CN111547275A - Spacecraft three-phase control robust self-adaptive multi-level cooperation method - Google Patents

Spacecraft three-phase control robust self-adaptive multi-level cooperation method Download PDF

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CN111547275A
CN111547275A CN202010351841.2A CN202010351841A CN111547275A CN 111547275 A CN111547275 A CN 111547275A CN 202010351841 A CN202010351841 A CN 202010351841A CN 111547275 A CN111547275 A CN 111547275A
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spacecraft
load
cmg
star
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CN111547275B (en
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姚宁
袁利
汤亮
关新
王有懿
宗红
郭子熙
张科备
郝仁剑
冯骁
刘昊
龚立纲
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Beijing Institute of Control Engineering
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces

Abstract

A spacecraft three-control robust self-adaptive multi-stage cooperation method is suitable for large satellite platforms with ultrahigh precision, ultrahigh stability and ultrahigh agility on payload postures, such as astronomical observation. Different from the traditional PID control algorithm, the method combines the robustness characteristic of sliding mode control on the sliding mode surface and the characteristic that adaptive control can estimate parameters on line, and carries out two-stage composite control of the star-active pointing hyperstatic platform. The multi-level cooperative control idea is as follows: 1) an active pointing hyperstatic platform is arranged between a load and a spacecraft body, and control parameters of the active pointing hyperstatic platform are designed according to the mass characteristics of the spacecraft body and the load; 2) and designing a star body robust adaptive controller considering bandwidth constraint by combining the ideas of sliding mode control and adaptive control, so that the star body controller can be matched with an active directional hyperstatic platform, and the three-mode hyperstatic control on the load is realized.

Description

Spacecraft three-phase control robust self-adaptive multi-level cooperation method
Technical Field
The invention belongs to the field of spacecraft attitude control, and relates to a multistage cooperative control method for a large-scale spacecraft system.
Background
With the continuous improvement of astronomical observation and on-orbit service requirements, the design ideas and practices of the modern large-scale satellite platform are revolutionarily changed, the complexity of a control system of the satellite is continuously improved, and meanwhile, the requirements of attitude precision and stability of satellite payloads are also developed towards individuation and refinement, so that a novel control method with three-super performance is necessary to be developed for one type of large-scale satellite platform, and the attitude requirements of various payloads in the future are met. The three-step process refers to 'ultrahigh precision, ultrahigh stability and hypersensitiveness'.
In the traditional spacecraft control system at present, a star body and a load are always in rigid connection, and high-frequency and low-frequency micro-vibration in the star body is directly transmitted to the load, so that the imaging quality of an optical load is influenced. The attitude control bandwidth of the star body is limited by the step length of the controller and the frequency of the flexible accessory, and the real-time compensation of high-frequency jitter cannot be realized. To solve the problem, scholars at home and abroad research a three-stage and super-control method based on spacecraft body-active pointing super-static platform two-stage composite control, because the control structure is complicated, PID control is often adopted in two-stage control loops, but the following defects still exist:
the simple PID control algorithm can only realize gradual stabilization of the system and cannot ensure the rapidity of state quantity convergence; after the feedforward control is added, the speed of attitude tracking error convergence can be improved, but the vibration of the flexible accessory is additionally excited, the feedback control needs to be further carried out on the part of vibration in the control of actively pointing to the hyperstatic platform, and the additional energy consumption is increased.
Disclosure of Invention
The technical problem solved by the invention is as follows: the method is high in universality, can be used for a large satellite platform which considers the flexibility of a spacecraft body and is provided with any number of flexible accessories and CMG, can realize the rapid tracking of the expected attitude of a load under the condition of considering the uncertainty of a spacecraft body model, and can realize the real-time estimation and compensation of low-frequency disturbance generated by the flexible accessories and the inhibition of low-frequency micro-vibration; the problem of near original point bandwidth infinitely increased in traditional slipform control is solved for the star controller can effectively with surpass quiet platform controller phase-match, realize the three super control to the load.
