer can estimate the angular velocity, and an angular acceleration estimated value can be obtained through differentiation; the second-order extended state observer can also estimate the total disturbance quantity borne by the system; and calculating the errors of the given values of the angular velocity and the angular acceleration and the corresponding estimated values through a nonlinear control rate to obtain a primary controlled variable, and adding the primary controlled variable into a disturbance estimated variable obtained by estimation of a second-order extended state observer to obtain final controlled variable output so as to realize excellent control on the system.
It can be seen that the greatest improvement point of the design of the auto-disturbance-rejection angular velocity controller is to improve the control quantity u (t) acting on the second-order extended state observer, and add the time delay representing the time from the generation of the control quantity to the feedback, so that all variables input to the observer are matched and aligned in time, and the disturbance of the system can be accurately estimated and compensated in real time.
In addition, because the attitude change of the quadrotor aircraft is large in the takeoff phase, the observer is easy to diverge to cause control instability, so that the PID attitude controller is adopted to control normal takeoff in the takeoff phase, and after the attitude of the quadrotor aircraft is stable, the control instability is switched to active disturbance rejection control. And the stable transition of the switching process of the PID attitude controller to the active disturbance rejection controller is ensured by designing a switching control strategy.
The invention is described in detail below by way of example with reference to the accompanying drawings.
Fig. 1 is a composite four-rotor aircraft attitude control structure based on active disturbance rejection + PID in an embodiment of the present invention, which includes an active disturbance rejection attitude controller and a PID attitude controller, which are switched by using a switch.
Fig. 2 is a detailed schematic diagram of the auto-disturbance-rejection attitude controller in fig. 1. The active disturbance rejection attitude controller is a dual-ring controller, namely an outer ring angle controller and an inner ring angular velocity controller, for controlling the attitude of the active disturbance rejection attitude controller. Through analysis, model uncertainty and disturbance of the four-rotor unmanned aerial vehicle can be considered to be completely applied to the inner ring angular velocity ring, so that the outer ring angular velocity ring still adopts a general PD angle controller, and only the inner ring angular velocity ring is provided with an active disturbance rejection angular velocity controller.
Fig. 3 is a block diagram showing the detailed components of the active disturbance rejection angular velocity controller in fig. 2. Compared with the existing active disturbance rejection controller, the time delay link is increased.
The following describes the design and workflow of the present invention.
Firstly, establishing a four-rotor aircraft attitude dynamics model, wherein the dynamics model is described as follows:
wherein, four independent control input values, U2For roll axis control, U3For pitch axis control, U4The control quantity of the yaw axis phi, theta and psi respectively represent the roll and the pitch of the attitude of the body, and the Euler angle I of the yaw axis under an inertial coordinate systemxx,Iyy,IzzRespectively, the moment of inertia of the corresponding channel, JrRepresenting the moment of inertia associated with the gyroscopic effect and/representing the arm length from the centre of each motor to the centre of the body. It can be seen that the three channels have the characteristics of multivariable, nonlinearity and strong coupling, so that generally, corresponding active disturbance rejection controllers are respectively designed for the three channels to compensate the mutual coupling effect between each channel and the internal and external disturbances.
The control model for each axis is described as follows:
wherein x is1,x2Representing the state variables of the system model, i.e. the values of the angle and angular velocity of the corresponding channel, f (x), respectively1,x2W (t), t) represents the uncertain part of the system model and the sum of internal and external disturbances suffered by the system, bu (t-tau) represents the part of the output control quantity, which takes the response delay of the actuating mechanism into consideration and acts on the system model, wherein tau is the response delay time of the system, and b is the estimated value of the control quantity amplification factor in the system model. For the control of each channel, the main objective of the invention is to estimate f (x) as accurately and quickly as possible1,x2W (t), t) and b, so that the uncertain and disturbed parts of the model can be accurately and quickly compensated on the basis of the PD angle controller finally.
