CN114987800A - Spacecraft high-precision attitude control method based on sliding mode interference observation - Google Patents
Spacecraft high-precision attitude control method based on sliding mode interference observation Download PDFInfo
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Abstract
A spacecraft high-precision attitude control method based on sliding mode interference observation comprises the following steps: determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft; the PD controller determines the attitude control moment of the spacecraft system when interference is not considered according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft; determining the attitude control moment of the spacecraft system under the interference consideration by utilizing the interference estimation value and the attitude control moment of the spacecraft system under the interference consideration; determining a three-axis attitude angle of the spacecraft and an angular speed of the spacecraft by utilizing total interference and an attitude control moment of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft; the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular velocity of the spacecraft, and the interference estimation value is fed back and then used for determining the attitude control moment of the spacecraft system when the interference is considered so as to realize closed-loop control.
Description
Technical Field
The invention relates to a spacecraft high-precision attitude control method based on sliding mode interference observation, and belongs to the technical field of spacecraft control.
Background
The current spacecraft undertakes a complex on-orbit task and needs to carry various large-area antenna loads, so that the structure of the spacecraft is more complex; the difficulty and importance of the attitude control problem of the flexible spacecraft with high precision and high stability are increasingly prominent. The spacecraft can be continuously influenced by various interferences and uncertainties from the outside and the spacecraft, and the large-amplitude vibration of the flexible accessories is extremely easy to excite when the spacecraft is subjected to attitude control, so that the stability of the spacecraft body is influenced.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the defects of the prior art are overcome, and the problem of high-precision attitude control of the spacecraft under the conditions of unknown interference and model uncertainty is solved.
The purpose of the invention is realized by the following technical scheme:
a spacecraft high-precision attitude control method based on sliding mode interference observation comprises the following steps:
determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
determining the attitude control moment of the spacecraft system under the interference consideration by utilizing the interference estimation value and the attitude control moment of the spacecraft system under the interference consideration;
determining a three-axis attitude angle of the spacecraft and an angular speed of the spacecraft by utilizing total interference and an attitude control moment of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular velocity of the spacecraft, and the interference estimation value is fed back and then used for determining the attitude control moment of the spacecraft system when the interference is considered so as to realize closed-loop control.
Preferably, the disturbance observer is in the specific form:
wherein s is called an auxiliary sliding modulus,
are state variables of the spacecraft control system,
represents the three-axis attitude angular velocity of the spacecraft,
ω＝[ω x ω y ω z ] T is the three-axis angular velocity vector, omega, of the spacecraft body relative to the inertial coordinate system x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertial angular velocity, J is the star rotational inertia matrix, z is the auxiliary state variable,
for the interference estimation, the output of the PD controller when considering the interference is u, alpha, beta and gamma are interference estimation coefficients, and p>q are all positive odd numbers, sgn represents a sign function, the sign omega × An antisymmetric matrix as follows:
preferably, the output u of the PD controller is selected irrespective of disturbances c The method comprises the following specific steps:
wherein, K p 、K d Are all 3X 3 dimensional positive definite diagonal control gain matrix, x e1 ＝θ d - θ, the attitude angle error of the spacecraft,
is the angular velocity error, theta, of the spacecraft d 、
Respectively, the three-axis expected attitude angle and the expected angular velocity, theta is respectively the three-axis attitude angle of the spacecraft, x 2 × ＝ω × 。
Preferably, the total interference is:
the point above the parameter in the formula represents the derivative, d 0 Is the environmental disturbance torque vector, F, to which the spacecraft is subjected si Is a 3 Xn dimensional modal vibration and spacecraft body rotation coupling coefficient matrix, n is the modal order, eta i In the n-dimensional flexible modal coordinate, i is 1, 2 represents sailboard number, xi i And Ω i And the n-dimensional diagonal matrix respectively represents the damping ratio of the flexible accessory and the modal frequency matrix.
Preferably, the attitude control moment of the spacecraft system in consideration of the disturbance is:
A spacecraft high-precision attitude control system based on sliding mode interference observation comprises: the system comprises a PD controller, a spacecraft model module, an interference observer, a first module and a second module;
the first module is used for determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
the second module determines the attitude control moment of the spacecraft system during the interference by using the interference estimated value and not considering the attitude control moment of the spacecraft system during the interference;
the spacecraft model module determines a three-axis attitude angle of a spacecraft and an angular speed of the spacecraft by utilizing total interference and attitude control torque of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
and the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular speed of the spacecraft, and feeds the interference estimation value back to the second module to realize closed-loop control.