The technical solution of the invention is as follows:
a spacecraft triple control robust self-adaptive multi-level cooperation method comprises the following steps:
(1) establishing a spacecraft system dynamic equation comprising a load dynamic equation, a spacecraft body dynamic equation and a flexible accessory vibration equation;
(2) designing a PD controller aiming at the active pointing hyperstatic platform;
(3) substituting the controller into the load dynamics equation to obtain an error equation of the load attitude:
(4) calculating a proportional control coefficient k from the expected natural frequency according to the calculation formula of the undamped natural frequency of the loadppAnd distributing the control quantity to each actuator;
(5) designing a sliding mode surface, and designing a star robust adaptive controller considering bandwidth constraint and an adaptive law of a nonlinear term based on the sliding mode surface;
(6) selecting boundary layer parameters according to the amplitude of the high-frequency jitter of the star attitude under the PD control action, and determining a star error equation in a segmented form;
(7) calculating the control bandwidth of the star body and the control bandwidth of the load controller, and designing the control parameters k and k2、k3And p, the control bandwidth of the spacecraft star state quantity error after entering the boundary layer is smaller than 1/10 of the bandwidth of the load controller, and meanwhile, the damping ratio of a star control loop is 2, so that the spacecraft triple control robust self-adaptive multi-stage cooperative control is realized.
Further, the spacecraft system comprises a flexible spacecraft body, any number of flexible accessories, a load, a control moment gyro CMG and an active pointing hyperstatic platform; the flexible accessory and the control moment gyroscope CMG are both arranged on the flexible spacecraft body; the active pointing hyperstatic platform is arranged between the load and the flexible spacecraft body, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the flexible spacecraft body; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor; the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the flexible spacecraft body; the linear motor is used for providing main power and realizing the attitude control of the load.
Furthermore, the control moment gyroscope CMG is arranged on the flexible spacecraft body in a distributed mode, so that attitude control of the star body and vibration suppression of the star body structure are achieved.
Further, the load dynamics equation is:
Figure BDA0002472184860000031
wherein the content of the first and second substances,
Figure BDA0002472184860000032
the spacecraft body dynamic equation is as follows:
Figure BDA0002472184860000033
wherein the content of the first and second substances,
Figure BDA0002472184860000034
the flexible attachment vibration equation is:
Figure BDA0002472184860000035
wherein E is3Is an identity matrix of order 3, EmbIs an mb-order identity matrix, rdbη as the relative position vector from the center of mass of the spacecraft body to the center of the lower plane of the active pointing hyperstatic platformbIs a modal coordinate array of the spacecraft body, ηakA modal coordinate array for the kth flexible attachment;
mps，Ips，mbs，Ibs，Pb，Hbsrespectively mass and inertia of a load, total mass and total inertia of a spacecraft platform, modal momentum and modal angular momentum to a body coordinate system, wherein the spacecraft platform comprises a spacecraft body, a flexible attachment and a CMG; fbak、RbakRespectively a translational coupling coefficient and a rotational coupling coefficient of the flexible accessory vibration to the spacecraft body, Λb、ξbRespectively eigenvalue matrix and damping ratio matrix of the spacecraft body, Λak、ξakRespectively, a matrix of eigenvalues and a matrix of damping ratios, h, of the flexible accessoriescIs the total angular momentum, v, of the CMGp、ωp、vb、ωbRespectively the speed and angular velocity of the load, the speed and angular velocity of the spacecraft body, Fd、Td、TcDisturbance force, disturbance torque and control torque borne by the spacecraft body are respectively; t isci、RciRespectively a translational mode matrix and a rotational mode matrix at the mounting point of the ith CMG, Fdi、Tdi、TciDisturbance force, disturbance torque and control torque output by the ith CMG respectively; fujAnd FsjRespectively the action force of the jth actuator on the load and on the spacecraft body, rrpjAnd rrbjRespectively the installation position coordinates of the jth actuator on the load and the spacecraft body.
Further, in the above-mentioned case,
Fdi＝-AbgiAgfiΩi ×Ωi ×AfwiSwi
Tdi＝-rfbi-b ×AbgiAgfiΩi ×Ωi ×AfwiSwi-AbgiAgfiΩi ×Iwi-fΩi
Figure BDA0002472184860000041
Figure BDA0002472184860000042
Fuj＝(-kj(lj-lj0)+uj)suj
Fsj＝-Fuj
wherein omegaiThe rotor speed vector for the ith CMG is at ffiCoordinates under the system, Iwi-fFor the rotor at ffiThe non-zero term of off-diagonal elements of the rotational inertia under the system represents the magnitude of the dynamic unbalance of the CMG rotor, SwiFor the rotor at fwiThe static moment under the system represents the magnitude of the static unbalance of the CMG rotor, and the two terms are disturbance sources of high-frequency micro vibration generated by the CMG; r isfbi-bFrom the mass center of the body coordinate system to the geometric center of the rotorCoordinates of the relative position vector of (2) in the system; k is a radical ofjIs the stiffness of the diaphragm spring in the jth actuator, ljIs the length of the jth actuator at any time, lj0Is the length of the initial moment of the j-th actuator, ujIs the control quantity of the j-th actuator, sujThe coordinate of the jth actuator direction vector at any moment in the inertial system;
ffiis defined as follows: f. offiThe system is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor;
fwiis defined as follows: f. ofwiThe system is a rotor fixed connection coordinate system of the ith CMG and rotates along with the rotor;
in the above variables, AabAs a coordinate system fbTo faOf the conversion matrix fa、fbCan be an arbitrary coordinate system and can be used,
Figure BDA0002472184860000056
is obTo oaSagittal diameter of oa、obIs the origin of any two coordinate systems;
subscript e is an inertia system, b is a body system, p is a load coordinate system, gi is a frame coordinate system of the ith CMG, and fi is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor; wi is a rotor fixed coordinate system of the ith CMG and rotates along with the rotor;
r×is an antisymmetric matrix of a vector, and the calculation formula is as follows:
Figure BDA0002472184860000051
where r can be any vector.