And secondly, designing a tracking differentiator in the attitude control of the active disturbance rejection control technology according to the established model.
The tracking differentiator is in the form of:
fh＝fhan(v1(k)-v(k),v2(k),r,h1)
v1(k+1)＝v1(k)+h2v2(k)
v2(k+1)＝v2(k)+h2fh
wherein, the input signal v (k) is the input discrete signal v1(k) And v2(k) Tracking the input signal itself and its derivative, v, respectively1(k) For the tracked angular velocity information, v2(k) Angular acceleration information for tracking is v1(k) The derivative of (c). fhan is the fastest synthesis function, and its expression is described in document 1 (active disturbance rejection control technology, han jingqing, leading edge science 20071, total phase 1). r, h1Respectively representing fast and filter factors, h2Representing the step factor. The parameters are selected to meet the requirement of minimum delay as much as possible, the filtering requirement can be properly reduced, and the method can be used as long as no step or peak signal with extremely large amplitude appears, so that the fast factor r of the invention adopts 104Magnitude data, filter factor h1By using 10-2Magnitude data.
And thirdly, designing a second-order Extended State Observer (ESO) according to a position control model of the four-rotor aircraft, and considering the response delay time of the added system, so as to estimate the disturbance applied to the inner ring angular velocity ring in real time and compensate in real time.
The invention designs a second-order extended state observer for observing each-order state of a system, simultaneously estimates internal and external disturbances borne by the actual system, and compensates in real time so as to enhance the robustness of the system.
The specifically designed second-order extended state observer is shown as follows:
e＝z1-y
where y is the feedback value of the actual system, i.e., the value of the angular velocity returned by each channel. z is a radical of1Representing the intrinsic state of the system, i.e. the angular velocity value of each channel; z is a radical of2I.e. variables representing the expansion state of the system, including the uncertainty of the system model and the sum of the total internal and external disturbances, z2The value multiplied by 1/b is used to compensate to the output control quantity u (t). bu represents a control amount applied to the model, b is an estimated value of an amplification factor of the control amount in the system model, u is an output control amount of the present control scheme, and bu (t- τ) represents a control amount input to the system model in consideration of a motor response delay since a feedback value y of the system has passed through the systemThe time delay is considered, so that all signals input into the observer are synchronized, and the obtained estimated value is more accurate. Where τ represents the time delay of action of the control quantity.
Essentially the observer remains an error-driven system, β01,β02Two observer gains are associated with the sampling step size of the system. Theoretically, the larger the gain of the observer is, the better the gain is, the observation speed can be accelerated, so that real-time compensation of system uncertainty and disturbance is realized, and an ideal control effect is achieved. The ideal effect can be obtained by adjusting parameters through experiments.
Fourthly, obtaining angular velocity estimation quantity z according to the obtained four-rotor attitude control model and the observer1The given amount v of angular velocity obtained by tracking the differentiator1An error e can be obtained1And the differential e of the error2。
e1＝v1-z1
Fifthly, calculating the error e of the angular velocity and the angular acceleration1、e2Calculating to obtain a preliminary control quantity u through a nonlinear control rate0。
For non-linear control rates, the present invention takes the form of:
u0＝b1fal(e1,α1,1)+b2fal(e2,α2,2)
wherein b is1,b2For the control gain, fal (e, α,) is a non-linear function, whose specific expression is shown below:
the nonlinear controller theoretically has small error and large gain; large error and small gain, so that the performance of the original controller can be slightly improved by adding a small amount of nonlinear part, therefore, alpha is selected to be a reasonable value near 1.
Sixthly, according to the basic principle of active disturbance rejection control, the system control rate after internal and external disturbance compensation is obtained as shown in the following formula:
wherein,
the method is a part for compensating the total disturbance value estimated by the second-order extended state observer into the control quantity. b is the estimated value of the control quantity amplification factor in the system model, and u is the actual output of the control scheme.