Compared with the prior art, the invention has the following beneficial effects:
(1) the method adopts the interference observer combined with PD control to process the attitude control problem of the flexible spacecraft, can effectively inhibit interference and flexible vibration, and improves the precision and stability of a control system;
(2) the invention can effectively estimate uncertain dynamic characteristics such as external interference and the like, and introduces the estimation result into the control system for compensation, thereby effectively enhancing the robustness of the system;
(3) the improved sliding mode disturbance observer is adopted to estimate external disturbance and flexible vibration influence, so that the convergence speed of the observer can be effectively improved;
(4) according to the invention, the improved sliding mode disturbance observer is adopted to estimate external disturbance and flexible vibration influence, and the total disturbance estimation result is more in line with the actual disturbance characteristic.
Drawings
FIG. 1 is a block diagram of a high-precision attitude control system of a spacecraft based on sliding mode interference observation provided by the invention.
Fig. 2 is a simulation diagram total interference estimation error of the present invention.
FIG. 3 is a diagram of the invention's pose roll angle.
Fig. 4 is a simulated map attitude pitch angle of the present invention.
FIG. 5 is a simulated graph attitude yaw angle of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
Example 1:
the invention provides a spacecraft high-precision attitude control method based on sliding mode interference observation, which comprises the following steps of:
simplifying the system into a state equation form according to the flexible spacecraft dynamics and kinematics equation, namely a spacecraft model is as follows;
wherein x is 1 ＝θ、
Is a state variable of a spacecraft control system, theta represents a three-axis attitude angle of the spacecraft,
the three-axis attitude angular velocity of the spacecraft is shown, u is the output of a PD controller when the interference is considered, namely the three-axis control moment vector of the spacecraft provided by an actuating mechanism,
ω＝[ω x ω y ω z ] T is a spacecraft body opposite toThree-axis angular velocity vector, ω, of the inertial frame x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertia angular velocity, J is the star rotational inertia matrix,
the sum of the environmental disturbance, the flexural vibration effect and the uncertainty, i.e. the total disturbance, is represented by the points above the parameters, d 0 Is the environmental disturbance torque vector, F, to which the spacecraft is subjected si Is a 3 Xn dimensional modal vibration and spacecraft body rotation coupling coefficient matrix, n is the modal order, eta i Is n-dimensional flexible modal coordinate, i is 1, 2 represents sailboard number, xi i And Ω i Is an n-dimensional diagonal matrix, and respectively represents the damping ratio and the modal frequency matrix of the flexible accessory, and the symbol omega × An antisymmetric matrix as follows:
determining a sliding mode disturbance observer;
wherein s is called an auxiliary sliding modulus, z is an auxiliary state variable,
for the interference estimation, α, β, γ are interference estimation coefficients, p>q are positive odd numbers, sgn represents a sign function.
Determining PD controller output u without considering interference c ；
Wherein u is c To disregard the PD controller output of the system when interference, K p 、K d For a positive definite diagonal control gain matrix of 3 x 3 dimensions, x e1 ＝x d -x 1 ＝θ d -θ、
Is the attitude angle and angular velocity error, x d ＝θ d ，
θ d 、
Attitude angles and angular velocities are desired for the three axes.
Using interference estimates
Determining a PD controller output u when considering interference;
the structural block diagram of the flexible spacecraft attitude control system is shown in figure 1. The simulation results are shown in fig. 2-5, which are the interference estimation error, the attitude roll angle, the attitude pitch angle and the attitude yaw angle, respectively. Wherein PD represents a simulation result of a single PD controller, SMDO represents a simulation result of PD control combined with general sliding mode disturbance observation, and ISMDO represents a simulation result of PD control combined with an improved sliding mode disturbance observer. The method can effectively estimate the environmental interference, the flexible vibration and the influence of model uncertainty, thereby realizing good compensation effect.
Example 2:
a spacecraft high-precision attitude control method based on sliding mode interference observation comprises the following steps:
determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
determining the attitude control moment of the spacecraft system under the interference consideration by utilizing the interference estimation value and the attitude control moment of the spacecraft system under the interference consideration;
determining a three-axis attitude angle of the spacecraft and an angular speed of the spacecraft by utilizing total interference and an attitude control moment of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular velocity of the spacecraft, and the interference estimation value is fed back and then used for determining the attitude control moment of the spacecraft system when the interference is considered so as to realize closed-loop control.