Further, the designing the PD controller by actively pointing to the hyperstatic platform in the step (2) is specifically:
Figure BDA0002472184860000052
wherein tau isppControl moment, J, output for actively directing the load to the hyperstatic platformpA Jacobian matrix from a load motion space to an actuator motion space, u is an array of actuator control variables, and a corner mark 4:6 represents 4 to 6 rows, omegaprAngular velocity, k, of the desired movement for the loadppAnd kpdProportional and differential coefficients, e, of the load controller, respectivelyperrIs the load error attitude angle.
Further, the error equation of the load attitude in the step (3) is specifically as follows:
Figure BDA0002472184860000053
wherein k isjIs a diagonal array of diaphragm spring stiffness0The length of the actuator and the length array of the initial moment of the actuator are respectively.
Further, the calculation formula of the undamped natural frequency of the load in the step (4) is as follows:
Figure BDA0002472184860000054
the control amount is distributed to each actuator as follows:
Figure BDA0002472184860000055
further, in the step (5), the slip-form surface is as follows:
Figure BDA0002472184860000061
wherein e ═ eperr T,eη T]T，eηIs the error value of the modal coordinate;
the star robust adaptive controller based on sliding mode surface design and considering bandwidth constraint comprises the following steps:
Figure BDA0002472184860000062
the adaptation law of the nonlinear term is:
Figure BDA0002472184860000063
wherein the content of the first and second substances,
Figure BDA0002472184860000064
Figure BDA0002472184860000065
wherein Hbs0Is HbsEstimated value of k, k2，k3P is the controller gain coefficient, 0<p<1, boundary layer parameters and adaptive law coefficients.
Further, the star error equation in the segmented form of the step (6) is as follows:
when | e | ≧ is,
Figure BDA0002472184860000066
when-is < e < is present,
Figure BDA0002472184860000067
wherein f (ω)b) Is a nonlinear term in the kinetic equation.
Further, the star control bandwidth is:
Figure BDA0002472184860000068
the control bandwidth of the load controller is the undamped natural frequency of the load, and specifically includes:
Figure BDA0002472184860000071
by designing the control parameters k, k2、k3P, of
Figure BDA0002472184860000072
Damping ratio of star control loop:
Figure BDA0002472184860000073
therefore, three-phase control robust self-adaptive multi-level cooperative control of the spacecraft is realized.
Compared with the prior art, the invention has the advantages that:
1. the load control index is improved, and the control performance is excellent
The active pointing hyperstatic platform can effectively inhibit the high-frequency vibration of the star, so that the precision and the stability of the load attitude are improved by 1-2 orders of magnitude; the star controller can realize the rapid tracking of the expected attitude under the condition of considering the uncertainty of the flexible coupling coefficient of the spacecraft body, and can simultaneously realize the real-time estimation and compensation of the low-frequency disturbance generated by the flexible accessory, thereby realizing the inhibition of the low-frequency micro-vibration; the problem of near original point bandwidth infinitely increased in traditional slipform control is solved for the star controller can effectively with surpass quiet platform controller phase-match, realize the three super control to the load.
2. The control method has strong engineering realizability
The composite control method provided by the invention is suitable for a large-scale spacecraft system, the uncertainty of a model is considered by the star controller, when the star body adopts the assumption of 'a central rigid body + a flexible accessory', the form of the star controller is unchanged, and the control method has certain universality; the invention provides a control parameter design method for a hyperstatic platform controller and a star controller, and the controller parameters can be obtained through expected control performance, so that the engineering application and popularization are facilitated.