When the flight control device is used, aiming at the control model shown in the figure 1, a reasonably designed controller switching strategy can realize the stable transition between the PID attitude controller adopted by manual flight and the auto-disturbance rejection attitude controller adopted by automatic flight. The basic idea of the switching control strategy is as follows: after the PID attitude controller is manually taken off and enters an automatic mode to be suspended stably basically, the PID attitude controller and the active disturbance rejection attitude controller are started to operate simultaneously, but the output quantity of the PID attitude controller is still adopted as the actual output control quantity at the moment; at the moment, whether the difference between the calculated control quantities is within a certain range is judged, if the output quantity of the active disturbance rejection attitude controller is not dispersed and the numerical difference between the two control quantities is within a certain range, the controller is switched to work in the active disturbance rejection attitude controller once the controller is switched and larger fluctuation of the output of the control quantities can not appear theoretically; if the auto-disturbance rejection attitude controller diverges or the difference between the two control quantities is large, switching cannot be carried out until a switching condition is met.
Based on the above method process, the invention also provides a four-rotor aircraft attitude control system based on the active disturbance rejection control technology, as shown in fig. 1, comprising an active disturbance rejection attitude controller, a PID attitude controller, a switching control module and a switch.
The switching control module enables the PID attitude controller to work by controlling the switching switch in a manual takeoff stage; after the aircraft enters an automatic mode and is stably suspended, the PID attitude controller and the active disturbance rejection attitude controller work simultaneously by controlling the selector switch, but the output quantity of the PID attitude controller is still adopted as the actual output control quantity at the moment; and judging whether the difference between the two control quantities calculated by the PID attitude controller and the active disturbance rejection attitude controller is within a certain range, and if the output quantity of the controller based on the active disturbance rejection is not dispersed and the difference between the control quantity numerical values calculated by the two controllers is within a certain range, switching to the operation of the active disturbance rejection attitude controller by controlling a change-over switch.
The active disturbance rejection attitude controller adopts a double-ring control structure and comprises an inner ring angular velocity ring and an outer ring angular ring; the inner ring angular velocity ring is realized by adopting an auto-disturbance rejection angular velocity controller, and the outer ring angular velocity ring is realized by adopting a PD (potential difference) angular velocity controller. As shown in fig. 4, the active disturbance rejection angular velocity controller includes a tracking differentiator, an observer, a nonlinear control rate calculation module, a time delay module, and a control amount compensation module.
A tracking differentiator for performing smoothing noise reduction processing on the given angular velocity signal v, and smoothing the input signal v subjected to the noise reduction processing1And its derivative v2As given angular velocity and angular acceleration for subsequent calculations;
the time delay module is used for adding the output control quantity u (t) into the motor response time delay tau to form u (t-tau) which is input into the observer;
the observer adopts a second-order extended state observer, and adopts the feedback angular velocity y of the controlled object and the control quantity u (t-tau) added with time delay to observe and estimate the state of each order of the angular velocity control model and the sum of internal and external disturbance acting on the model and unmodeled dynamics of the system to obtain the angular velocity estimation quantity z1And a disturbance estimator, and an angular velocity estimator z1Differentiating to obtain angular acceleration estimator
A non-linear control rate calculation module for estimating the angular velocity1And angular acceleration estimator
With a given quantity v from the tracking differentiator1And v2Correspondingly making difference, and calculating the two difference values through the nonlinear control rate to obtain a primary control quantity u0；
A control quantity compensation module for utilizing the disturbance estimation quantity obtained by the observer to perform the control on the preliminary control quantity u0And compensating to obtain the final control quantity output u (t).
In order to verify the effectiveness of the designed active disturbance rejection plus PID-based composite control technology four-rotor attitude controller, the invention utilizes the QAV-250 frame-based four-rotor aircraft to carry out actual flight experiments.