Optionally, the disturbance observer is in a specific form:
wherein s is called an auxiliary sliding modulus,
are state variables of the spacecraft control system,
represents the three-axis attitude angular velocity of the spacecraft,
ω＝[ω x ω y ω z ] T is the three-axis angular velocity vector, omega, of the spacecraft body relative to the inertial coordinate system x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertial angular velocity, J is the star rotational inertia matrix, z is the auxiliary state variable,
for the interference estimation, the output of the PD controller when considering the interference is u, alpha, beta and gamma are interference estimation coefficients, and p>q are all positive odd numbers, sgn represents a sign function, the sign omega × An antisymmetric matrix as follows:
optionally, the output u of the PD controller when interference is not considered c The method specifically comprises the following steps:
wherein, K p 、K d Are all 3X 3 dimensional positive definite diagonal control gain matrix, x e1 ＝θ d - θ, the attitude angle error of the spacecraft,
is the angular velocity error, theta, of the spacecraft d 、
Three-axis desired attitude angles and desired angular velocities, theta is the three-axis attitude angle of the spacecraft,
optionally, the total interference is:
the point above the parameter in the formula represents the derivative, d 0 Is the environmental disturbance torque vector, F, to which the spacecraft is subjected si Is a 3 Xn dimensional modal vibration and spacecraft body rotation coupling coefficient matrix, n is the modal order, eta i Is n-dimensional flexible modal coordinate, i is 1, 2 represents sailNumber of plate, xi i And Ω i And the n-dimensional diagonal matrix respectively represents the damping ratio of the flexible accessory and the modal frequency matrix.
Optionally, the total disturbance is a sum of an environmental disturbance, a flexural vibration influence and an uncertainty.
Optionally, the attitude control moment of the spacecraft system when considering the interference is:
A computer-readable storage medium, on which a computer program is stored, which, when executed by a processor, implements the spacecraft high-precision attitude control method described above.
A spacecraft high-precision attitude control system based on sliding mode interference observation comprises: the system comprises a PD controller, a spacecraft model module, an interference observer, a first module and a second module;
the first module is used for determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
the second module determines the attitude control moment of the spacecraft system during interference by using the interference estimated value and without considering the attitude control moment of the spacecraft system during interference;
the spacecraft model module determines a three-axis attitude angle of a spacecraft and an angular speed of the spacecraft by utilizing total interference and attitude control torque of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
and the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular speed of the spacecraft, and feeds the interference estimation value back to the second module to realize closed-loop control.
Optionally, the disturbance observer is in a specific form:
wherein s is called an auxiliary sliding modulus,
are state variables of the spacecraft control system,
represents the three-axis attitude angular velocity of the spacecraft,
ω＝[ω x ω y ω z ] T is the three-axis angular velocity vector, ω, of the spacecraft body relative to the inertial coordinate system x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertial angular velocity, J is the star rotational inertia matrix, z is the auxiliary state variable,
for the interference estimation, the output of the PD controller when considering the interference is u, alpha, beta and gamma are interference estimation coefficients, and p>q are all positive odd numbers, sgn represents a sign function, the sign omega × An antisymmetric matrix as follows:
optionally, the output u of the PD controller when interference is not considered c The method specifically comprises the following steps:
wherein, K p 、K d Are all a 3 x 3 dimensional positive definite diagonal control gain matrix, x e1 ＝θ d - θ, the attitude angle error of the spacecraft,
is the angular velocity error, theta, of the spacecraft d 、
Respectively, the three-axis expected attitude angle and the expected angular velocity, theta is respectively the three-axis attitude angle of the spacecraft,
those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (10)
1. A spacecraft high-precision attitude control method based on sliding mode interference observation is characterized by comprising the following steps:
determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
determining the attitude control moment of the spacecraft system under the interference consideration by utilizing the interference estimation value and the attitude control moment of the spacecraft system under the interference consideration;
determining a three-axis attitude angle of the spacecraft and an angular speed of the spacecraft by using total interference and an attitude control moment of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
the interference observer determines an interference estimation value by utilizing the three-axis attitude angle of the spacecraft and the angular velocity of the spacecraft, and the interference estimation value is fed back and then used for determining the attitude control moment of the spacecraft system when the interference is considered so as to realize closed-loop control.
