CN115509249A - Attitude control method of four-rotor aircraft in Mars gust strong interference environment - Google Patents
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Abstract
The invention discloses a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment, which comprises the steps of constructing an attitude dynamics model of the four-rotor aircraft, analyzing the internal characteristics of the four-rotor aircraft, designing a double-ring active disturbance rejection nonlinear controller on the basis, compensating external disturbance through an extended state observer of an inner ring and an outer ring, and realizing the accurate tracking of the four-rotor aircraft on a set expected attitude angle, thereby completing the control of the attitude of the four-rotor aircraft.
Description
Technical Field
The invention belongs to the technical field of unmanned aerial vehicles and active disturbance rejection control, and particularly relates to a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment.
Background
The multi-rotor unmanned aerial vehicle is a very flexible aircraft, and has wide application in aspects such as military reconnaissance, agricultural surveying and mapping, aerial photography and the like due to the special advantages of hovering and vertical takeoff and landing. How to effectively control the unmanned aerial vehicle has been a hot problem of research in the industrial and academic fields in the past decades.
The currently determined scientific target for Mars exploration mainly comprises the steps of determining whether life exists on a Mars, knowing the evolution process and history of Mars climate, knowing the origin and evolution of Mars as a geological system and the like. Even though people have been exploring mars and researching the universe for more than half a century, the detection level at present is far from the clear detection level of mars.
The characteristics of severe mars surface flight condition, violent flow field change and the like have severe requirements on the performance indexes of the unmanned aerial vehicle, and in order to ensure that the mars unmanned aerial vehicle completes the mars detection task, the technical problems of flight control and navigation of the unmanned aerial vehicle under the condition without a GPS (global positioning system) and system integration of the unmanned aerial vehicle need to be solved urgently. In rotor type mars unmanned aerial vehicle's flight control, thin mars atmosphere can lead to unmanned aerial vehicle rotor lift to be far less than at earth's environment's range of variation along with the range of rotor speed variation, therefore mars unmanned aerial vehicle's flight attitude adjustment process can be very slow. However, the phenomena of the mars wind, the dust storm and the like can seriously affect the flying stability of the mars unmanned aerial vehicle, and the mars unmanned aerial vehicle is required to be capable of making quick adjustment aiming at the change of the environment so as to ensure the flying safety of the unmanned aerial vehicle. At present, the control of the rotary wing type mars unmanned aerial vehicle mainly realizes the steering control and the attitude adjustment of the mars unmanned aerial vehicle by adjusting the relative position of a Tip Path Plane (TPP) and the mass center of the unmanned aerial vehicle.
Especially, four-rotor aircraft has received extensive attention from a plurality of trades because of its characteristics such as small, mechanical design is simple, flight simple operation. However, a quad-rotor aircraft is a nonlinear system with strong coupling and under-actuation characteristics, which has six degrees of freedom, but only four control inputs; meanwhile, due to the fact that the size and the weight of the wind-driven generator are small, external interference such as gust is caused to have large influence on the wind-driven generator. Therefore, designing a reliable control strategy for stable attitude of a four-rotor system is a challenging task.
Disclosure of Invention
In view of this, the invention provides a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment, which realizes the attitude control of the four-rotor aircraft in a simulated Mars gust strong interference environment.
The invention provides a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment, which comprises the following steps:
step 2, constructing an outer ring controller and an inner ring controller, wherein the outer ring controller comprises an outer ring nonlinear extended state observer and an outer ring nonlinear controller, and the inner ring controller comprises an inner ring nonlinear extended state observer and an inner ring nonlinear controller;
and 3, in actual control, the outer ring nonlinear extended state observer takes the angular speed expected value output by the outer ring nonlinear controller and the angle feedback value output by the four-rotor aircraft as input, and the output state estimated value is obtained through calculation
Outer loop nonlinear controller with desired angle of direct input, angle of quad-rotor aircraft output, and state estimate
For input, the expected value v of the angular velocity is calculated j As an output; the inner-loop nonlinear extended state observer takes an angular velocity feedback value output by the quadrotor aircraft and a control quantity output by the inner-loop nonlinear controller as inputs, and calculates to obtain an output state estimation value
The inner ring linear controller outputs angular speed expected value from the outer ring controller, angular speed feedback value from the four-rotor aircraft, and state estimation value from the inner ring nonlinear extended state observer
For input, calculating the control quantity u of the four-rotor aircraft to obtain output i (ii) a Will control the quantity u i And inputting the data into a posture dynamics model of the four-rotor aircraft to realize the control of the four-rotor aircraft.
