CN113002803B - Multi-satellite overlength baseline composite formation method - Google Patents
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Abstract
The invention relates to a multi-satellite ultra-long baseline composite formation method, and belongs to the technical field of spaceflight. The satellite body is connected with the load modules through a separated driving platform, the ultra-static ultra-stable satellite is composed of the two load modules and a service module, and ultra-long baseline composite formation is achieved through coarse and fine two-stage formation. Firstly, the satellite body and the load module are electromagnetically locked or actively controlled to realize rigid connection, and the satellite body carries the load module to enter a preset orbit and perform relative motion and pointing control between the satellite bodies. And then, the separated electromagnetic actuator is released and used as an actuating mechanism to generate three-axis control force and three-axis control moment in a matching manner, so that the relative motion and the pointing direction between the load modules are controlled. And finally, releasing the non-dragging checking quality, realizing non-dragging control of the load module through a separated electromagnetic actuator, following the load module by the satellite body through an on-satellite executing mechanism, and simultaneously controlling a plurality of satellites to realize the ultra-long baseline composite formation system.
Description
Technical Field
The invention belongs to the technical field of spaceflight, and relates to a multi-satellite flight formation control method, in particular to a multi-satellite ultra-long baseline composite formation method.
Background
Because the ground gravitational wave experiment is difficult to isolate low-frequency disturbance, the demand for space-based multi-satellite gravitational wave detection is increasing in recent years. The space-based gravitational wave detection is to carry out formation flying on a plurality of satellites at the track height of hundred thousand kilometers, the inter-satellite baseline is hundred thousand to million kilometers, every two satellites form an interferometer, and the distance change between the inspection masses carried by the adjacent satellites is accurately measured through a laser range finder. In order to ensure the laser ranging performance of the ultra-long baseline, the pointing precision between satellites needs to be controlled in the magnitude of nano radian, which has extremely high requirements on the hyperstatic and ultra-stable design and high-precision control of the satellite.
In order to meet the requirements of formation ultra-precise pointing and ultra-static and ultra-stable of the ultra-long baseline space science task, the traditional gravitational wave detection satellite adopts a non-towing control technology, the satellite body is required to be designed ultra-statically and ultra-stably, the method comprises the steps of adopting a high-rigidity structure to prevent the coupling vibration of a flexible accessory, and adopting a micro-Newton micro-thruster for controlling the satellite attitude so as to avoid the micro-vibration of a rotating part, so that the complexity and the difficulty of the satellite system design are increased. Aiming at the application background of multi-satellite high-precision formation with an ultra-long baseline, the invention provides a multi-satellite composite formation method adopting a separated satellite, in order to reduce the design complexity of a satellite platform and further improve the formation control precision. The separated design can isolate the disturbance of the satellite platform, the satellite body does not need to be subjected to ultra-static and ultra-stable design, and the load module can realize ultra-high precision pointing control and ultra-static drag-free control among satellites.
The document 'Chinese invention patent with application publication number CN 107554817A' discloses a satellite composite formation method, aiming at the problem of master-slave type formation satellite attitude and orbit control, two-stage coarse-fine composite formation is carried out respectively through a micro thruster of a satellite body and a separated active vibration isolation and six-degree-of-freedom control system designed between the satellite body and a load, and six degrees of freedom of the satellite body and the load are accurately controlled. The formation mode is only suitable for the fields of distributed optical imaging of short baselines and the like, and is limited in application under the condition of an overlong baseline.
Disclosure of Invention
Technical problem to be solved
The invention provides a multi-satellite ultra-long baseline composite formation method, aiming at overcoming the defects of the existing gravitational wave interference satellite ultra-static ultra-stable formation design and incapable of realizing composite formation under an ultra-long baseline.
Technical scheme
The satellite body and the load modules are connected by adopting a separated driving platform, the hyperstatic hyperstable satellite is composed of the two load modules and a service module, and the ultralong baseline composite formation is realized by coarse-fine two-stage formation. Firstly, the satellite body and the load module are electromagnetically locked or actively controlled to realize rigid connection, and the satellite body carries the load module to enter a preset orbit and perform relative motion and pointing control between the satellite bodies. And then, the separated electromagnetic actuator is released and used as an actuating mechanism to generate three-axis control force and three-axis control moment in a matching manner, so that the relative motion and the pointing direction between the load modules are controlled. And finally, releasing the drag-free inspection quality, realizing drag-free control of the load module through a separated electromagnetic actuator, following the load module by the satellite body through an on-satellite executing mechanism, and simultaneously controlling a plurality of satellites to realize the ultra-long baseline composite formation system.