2. A spacecraft high precision attitude control method according to claim 1, wherein the disturbance observer is in the form of:
wherein s is called an auxiliary sliding modulus,
are state variables of the spacecraft control system,
represents the three-axis attitude angular velocity of the spacecraft,
ω＝[ω x ω y ω z ] T is the three-axis angular velocity vector, omega, of the spacecraft body relative to the inertial coordinate system x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertial angular velocity, J is the star rotational inertia matrix, z is the auxiliary state variable,
for the interference estimation, the output of the PD controller when considering the interference is u, alpha, beta and gamma are interference estimation coefficients, and p>q are all positive odd numbers, sgn represents a sign function, the sign omega × Is an antisymmetric momentArraying:
3. a spacecraft high accuracy attitude control method according to claim 2, characterized in that the output u of the PD controller without considering disturbances c The method specifically comprises the following steps:
wherein, K p 、K d Are all 3X 3 dimensional positive definite diagonal control gain matrix, x e1 ＝θ d - θ, the attitude angle error of the spacecraft,
is the angular velocity error, theta, of the spacecraft d 、
Respectively, the three-axis expected attitude angle and the expected angular velocity, theta is respectively the three-axis attitude angle of the spacecraft,
4. a spacecraft high accuracy attitude control method according to claim 2, wherein said total disturbance is:
the point above the parameter in the formula represents the derivative, d 0 Is the environmental disturbance torque vector, F, to which the spacecraft is subjected si Is a 3 Xn dimensional modal vibration and spacecraft body rotation coupling coefficient matrix, n is a modal order，η i In the n-dimensional flexible modal coordinate, i is 1, 2 represents sailboard number, xi i And Ω i And the n-dimensional diagonal matrix respectively represents the damping ratio of the flexible accessory and the modal frequency matrix.
5. A spacecraft high accuracy attitude control method according to claim 1 wherein the total disturbance is the sum of environmental disturbance, flexural vibration effect and uncertainty.
6. A spacecraft high accuracy attitude control method according to any of claims 1 to 5, wherein the attitude control moments of the spacecraft system under consideration of the disturbance are:
7. A computer-readable storage medium, characterized in that a computer program is stored thereon, which when executed by a processor, implements the spacecraft high precision attitude control method of any of claims 1 to 5.
8. A spacecraft high-precision attitude control system based on sliding mode interference observation is characterized by comprising: the system comprises a PD controller, a spacecraft model module, an interference observer, a first module and a second module;
the first module is used for determining attitude angle errors of the spacecraft and angular speed errors of the spacecraft;
the PD controller determines the attitude control moment of the spacecraft system without considering the interference according to the attitude angle error of the spacecraft and the angular velocity error of the spacecraft;
the second module determines the attitude control moment of the spacecraft system during the interference by using the interference estimated value and not considering the attitude control moment of the spacecraft system during the interference;
the spacecraft model module determines a three-axis attitude angle of a spacecraft and an angular speed of the spacecraft by utilizing total interference and attitude control torque of a spacecraft system during interference so as to complete high-precision attitude control of the spacecraft;
the interference observer determines an interference estimation value by using the three-axis attitude angle of the spacecraft and the angular velocity of the spacecraft, and feeds the interference estimation value back to the second module to realize closed-loop control.
9. A spacecraft high accuracy attitude control system according to claim 8, wherein the disturbance observer is in the form of:
wherein s is called an auxiliary sliding modulus,
are state variables of the spacecraft control system,
representing the three-axis attitude angular velocity of the spacecraft,
ω＝[ω x ω y ω z ] T is the three-axis angular velocity vector, ω, of the spacecraft body relative to the inertial coordinate system x Is the angular velocity of inertia of the rolling axis, omega y Is the angular velocity of inertia of the pitch axis, omega z Is the yaw axis inertial angular velocity, J is the star rotational inertia matrix, z is the auxiliary state variable,
for interference estimationThe value of the PD controller output when considering the interference is u, alpha, beta, gamma are interference estimation coefficients, p>q are all positive odd numbers, sgn represents a sign function, the sign omega × An antisymmetric matrix as follows:
10. a spacecraft high accuracy attitude control system according to claim 9, wherein the output u of the PD controller without taking interference into account c The method comprises the following specific steps:
wherein, K p 、K d Are all 3X 3 dimensional positive definite diagonal control gain matrix, x e1 ＝θ d - θ, an attitude angle error of the spacecraft,
is the angular velocity error, theta, of the spacecraft d 、
Respectively, the three-axis expected attitude angle and the expected angular velocity, theta is respectively the three-axis attitude angle of the spacecraft,
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| KR102605904B1 (en) * | 2023-04-06 | 2023-11-27 | 세종대학교 산학협력단 | Method and apparatus for attitude tracking control for rigid spacecraft with arbitrary disturbance |
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