Further, the outer ring nonlinear extended state observer in step 2 is constructed as follows:
wherein, y 1 ＝η＝[φ,θ,ψ] T ＝[y 1j ] 3×1 ∈R 3×1 J = phi, theta, psi andy 2 ＝[y 2j ] 3×1 ∈R 3×1 j = phi, theta, psi respectively represent external disturbance values, phi being the pitch angle of the quad-rotor craft, theta being the roll angle of the quad-rotor craft, psi being the yaw angle of the quad-rotor craft,
and
respectively represent y 1 And y 2 Corresponding state estimation value, v j Represents the output value of the outer loop non-linear controller,
as state estimates
The value of the first derivative of (a),
is a state estimation value
First derivative value of, beta 1j And ε 1j Are all positive adjustable constants which are constant in number,
the expression for the nonlinear function is:
where σ is a positive adjustable constant.
Further, the outer loop nonlinear controller is configured as follows:
wherein e is η ＝η-η d ＝[e ηj ] 3×1 ∈R 3×1 J = phi, theta, psi is the error value of the pitch angle, the roll angle and the yaw angle; k is a radical of 0j ,k 1j ,k 2j Are all normal numbers;
is an adjustable normal constant; v. of j The angular speed expected value output by the outer loop nonlinear controller.
Further, the inner-loop nonlinear extended state observer is constructed as follows:
wherein alpha is 1i 、α 2i And epsilon 0i Are all non-negative constants, and are,
in the form of a non-linear function,
and
are each x 1i And x 2i Estimated values of two states, b i And the reciprocal of the moment of inertia of each channel, and the angular speed of the four-rotor aircraft in a body coordinate system.
Further, the inner loop nonlinear controller is constructed as follows:
wherein e is Ω ＝ω-ω r ＝[e ωi ] 3×1 ∈R 3×1 Error value, k, for three channel angular velocities pi ,k di ,k Ni Proportional, differential and non-linear terms, respectivelyIs a positive constant of (a) or (b),
is the reciprocal of the moment of inertia of each channel,
is a non-linear function.
Has the beneficial effects that:
according to the invention, the attitude of the four-rotor aircraft is controlled by constructing an attitude dynamics model of the four-rotor aircraft, analyzing the internal characteristics of the four-rotor aircraft, designing a double-loop active disturbance rejection nonlinear controller on the basis, and compensating external disturbance through the extended state observers of the inner loop and the outer loop, so that the four-rotor aircraft can accurately track a set expected attitude angle.
Drawings
Fig. 1 is a control schematic diagram of an attitude control method of a four-rotor aircraft in a Mars gust strong interference environment provided by the invention.
Detailed Description
The invention is described in detail below by way of example with reference to the accompanying drawings.
The active disturbance rejection control technology has the capability of processing wide problems such as unknown dynamics and external disturbance inside a system, and becomes an emerging technology in control engineering. The active disturbance rejection controller consists of a tracking differentiator, an extended state observer and a nonlinear control law, wherein the extended state observer is the most important part in the whole controller and estimates the state of the system through the output of the system. Therefore, under the condition of strong gust wind interference of a simulated mars, the active disturbance rejection control technology is applied to a four-rotor attitude system, and the influence of external disturbance with uncertainty on the flight effect of a four-rotor aircraft is resisted, so that stable flight of the unmanned aerial vehicle is realized.
Based on the consideration, the invention provides a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment, which has the core idea that: firstly, inputting an expected attitude angle into a four-rotor aircraft, then respectively controlling and calculating errors of the three attitude angles through an outer ring angle controller, and compensating disturbance of the angle through an outer ring nonlinear extended state observer so as to solve the expected angular speed; the obtained expected angular velocity and an angular velocity feedback value obtained from the system are further input into an inner ring angular velocity controller, the angular velocity of the aircraft is controlled through an inner ring active disturbance rejection controller, the angular disturbance is compensated through an inner ring nonlinear expansion state observer, three torque forces required to be input into the aircraft are obtained, the pitch angle, the roll angle and the yaw angle of the aircraft system are respectively controlled through the torque of three channels, the obtained output value is fed back to the controller, and therefore the closed-loop attitude control of the whole system of the aircraft is achieved.
Therefore, different from the algorithm that the active disturbance rejection control is only adopted in the inner ring of the aircraft attitude control in the prior art, the method is additionally provided with the outer ring active disturbance rejection controller and the outer ring nonlinear state observer so as to deal with the situation of strong gust wind disturbance in a mars background.