In the experiment, the four-rotor aircraft is developed mainly based on ANO-Pioneer open source flight control and is built by utilizing a QAV-250 frame. The flight control is based on an STM32F407VG main controller, the highest main frequency can reach 168MHz, an ICM20602 module is adopted as an inertia measurement module, an EMAX MT2204KV2300 direct-current brushless motor, an EMAX sink 12A electronic speed regulator and 5045 blades are adopted as a power system. The total mass of the body in flight is about 650g and the load mass to which the disturbance is applied is about 150 g.
In order to verify the excellent anti-interference performance of the inner ring attitude controller in an actual flight experiment, a certain degree of disturbance is applied to an actual four-rotor aircraft under the condition of a fixed position, the angular response conditions of two channels of attitude rolling and attitude pitching are respectively observed, as shown in fig. 5 and 6, and the control quantity of the two corresponding channels is analyzed, as shown in fig. 7 and 8.
During the whole flight, the 0s-15s aircraft is in a hovering state without obvious disturbance, a load is hung on one arm of the four-rotor aircraft at 15s, disturbance in the horizontal direction with different sizes is applied to the suspended load during 15s-40s, the suspended load is removed at 40s, and the aircraft is still in the hovering state for the rest of time. As can be seen from FIGS. 5 and 6, the angular errors of the roll and pitch channels of the aircraft in the hovering state before 15s are basically kept within + -1deg considering the influence of the sensor error, which indicates that the control accuracy of the active disturbance rejection controller in the hovering state is high. When a larger disturbance which changes more gently is applied in 15s, the angle errors of the two channels can be still kept within +/-3 deg under the conditions that the load is not subjected to violent disturbance which changes in the horizontal direction excessively and the measurement error of the sensor is considered, even if the violent disturbance is applied to the horizontal direction of the suspended load in about 33s, the angle can be still in a controllable state, the error is in an acceptable range, the control delay of the whole process is lower, the performance of the observer is better, the rapid and accurate estimation and compensation of the larger and more violent disturbance can be realized, and the requirement of certain disturbance resistance is met.
Meanwhile, as can be seen from fig. 7 and 8, in the hovering state under the normal condition, the output of the two-channel controller is substantially zero, after the load is suspended for 15s, the observer quickly estimates the disturbance, which is reflected in the controlled variable that the controlled variable rapidly and smoothly reaches a stable value, when severe disturbance is applied for 33s, the controlled variable output slightly oscillates but still converges rapidly, and at this time, the output of the controlled variable is close to the saturation state, it can be seen that the controller still has a stable performance under the extreme state, and the expected requirements on the control accuracy and the robustness are met.
In the aspect of position control, as shown in fig. 9 and 10, the attitude controller based on the active disturbance rejection technology is used as a bottom controller, a nonlinear position controller is adopted in an outer ring to track a circular track, and it can be seen that the control accuracy in the x and y axial directions is controlled to be + -5cm, and basically no tracking control delay exists, the overall effect is good, the accuracy is comparable to or even superior to that of a common PID attitude controller, but the disturbance rejection capability is far superior to that of the common PID attitude controller.