Drawings
FIG. 1 is a block diagram of a large spacecraft system to which the present invention is directed;
FIG. 2 is a flow chart of the method of the present invention;
FIG. 3 is a schematic diagram of loaded triaxial attitude accuracy without using multi-level cooperative control;
FIG. 4 is a schematic view of loaded triaxial attitude stability under a condition of not using multi-level cooperative control;
FIG. 5 is a schematic view of loaded triaxial attitude accuracy under a multi-level cooperative control scenario;
FIG. 6 is a schematic view of loaded triaxial attitude stability under a multi-level cooperative control condition.
Detailed Description
The invention provides a spacecraft three-control robust self-adaptive multi-stage cooperation method which is suitable for large satellite platforms with ultrahigh precision, ultrahigh stability and ultrahigh agility on payload postures, such as astronomical observation. Different from the traditional PID control algorithm, the method combines the robustness characteristic of sliding mode control on the sliding mode surface and the characteristic that adaptive control can estimate parameters on line, and carries out two-stage composite control of the star-active pointing hyperstatic platform. The multi-level cooperative control idea is as follows: 1) an active pointing hyperstatic platform is arranged between a load and a spacecraft body, and control parameters of the active pointing hyperstatic platform are designed according to the mass characteristics of the spacecraft body and the load; 2) and designing a star body robust adaptive controller considering bandwidth constraint by combining the ideas of sliding mode control and adaptive control, so that the star body controller can be matched with an active directional hyperstatic platform, and the three-mode hyperstatic control on the load is realized.
As shown in fig. 2, the three-phase control robust adaptive multi-level collaborative method for the spacecraft of the present invention specifically includes the following steps:
(1) as shown in fig. 1, the spacecraft system comprises a flexible spacecraft body, any number of flexible appendages, a load, control moment gyro CMG, and an active pointing hyperstatic platform; the flexible accessory and the control moment gyroscope CMG are both arranged on the flexible spacecraft body; the active pointing hyperstatic platform is arranged between the load and the flexible spacecraft body, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the flexible spacecraft body; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor; the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the flexible spacecraft body; the linear motor is used for providing main power and realizing the attitude control of the load.
The control moment gyroscope CMG is arranged on the flexible spacecraft body in a distributed mode, so that attitude control of a star body and vibration suppression of the star body structure are achieved.
Establishing a spacecraft system dynamic equation which comprises a load dynamic equation, a spacecraft body dynamic equation and a flexible accessory vibration equation as follows:
the load dynamics equation is:
Figure BDA0002472184860000091
wherein the content of the first and second substances,
Figure BDA0002472184860000092
the spacecraft body dynamic equation is as follows:
Figure BDA0002472184860000093
wherein the content of the first and second substances,
Figure BDA0002472184860000094
the flexible attachment vibration equation is:
Figure BDA0002472184860000095
wherein E is3Is an identity matrix of order 3, EmbIs an mb-order identity matrix, rdbη as the relative position vector from the center of mass of the spacecraft body to the center of the lower plane of the active pointing hyperstatic platformbIs a modal coordinate array of the spacecraft body, ηakA modal coordinate array for the kth flexible attachment;
mps，Ips，mbs，Ibs，Pb，Hbsmass, inertia, and spacecraft of the load, respectivelyThe spacecraft platform comprises a spacecraft body, a flexible accessory and a CMG (China Mobile gateway), wherein the spacecraft platform comprises a platform total mass, a total inertia, a modal momentum and a modal angular momentum to a body coordinate system; fbak、RbakRespectively a translational coupling coefficient and a rotational coupling coefficient of the flexible accessory vibration to the spacecraft body, Λb、ξbRespectively eigenvalue matrix and damping ratio matrix of the spacecraft body, Λak、ξakRespectively, a matrix of eigenvalues and a matrix of damping ratios, h, of the flexible accessoriescIs the total angular momentum, v, of the CMGp、ωp、vb、ωbRespectively the speed and angular velocity of the load, the speed and angular velocity of the spacecraft body, Fd、Td、TcDisturbance force, disturbance torque and control torque borne by the spacecraft body are respectively; t isci、RciRespectively a translational mode matrix and a rotational mode matrix at the mounting point of the ith CMG, Fdi、Tdi、TciDisturbance force, disturbance torque and control torque output by the ith CMG respectively; fujAnd FsjRespectively the action force of the jth actuator on the load and on the spacecraft body, rrpjAnd rrbjRespectively the installation position coordinates of the jth actuator on the load and the spacecraft body.