The invention provides a method for controlling the attitude of a four-rotor aircraft in a Mars gust strong interference environment, which specifically comprises the following steps of:
The existing four-rotor aircraft attitude dynamics model is adopted and is expressed as follows:
wherein ω = [ p, q, r] T ∈R 3 Eta = [ phi, theta, psi ] for the angular velocity of the aircraft in the coordinate system of the aircraft body] T ∈R 3 Is the Euler angle of the aircraft in an inertial coordinate system, phi is the pitch angle of the aircraft, theta is the roll angle of the aircraft, and psi is the yaw angle of the aircraft, W -1 As a rotating matrix such asShown below:
τ is the aircraft input (torque) as follows:
wherein k is 1 Is the thrust factor, k 2 Is the tension factor, omega i Is the angular velocity of the ith motor, i is a positive integer, l 1 And l 2 Representing the distance of the rotor to the x and y axes, respectively.
And 2, constructing an outer ring controller and an inner ring controller, wherein the outer ring controller comprises an outer ring nonlinear extended state observer and an outer ring nonlinear controller, and the inner ring controller comprises an inner ring nonlinear extended state observer and an inner ring nonlinear controller.
Different from the structural form of the existing linear state observer, on the basis of linearity, the invention adds nonlinear terms to the calculation of two state estimation values, and realizes better anti-disturbance effect by adjusting the parameter sigma of a nonlinear function. The outer ring nonlinear extended state observer constructed by the invention has the following structural form:
wherein, y 1 ＝η＝[φ,θ,ψ] T ＝[y 1j ] 3×1 ∈R 3×1 J = phi, theta, psi and y 2 ＝[y 2j ] 3×1 ∈R 3×1 J = phi, theta, psi respectively denote external disturbance values,
and
respectively representy 1 And y 2 Corresponding state estimation value, v j Represents the output value of the outer-loop controller,
as state estimates
The value of the first derivative of (a),
as state estimates
First derivative value of, beta 1j And ε 1j Are all positive adjustable constants that can be adjusted,
the expression for a nonlinear function is:
wherein, sigma is a positive adjustable constant, and the tracking effect of the whole system to the expected attitude angle is improved by adjusting the value of sigma.
Different from the existing linear active disturbance rejection controller, the outer loop controller further increases the nonlinear control law on the basis of the linear PI controller and the state feedback quantity in order to enable the error value to tend to 0 in a limited time, and meanwhile, the PI control rate and the nonlinear control rate supplement each other, so that on one hand, the outer loop controller has a good disturbance rejection effect, and on the other hand, convergence is accelerated. The structural form of the outer ring nonlinear controller constructed by the invention is as follows:
wherein e is η ＝η-η d ＝[e ηj ] 3×1 ∈R 3×1 ,j＝Phi, theta, psi, error values representing pitch, roll and yaw angles; k is a radical of formula 0j ,k 1j ,k 2j Is a normal number;
is an adjustable normal number; v. of j Is the output of the outer loop nonlinear controller, i.e. the desired value of angular velocity, and is also the input value of the inner loop controller.
Furthermore, the attitude dynamics model of a four-rotor aircraft can also be expressed in the form:
further translated into a state space expression:
wherein x is 1 (t)＝ω，x 2 (t) = f (t), and
is bounded.
On the basis, the invention carries out the following improvement on the representation method of three channels, and the attitude dynamic model is represented as follows:
on the basis, the structural form of the inner-loop nonlinear extended state observer constructed by the invention is as follows:
wherein alpha is 1i 、α 2i And ε 0i Are all non-negative constants, and are,
in the form of a non-linear function,
and
are respectively x 1i And x 2i Estimated values of two states, b i Is the reciprocal of the moment of inertia of each channel. At the same time, the user can select the required time,
and the inner-loop disturbance compensation term is input into the inner-loop nonlinear controller.
The inner loop nonlinear controller constructed by the invention adds a nonlinear term taking a nonlinear function as a control rate on the basis of a PI control law and a state feedback quantity, so that the controller has better adaptability and robustness, wherein the control rate is expressed as follows:
wherein e is Ω ＝ω-ω r ＝[e ωi ] 3×1 ∈R 3×1 Error value, k, for three channel angular velocities pi ,k di ,k Ni Respectively, proportional, differential and non-linear terms,
is the reciprocal of the moment of inertia of each channel, u i Is a four-rotor aircraft control variable.