In summary, the above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (4)
1. A four-rotor aircraft attitude control method based on an active disturbance rejection control technology is characterized in that attitude control adopts a double-ring control structure, an inner ring angular velocity ring is realized by adopting an active disturbance rejection angular velocity controller, and an outer ring angular velocity ring is realized by adopting a PD angular velocity controller;
in the auto-disturbance-rejection angular velocity controller, a tracking differentiator is adopted to carry out smooth noise reduction processing on a given angular velocity signal v, and the input signal v subjected to smooth noise reduction processing is used1And its derivative v2As given angular velocity and angular acceleration; the observer adopts a second-order extended state observer, and adopts the feedback angular velocity y of a controlled object and a control quantity u (t-tau) added with the motor response delay tau to observe and estimate the sum of each order state of the angular velocity control model, internal disturbance and external disturbance acting on the model and unmodeled dynamic state of the system; estimating the angular velocity z output by an observer1With a given amount v of angular velocity1Making a difference, and estimating the angular velocity by using the angular velocity estimation quantity z1Differentiating to obtain angular acceleration estimation quantity and angular acceleration given quantity v2Making a difference, and calculating the two difference values through a nonlinear control rate to obtain a primary control quantity u0Using the disturbance estimator obtained by the observer to estimate the initial control quantity u0Compensating on the basis to obtain final control quantity output u (t);
establishing a PID attitude controller; adopting a PID attitude controller to control a manual takeoff process, starting to simultaneously operate the PID attitude controller and an active disturbance rejection attitude controller consisting of an active disturbance rejection angular velocity controller and a PD angle controller after the aircraft enters an automatic mode and is stably hovered, and still adopting the output quantity of the PID attitude controller as an actual output control quantity at the moment; and judging whether the difference between the control quantities calculated by the active disturbance rejection attitude controller and the PID attitude controller is within a certain range, and if the output quantity of the active disturbance rejection attitude controller is not dispersed and the difference between the control quantity values output by the two controllers is within a set range, switching to the control of the active disturbance rejection attitude controller.
2. The method of claim 1, wherein the tracking differentiator is constructed using a steepest synthesis function fhan; in the steepest synthesis function, the fast factor r is 104Magnitude data, filter factor h1By using 10-2Magnitude data.
3. A four-rotor aircraft attitude control system based on active disturbance rejection control technology comprises an active disturbance rejection attitude controller, and is characterized in that the active disturbance rejection attitude controller comprises an inner ring angular velocity controller and an outer ring angular velocity controller; the angular velocity controller is an active disturbance rejection angular velocity controller and comprises a tracking differentiator, an observer, a nonlinear control rate calculation module, a time delay module and a control quantity compensation module; the system further comprises a PID attitude controller, a switching control module and a switch;
a tracking differentiator for performing smoothing noise reduction processing on the given angular velocity signal v, and smoothing the input signal v subjected to the noise reduction processing1And its derivative v2As given angular velocity and angular acceleration for subsequent calculations;
the time delay module is used for adding the output control quantity u (t) into the motor response time delay tau to form u (t-tau) which is input into the observer;
the observer adopts a second-order extended state observer, and adopts the feedback angular velocity y of the controlled object and the control quantity u (t-tau) added with time delay to observe and estimate the state of each order of the angular velocity control model and the sum of internal and external disturbance acting on the model and unmodeled dynamics of the system to obtain the angular velocity estimation quantity z1And a disturbance estimator, and an angular velocity estimator z1Differentiating to obtain angular acceleration estimator
A non-linear control rate calculation module for estimating the angular velocity1And angular acceleration estimator
With a given quantity v from the tracking differentiator1And v2Correspondingly making difference, and calculating the two difference values through the nonlinear control rate to obtain a primary control quantity u0；
A control quantity compensation module for utilizing the disturbance estimation quantity obtained by the observer to perform the control on the preliminary control quantity u0Compensating on the basis to obtain final control quantity output u (t);
the switching control module is used for controlling the switching switch to enable the PID attitude controller to work in a manual takeoff stage; after the aircraft enters an automatic mode and is stably suspended, the PID attitude controller and the active disturbance rejection attitude controller work simultaneously by controlling the selector switch, but the output quantity of the PID attitude controller is still adopted as the actual output control quantity at the moment; and judging whether the difference between the control quantities calculated by the PID attitude controller and the active disturbance rejection attitude controller is within a certain range, and if the output quantity of the active disturbance rejection attitude controller is not dispersed and the difference between the two control quantity values is within a set range, switching to the active disturbance rejection attitude controller to work by controlling a selector switch.
4. The system of claim 3, wherein the tracking differentiator is constructed using a steepest synthesis function fhan; in the steepest synthesis function, the fast factor r is 104Magnitude data, filter factor h1By using 10-2Magnitude data.
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