Fdi＝-AbgiAgfiΩi ×Ωi ×AfwiSwi
Tdi＝-rfbi-b ×AbgiAgfiΩi ×Ωi ×AfwiSwi-AbgiAgfiΩi ×Iwi-fΩi
Figure BDA0002472184860000101
Figure BDA0002472184860000102
Fuj＝(-kj(lj-lj0)+uj)suj
Fsj＝-Fuj
Wherein omegaiThe rotor speed vector for the ith CMG is at ffiCoordinates under the system, Iwi-fFor the rotor at ffiThe non-zero term of off-diagonal elements of the rotational inertia under the system represents the magnitude of the dynamic unbalance of the CMG rotor, SwiFor the rotor at fwiThe static moment under the system represents the magnitude of the static unbalance of the CMG rotor, and the two terms are disturbance sources of high-frequency micro vibration generated by the CMG; r isfbi-bCoordinates of a relative position vector from a mass center of the body coordinate system to a geometric center of the rotor under the body coordinate system; k is a radical ofjIs the stiffness of the diaphragm spring in the jth actuator, ljIs the length of the jth actuator at any time, lj0Is the length of the initial moment of the j-th actuator, ujIs the control quantity of the j-th actuator, sujThe coordinate of the jth actuator direction vector at any moment in the inertial system;
ffiis defined as follows: f. offiThe system is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor;
fwiis defined as follows: f. ofwiThe system is a rotor fixed connection coordinate system of the ith CMG and rotates along with the rotor;
in the above variables, AabAs a coordinate system fbTo faOf the conversion matrix fa、fbCan be an arbitrary coordinate system and can be used,
Figure BDA0002472184860000103
is obTo oaSagittal diameter of oa、obIs the origin of any two coordinate systems;
subscript e is an inertia system, b is a body system, p is a load coordinate system, gi is a frame coordinate system of the ith CMG, and fi is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor; wi is a rotor fixed coordinate system of the ith CMG and rotates along with the rotor;
r×is an antisymmetric matrix of a vector, and the calculation formula is as follows:
Figure BDA0002472184860000111
where r can be any vector.
(2) Designing a PD controller aiming at an active pointing hyperstatic platform:
Figure BDA0002472184860000112
wherein tau isppControl moment, J, output for actively directing the load to the hyperstatic platformpA Jacobian matrix from a load motion space to an actuator motion space, u is an array of actuator control variables, and a corner mark 4:6 represents 4 to 6 rows, omegaprAngular velocity, k, of the desired movement for the loadppAnd kpdProportional and differential coefficients, e, of the load controller, respectivelyperrIs the load error attitude angle.
Under the assumption of slight amplitude motion and small angle, the controller is substituted into the load dynamics equation given in the step (1) to obtain an error equation of the load attitude:
Figure BDA0002472184860000113
wherein k isjIs a diagonal array of diaphragm spring stiffness0The length of the actuator and the length array of the initial moment of the actuator are respectively.
(3) And selecting the natural frequency of the load loop according to the frequency of the CMG disturbance, so that the high-frequency disturbance on the spacecraft body can be attenuated in a larger amplitude. Calculating a formula according to the undamped natural frequency of the load:
Figure BDA0002472184860000114
calculating a proportional control coefficient k from the desired natural frequency, the mass characteristics of the spacecraft body and the loadpp＝9×104N/m, designing a differential coefficient k through a desired system disturbance transfer characteristicpd800 Ns/m. Go toAfter the load controller and the controller parameters are designed, the control quantity is distributed to each actuator according to the following formula:
Figure BDA0002472184860000115
(4) designing a slip form surface:
Figure BDA0002472184860000121
wherein e ═ eperr T,eη T]T，eηIs the error value of the modal coordinates.
A star robust adaptive controller considering bandwidth constraint based on sliding mode surface design comprises:
Figure BDA0002472184860000122
the adaptation rate of the nonlinear term is:
Figure BDA0002472184860000123
wherein the content of the first and second substances,
Figure BDA0002472184860000124
Figure BDA0002472184860000125
wherein Hbs0Is HbsEstimated value of k, k2，k3P is the controller gain coefficient, 0<p<1, boundary layer parameter and adaptive rate coefficient.
(5) Selecting a proper boundary layer parameter of 1 × 10 according to the amplitude of the high-frequency jitter of the star attitude under the PD control action-2And determining a sectional form star error equation:
when | e | ≧ is,
Figure BDA0002472184860000126
when-is < e < is present,
Figure BDA0002472184860000127
wherein f (ω)b) Is a nonlinear term in the kinetic equation.