Step 3, in actual control, the outer ring nonlinear extended state observer outputs the expected angular velocity value by the outer ring nonlinear controller and the value output by the four-rotor aircraftThe angle feedback value is used as input, and the output state estimation value is obtained through calculation
Outer loop nonlinear controller with desired angle of direct input, angle of quad-rotor aircraft output, and state estimate
Calculating the expected value v of the angular velocity of the output as input i (ii) a The inner-loop nonlinear extended state observer takes an angular velocity feedback value output by the quadrotor aircraft and an angular acceleration value output by the inner-loop nonlinear controller as inputs, and calculates to obtain an output state estimation value
The inner loop linear controller outputs angular speed expected value from the outer loop controller, angular speed feedback value from the four-rotor aircraft, and state estimation value from the inner loop nonlinear extended state observer
For input, calculating the control quantity u of the four-rotor aircraft to obtain output i (ii) a Will control the quantity u i And inputting the data into a posture dynamics model of the four-rotor aircraft to realize the control of the four-rotor aircraft.
Wherein, four rotor crafts are equivalent to the aircraft attitude dynamic system in the actual control.
In the aircraft framework, under the control of the inner ring and the outer ring, compared with a common active disturbance rejection controller, the double-ring active disturbance rejection controller provided by the invention can enable the aircraft to have better anti-jamming capability, and can also enable an error value to be more rapidly converged, so that an aircraft system can rapidly and accurately reach an expected attitude angle, and effective attitude control of the aircraft under a strong interference environment is realized.
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 (5)
1. A method for controlling the attitude of a four-rotor aircraft in a Mars gust interference environment is characterized by comprising the following steps:
step 1, establishing a posture dynamics model of a four-rotor aircraft;
step 2, constructing an outer ring controller and an inner ring controller, wherein the outer ring controller comprises an outer ring nonlinear extended state observer and an outer ring nonlinear controller, and the inner ring controller comprises an inner ring nonlinear extended state observer and an inner ring nonlinear controller;
and 3, in actual control, the outer ring nonlinear extended state observer takes the angular speed expected value output by the outer ring nonlinear controller and the angle feedback value output by the four-rotor aircraft as input, and the output state estimated value is obtained through calculation
Outer loop nonlinear controller with desired angle of direct input, angle of quad-rotor aircraft output, and state estimate
For input, the expected value v of the angular velocity is calculated j As an output; the inner-loop nonlinear extended state observer takes an angular velocity feedback value output by the quadrotor aircraft and a control quantity output by the inner-loop nonlinear controller as inputs, and calculates to obtain an output state estimation value
The inner loop linear controller outputs angular speed expected value from the outer loop controller, angular speed feedback value from the four-rotor aircraft, and state estimation value from the inner loop nonlinear extended state observer
Calculating to obtain output for inputControl u of four-rotor aircraft i (ii) a Will control the quantity u i And inputting the data into a posture dynamics model of the four-rotor aircraft to realize the control of the four-rotor aircraft.
2. The attitude control method according to claim 1, wherein the outer-loop nonlinear extended state observer in step 2 is constructed in the form as follows:
wherein, y 1 ＝η＝[φ,θ,ψ] T ＝[y 1j ] 3×1 ∈R 3×1 J = phi, theta, psi and y 2 ＝[y 2j ] 3×1 ∈R 3×1 J = phi, theta, psi respectively represent external disturbance values, phi being the pitch angle of the quad-rotor craft, theta being the roll angle of the quad-rotor craft, psi being the yaw angle of the quad-rotor craft,
and
respectively represent y 1 And y 2 Corresponding state estimation value, v j Represents the output value of the outer loop non-linear controller,
is a state estimation value
The value of the first derivative of (a),
as state estimates
First derivative value of, beta 1j And ε 1j Are all positive adjustable constants which are constant in number,
the expression for the nonlinear function is:
where σ is a positive adjustable constant.
3. The attitude control method according to claim 2, wherein the outer-loop nonlinear controller is constructed in the form as follows:
wherein e is η ＝η-η d ＝[e ηj ] 3×1 ∈R 3×1 J = phi, theta, psi is the error value of the pitch angle, the roll angle and the yaw angle; k is a radical of formula 0j ,k 1j ,k 2j Are all normal numbers;
is an adjustable normal number; v. of j Is the angular velocity desired value output by the outer loop nonlinear controller.
4. The attitude control method according to claim 1, wherein the inner-loop nonlinear extended state observer is configured as follows:
wherein alpha is 1i 、α 2i And ε 0i Are all non-negative constants, and are,
in the form of a non-linear function,
and
are each x 1i And x 2i Estimated values of two states, b i And the reciprocal of the moment of inertia of each channel, and the angular speed of the four-rotor aircraft in a body coordinate system.
5. The attitude control method according to claim 1, wherein the inner-loop nonlinear controller is configured as follows:
wherein e is Ω ＝ω-ω r ＝[e ωi ] 3×1 ∈R 3×1 Error value, k, for angular velocity of three channels pi ,k di ,k Ni Respectively, proportional, differential and non-linear terms,
is the reciprocal of the moment of inertia of each channel,
is a non-linear function.
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