A multi-satellite overlong baseline composite formation method is characterized by comprising the following steps: the satellite carries two load modules, each load module is provided with a gravitational wave detection mass block, and the satellite body is connected with the two load modules through 8 separated electromagnetic actuators respectively; the separated connecting interface adopts an eight-rod isotropic configuration; the method comprises the following steps:
step 1: in the coarse formation stage, the satellite body is respectively and rigidly connected with the separated electromagnetic actuators between the two load modules; establishing a local reference orbit coordinate system between every two satellites, describing the relative motion of the double-satellite body and the load, and establishing an inter-satellite directional motion model; the local reference orbit coordinate system is as follows:
wherein, Δ r 12 ＝r 1 -r 2 ，r 1 、r 2 Respectively the position vectors, x, of two satellites in an inertial frame F 、y F 、z F Three axes of a local reference orbit coordinate system are respectively, and the three axes follow a right-hand rule;
and 2, step: taking a formation satellite operating in a gravity field at the center of the earth as an example, an orbital dynamics equation of a satellite i in an orbit coordinate system at the center of the earth is as follows:
wherein r is i 、v i Respectively the displacement and velocity vectors of the satellite i, d i Is subject to disturbance;
the attitude kinetic equation of the satellite i is
Wherein, T ic 、T id Respectively the control moment and the disturbance moment borne by the satellite body;
the satellite relative kinematics equation can be obtained by converting a relative coordinate system, wherein the relative change rate of the 6-degree-of-freedom vector can be obtained by an equation (4); the rate of change of the known vector with respect to the fixed reference coordinate is equal to the sum of the rate of change of the vector in the dynamic coordinate and the product of the rotation speed vector ω of the dynamic coordinate with respect to the reference coordinate and x of the vector:
by measuring the relative position and the relative attitude among the formation satellites, controlling the relative motion among the formation satellite bodies according to the kinematics and the dynamics model obtained by derivation and a conventional pose control method, determining a local reference orbit coordinate system, and controlling the pointing direction of the satellites to preliminarily track the origin of the local reference orbit coordinate system;
and step 3: when the relative motion of the satellite body is stable and the inter-satellite coarse pointing accuracy meets the requirement of a load module working domain, the separated electromagnetic actuator starts to work and enters a fine formation stage; the separated electromagnetic actuator is unlocked, the load module and the satellite body do not maintain a rigid connection state any more, the satellite body and the two load modules are converted from the connection state to the separation state, at the moment, a micro-vibration transmission path from the satellite platform to the load module is physically isolated, and the load module realizes hyperstatic hyperstability; two load modules of the satellite i point to load modules of adjacent satellites respectively;
calculating by a measuring system to obtain relative attitude deviation between different load modules of adjacent satellites and a local orbit coordinate system of the adjacent satellites, and controlling the attitude of the load module to track the local orbit coordinate system by a separated electromagnetic actuator;
according to the structural layout mode of the separated electromagnetic actuators, the separated electromagnetic actuators output forces at different positions of the load module, and finally control forces and control moments of various degrees of freedom are obtained;
the dynamic equation of the load module is
Wherein, χ B ＝(x B ,y B ,z B ,α B ,β B ,γ B ) T Is a generalized coordinate of the satellite body, χ P ＝(x P ,y P ,z P ,α P ,β P ,γ P ) T Generalized coordinates, M, for separate load modules P C and K respectively represent an equivalent damping matrix and an equivalent stiffness matrix of the actuator, wherein subscripts B and P respectively represent a satellite body and a load module; the relative motion between the load modules is obtained by transforming the relative coordinate system through the absolute motion equation of the load modules, F ic 、T ic Respectively a control force and a control moment, F iPd 、T iPd Respectively a disturbance force and a disturbance torque; is provided with
Wherein F is a vector formed by output forces of 8 separate electromagnetic actuators; the relationship between the output force F and the current I of the single split electromagnetic actuator is given by the following formula:
F＝kBLI (7)
wherein L is the effective length of the wire, B is the magnetic field intensity, and k is the inductive coefficient of the wire;
and 4, step 4: after the motion and the inter-satellite pointing of the load modules meet gravitational wave measurement conditions, unlocking the detection mass block, measuring the non-gravitational acceleration borne by the load modules by adopting a displacement mode or an acceleration mode, decoupling the six-degree-of-freedom motion of the load modules, selecting 6 degrees of freedom as sensitive degrees of freedom for 12 degrees of freedom of the two load modules, and selecting the rest 6 degrees of freedom as non-sensitive degrees of freedom; at the moment, the control of the load module comprises sensitive freedom degree drag-free control, direction control and insensitive freedom degree stable control; the method comprises the steps that drag-free control and inter-satellite pointing control are simultaneously carried out on the sensitive freedom degree of a load module through a separated electromagnetic actuator, the drag-free control is used for following detection quality, and non-gravitational acceleration borne by the load module in a scientific task measurement frequency range is eliminated; the pointing control is used for controlling the pointing between load modules of adjacent satellites in a non-measurement frequency band so as to ensure the laser measurement precision; the 6 non-sensitive degrees of freedom of the load module are stably controlled through a separated electromagnetic actuator; the load module realizes the ultra-static environment and ultra-high precision inter-satellite pointing required by scientific detection under the simultaneous action of non-dragging control, pointing control and stable control; the satellite body tracks the two load modules through a conventional pose control method, relative motion between the satellite body and the load modules is kept and controlled, and collision between the satellite body and the load modules is avoided, so that a measurement result is prevented from being interfered; and a plurality of satellites form an ultra-long baseline composite formation system.