(6) Calculating the control bandwidth of the star body and the control bandwidth of the load controller, and designing the control parameters k and k2、k3And p, the control bandwidth of the spacecraft star state quantity error after entering the boundary layer is smaller than 1/10 of the bandwidth of the load controller, and meanwhile, the damping ratio of a star control loop is 2, so that the spacecraft triple control robust self-adaptive multi-stage cooperative control is realized.
Example (b):
due to 0<p<1, according to the image of the fal function, the bandwidth of the spacecraft body loop under the action of the controller is continuously increased along with the reduction of the error, and the control bandwidth is kept constant after the spacecraft body loop enters an error boundary layer. In order to ensure that the two-stage compound control system can be sufficiently stable, a control parameter k is designed to be 0.8/2, k2＝2，k30.8/2, p 0.5, adaptive rate parameter 1 × 101The control bandwidth of the system entering the boundary layer is smaller than 1/10 of the bandwidth of the load controller, and the damping ratio of the system is about 2, so that the system can be ensured to be fully stable in the tracking control process, and the overshoot is avoided in the tracking process.
Adopting a star controller and a load controller to carry out multi-level collaborative attitude control on the spacecraft, observing the steady-state response condition of the load after the spacecraft maneuvers 30deg, and respectively giving attitude control results of the load within 0-200 s under the conditions of nonuse and multi-level collaborative control, wherein the convergence time of the load attitude is about 100s before the multi-level collaborative control is applied and the precision of the three-axis attitude is 1.14 × 10 through statistics-3deg，5.30×10-6deg，1.07×10-4deg, three-axis attitude stability of 6.93 × 10-2deg/s，1.50×10-3deg/s，6.43×10-2deg/s, the convergence time of the load attitude is about 10s after the multi-level cooperative control is applied, and the three-axis attitude precision is 1.54 × 10-6deg，1.25×10-7deg，1.85×10-6deg, three-axis attitude stability of 1.47 × 10-4deg/s，1.14×10-5deg/s，1.15×10-3deg/s, load attitude tracking speed, attitude precision and stability are all obviously improved.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.

Claims (11)

1. A spacecraft triple control robust self-adaptive multi-level cooperation method is characterized by comprising the following steps:
(1) establishing a spacecraft system dynamic equation comprising a load dynamic equation, a spacecraft body dynamic equation and a flexible accessory vibration equation;
(2) designing a PD controller aiming at the active pointing hyperstatic platform;
(3) substituting the controller into the load dynamics equation to obtain an error equation of the load attitude:
(4) calculating a proportional control coefficient k from the expected natural frequency according to the calculation formula of the undamped natural frequency of the loadppAnd distributing the control quantity to each actuator;
(5) designing a sliding mode surface, and designing a star robust adaptive controller considering bandwidth constraint and an adaptive law of a nonlinear term based on the sliding mode surface;
(6) selecting boundary layer parameters according to the amplitude of the high-frequency jitter of the star attitude under the PD control action, and determining a star error equation in a segmented form;
(7) calculating the control bandwidth of the star body and the control bandwidth of the load controller, and designing the control parameters k and k2、k3And p, the control bandwidth of the spacecraft star state quantity error after entering the boundary layer is smaller than 1/10 of the bandwidth of the load controller, and meanwhile, the damping ratio of a star control loop is 2, so that the spacecraft triple control robust self-adaptive multi-stage cooperative control is realized.
2. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 1, characterized in that: the spacecraft system comprises a flexible spacecraft body, any number of flexible accessories, a load, a control moment gyro CMG and an active pointing hyperstatic platform; the flexible accessory and the control moment gyroscope CMG are both arranged on the flexible spacecraft body; the active pointing hyperstatic platform is arranged between the load and the flexible spacecraft body, the upper plane of the active pointing hyperstatic platform is connected with the load, and the lower plane of the active pointing hyperstatic platform is connected with the flexible spacecraft body; the active pointing hyperstatic platform consists of six actuators, and each actuator comprises a displacement sensor, a spring-damping structure and a linear motor; the displacement sensor is used for measuring the translational displacement of the linear motor; the spring-damping structure is used for isolating the high-frequency vibration of the flexible spacecraft body; the linear motor is used for providing main power and realizing the attitude control of the load.
3. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 2, wherein: the control moment gyroscope CMG is arranged on the flexible spacecraft body in a distributed mode, so that attitude control of a star body and vibration suppression of the star body structure are achieved.
4. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 1, characterized in that: the load dynamics equation is:
Figure FDA0002472184850000021
wherein the content of the first and second substances,
Figure FDA0002472184850000022
the spacecraft body dynamic equation is as follows:
Figure FDA0002472184850000023
wherein the content of the first and second substances,
Figure FDA0002472184850000024
the flexible attachment vibration equation is:
Figure FDA0002472184850000025
wherein E is3Is an identity matrix of order 3, EmbIs an mb-order identity matrix, rdbη as the relative position vector from the center of mass of the spacecraft body to the center of the lower plane of the active pointing hyperstatic platformbIs a modal coordinate array of the spacecraft body, ηakA modal coordinate array for the kth flexible attachment;
mps，Ips，mbs，Ibs，Pb，Hbsrespectively mass and inertia of a load, total mass and total inertia of a spacecraft platform, modal momentum and modal angular momentum to a body coordinate system, wherein the spacecraft platform comprises a spacecraft body, a flexible attachment and a CMG; fbak、RbakRespectively a translational coupling coefficient and a rotational coupling coefficient of the flexible accessory vibration to the spacecraft body, Λb、ξbRespectively eigenvalue matrix and damping ratio matrix of the spacecraft body, Λak、ξakRespectively, a matrix of eigenvalues and a matrix of damping ratios, h, of the flexible accessoriescIs the total angular momentum, v, of the CMGp、ωp、vb、ωbRespectively the speed and angular velocity of the load, the speed and angular velocity of the spacecraft body, Fd、Td、TcDisturbance force, disturbance torque and control torque borne by the spacecraft body are respectively; t isci、RciRespectively a translational mode matrix and a rotational mode matrix at the mounting point of the ith CMG, Fdi、Tdi、TciDisturbance force, disturbance torque and control torque output by the ith CMG respectively; fujAnd FsjRespectively the action force of the jth actuator on the load and on the spacecraft body, rrpjAnd rrbjRespectively the installation position coordinates of the jth actuator on the load and the spacecraft body.
5. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 4, wherein:
Fdi＝-AbgiAgfiΩi ×Ωi ×AfwiSwi
Tdi＝-rfbi-b ×AbgiAgfiΩi ×Ωi ×AfwiSwi-AbgiAgfiΩi ×Iwi-fΩi
Figure FDA0002472184850000031
Figure FDA0002472184850000032
Fuj＝(-kj(lj-lj0)+uj)suj
Fsj＝-Fuj
wherein omegaiThe rotor speed vector for the ith CMG is at ffiCoordinates under the system, Iwi-fFor the rotor at ffiThe non-zero term of off-diagonal elements of the rotational inertia under the system represents the magnitude of the dynamic unbalance of the CMG rotor, SwiFor the rotor at fwiThe static moment under the system represents the magnitude of the static unbalance of the CMG rotor, and the two terms are disturbance sources of high-frequency micro vibration generated by the CMG; r isfbi-bCoordinates of a relative position vector from a mass center of the body coordinate system to a geometric center of the rotor under the body coordinate system; k is a radical ofjIs the stiffness of the diaphragm spring in the jth actuator, ljIs the length of the jth actuator at any time, lj0Is the length of the initial moment of the j-th actuator, ujIs the control quantity of the j-th actuator, sujThe coordinate of the jth actuator direction vector at any moment in the inertial system;
ffiis defined as follows: f. offiThe system is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor;
fwiis defined as follows: f. ofwiThe system is a rotor fixed connection coordinate system of the ith CMG and rotates along with the rotor;
in the above variables, AabAs a coordinate system fbTo faOf the conversion matrix fa、fbCan be an arbitrary coordinate system and can be used,
Figure FDA0002472184850000041
is obTo oaSagittal diameter of oa、obIs the origin of any two coordinate systems;
subscript e is an inertia system, b is a body system, p is a load coordinate system, gi is a frame coordinate system of the ith CMG, and fi is a rotor geometric coordinate system of the ith CMG and does not rotate along with the rotor; wi is a rotor fixed coordinate system of the ith CMG and rotates along with the rotor;
r×is an antisymmetric matrix of a vector, and the calculation formula is as follows:
Figure FDA0002472184850000042
where r can be any vector.
6. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 5, wherein: the step (2) of designing the PD controller by actively pointing to the hyperstatic platform specifically comprises the following steps:
Figure FDA0002472184850000043
wherein tau isppControl moment, J, output for actively directing the load to the hyperstatic platformpIs Jacobian matrix of load motion space to actuator motion space, and u is column of actuator control quantityThe matrix, corner mark 4:6, takes 4 to 6 rows, ωprAngular velocity, k, of the desired movement for the loadppAnd kpdProportional and differential coefficients, e, of the load controller, respectivelyperrIs the load error attitude angle.
7. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 6, wherein: the error equation of the load attitude in the step (3) is specifically as follows:
Figure FDA0002472184850000044
wherein k isjIs a diagonal array of diaphragm spring stiffness0The length of the actuator and the length array of the initial moment of the actuator are respectively.
8. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 7, wherein: the calculation formula of the undamped natural frequency of the load in the step (4) is as follows:
Figure FDA0002472184850000051
the control amount is distributed to each actuator as follows:
Figure FDA0002472184850000052
9. the spacecraft three-control robust adaptive multi-level collaborative method according to claim 8, wherein: the sliding mode surface in the step (5) is as follows:
Figure FDA0002472184850000053
wherein e ═ eperr T,eη T]T，eηIs of a modal shapeAn error value of the coordinates;
the star robust adaptive controller based on sliding mode surface design and considering bandwidth constraint comprises the following steps:
Figure FDA0002472184850000054
the adaptation law of the nonlinear term is:
Figure FDA0002472184850000055
wherein the content of the first and second substances,
Figure FDA0002472184850000056
Figure FDA0002472184850000057
wherein Hbs0Is HbsEstimated value of k, k2，k3P is the controller gain coefficient, 0<p<1, boundary layer parameters and adaptive law coefficients.
10. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 1, characterized in that: the star error equation in the segmented form in the step (6) is as follows:
when | e | ≧ is,
Figure FDA0002472184850000061
when-is < e < is present,
Figure FDA0002472184850000062
wherein f (ω)b) Is a nonlinear term in the kinetic equation.
11. The spacecraft three-control robust adaptive multi-level collaborative method according to claim 10, wherein: the star control bandwidth is:
Figure FDA0002472184850000063
the control bandwidth of the load controller is the undamped natural frequency of the load, and specifically includes:
Figure FDA0002472184850000064
by designing the control parameters k, k2、k3P, of
Figure FDA0002472184850000065
Damping ratio of star control loop:
Figure FDA0002472184850000066
therefore, three-phase control robust self-adaptive multi-level cooperative control of the spacecraft is realized.
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112015196A (en) * 2020-10-21 2020-12-01 蓝箭航天空间科技股份有限公司 Attitude control system amplitude limiting value design method, storage medium and server
CN114039854A (en) * 2021-10-26 2022-02-11 北京航天科工世纪卫星科技有限公司 Satellite dynamic bandwidth self-adaptive adjusting method based on PID algorithm
CN114408219A (en) * 2022-01-21 2022-04-29 北京控制工程研究所 Reconstruction method for command force distribution array of active pointing hyperstatic platform

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0949143A2 (en) * 1998-04-10 1999-10-13 Laser Communications International, L.L.C. Method and apparatus for improved attitude determination of spacecraft
CN103412491A (en) * 2013-08-27 2013-11-27 北京理工大学 Method for controlling index time-varying slide mode of flexible spacecraft characteristic shaft attitude maneuver
CN106406086A (en) * 2016-05-26 2017-02-15 北京航空航天大学 Large flexible spacecraft interference compensation method based on sliding mode disturbance observer
CN108646775A (en) * 2018-06-08 2018-10-12 北京控制工程研究所 One kind three surpassing the quick motor-driven and fast and stable control method of platform

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0949143A2 (en) * 1998-04-10 1999-10-13 Laser Communications International, L.L.C. Method and apparatus for improved attitude determination of spacecraft
CN103412491A (en) * 2013-08-27 2013-11-27 北京理工大学 Method for controlling index time-varying slide mode of flexible spacecraft characteristic shaft attitude maneuver
CN106406086A (en) * 2016-05-26 2017-02-15 北京航空航天大学 Large flexible spacecraft interference compensation method based on sliding mode disturbance observer
CN108646775A (en) * 2018-06-08 2018-10-12 北京控制工程研究所 One kind three surpassing the quick motor-driven and fast and stable control method of platform

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
郝仁剑等: "基于观测器的超静卫星平台关节-任务空间鲁棒控制方法", 《空间控制技术与应用》 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112015196A (en) * 2020-10-21 2020-12-01 蓝箭航天空间科技股份有限公司 Attitude control system amplitude limiting value design method, storage medium and server
CN114039854A (en) * 2021-10-26 2022-02-11 北京航天科工世纪卫星科技有限公司 Satellite dynamic bandwidth self-adaptive adjusting method based on PID algorithm
CN114408219A (en) * 2022-01-21 2022-04-29 北京控制工程研究所 Reconstruction method for command force distribution array of active pointing hyperstatic platform
CN114408219B (en) * 2022-01-21 2023-06-16 北京控制工程研究所 Method for reconstructing command force distribution array of active pointing hyperstatic platform

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