The technical scheme of the invention is further that: the separated electromagnetic actuator selects a voice coil actuator, the output of force is realized by introducing direct current into a coil, and the magnitude and the direction of output force are changed by adjusting the magnitude and the direction of input current.
The technical scheme of the invention is further that: step 1, the rigid connection mode of the separated electromagnetic actuator between the satellite body and the two load modules is electromagnetic locking or active control.
The technical scheme of the invention is further described as follows: and the active cells and the stators of the 8 separated electromagnetic actuators are respectively connected with the satellite body and the load module through bolts.
Advantageous effects
The invention provides a multi-satellite ultra-long baseline composite formation method, which is characterized in that a separate active vibration isolation and six-degree-of-freedom control system is designed between bodies and load modules of a plurality of satellites, the bodies and the load modules of the satellites are electromagnetically locked or actively controlled to realize rigid connection, a conventional method is adopted to carry out coarse formation and coarse pointing on the plurality of satellites, relative motion between the bodies of the satellites and a local reference orbit coordinate system are derived according to an absolute motion equation aiming at the problem of controlling the attitude and the orbit of the ultra-long baseline formation satellites, and the relative motion between the bodies of the satellites and the inter-satellite pointing are controlled by the conventional position and pose control method. And then releasing the separated system, deducing and obtaining relative motion and pointing information among the load modules according to an absolute motion equation of the load modules, generating control force and control moment by using the cooperation of 8 separated electromagnetic actuators, stably controlling the three-degree-of-freedom displacement of the load, and controlling the posture of the load modules to realize high-precision pointing among the satellites. And by combining a non-dragging control technology, the load module is subjected to non-dragging and inter-satellite pointing coordination control, the satellite body tracks the motion of the load module, the requirements of an ultra-static environment and ultra-high precision inter-satellite pointing required by scientific detection are met, and a multi-satellite is controlled to realize an ultra-long baseline composite formation system.
A separated active vibration isolation and six-degree-of-freedom control system is designed between a satellite body and a load module, a separated interface can isolate interference of various micro-vibrations of the satellite body on a precise load from a physical full frequency band, and the satellite body can adopt a traditional satellite design, so that the difficulty and complexity of the satellite design are reduced.
The separated active vibration isolation and six-degree-of-freedom control system can provide ultra-precise load module control, secondary formation stability and pointing control are carried out on the load on the basis of coarse formation and coarse pointing of the satellite body, formation and pointing precision are further improved, and the ultra-precise pointing requirement of scientific detection is met.
The separated interface can provide an ultra-static environment for the load module, and further meets the ultra-static requirement of scientific detection by combining non-dragging control.
The invention can break through the limitation of the base line, and multi-star can carry out composite formation on the base line of hundred thousand to million kilometers.
Drawings
FIG. 1 is a flow chart of a multi-satellite overlength baseline composite formation method of the present invention.
FIG. 2 is a schematic diagram of a multi-satellite overlength baseline composite formation method of the present invention.
FIG. 3 is a schematic diagram of a separate active vibration isolation and six-degree-of-freedom control system in the method of the present invention.
FIG. 4 is a top view of the split active vibration isolation and six-degree-of-freedom control system of the method of the present invention.
FIG. 5 is a block diagram of coordination control of no-drag and inter-satellite pointing of the loading platform in the method of the present invention.
In the figure, 1-satellite body, 2-first load module, 3-second load module, 4-proof mass block, 5-load module connection interface, 6-satellite body connection interface, 7-first separated electromagnetic actuator, 8-second separated electromagnetic actuator, 9-third separated electromagnetic actuator, 10-fourth separated electromagnetic actuator, 11-fifth separated electromagnetic actuator, 12-sixth separated electromagnetic actuator, 13-seventh separated electromagnetic actuator, 14-eighth separated electromagnetic actuator.
Detailed Description
The invention will now be further described with reference to the following examples, and the accompanying drawings:
reference is made to fig. 1-5.
Step one, designing a separated active vibration isolation and six-degree-of-freedom control system. One satellite carries two load modules, and the two load modules comprise two separated active vibration isolation and six-degree-of-freedom control systems, each system consists of 8 separated electromagnetic actuators, and the system is shown in fig. 3. Each load module is provided with a gravitational wave detection mass block, and the satellite body is connected with the two load modules through 8 separated electromagnetic actuators respectively. The active cells and the stators of the 8 separated electromagnetic actuators are connected with the satellite body and the load module through bolts respectively. The split connection interface adopts an eight-bar isotropic configuration.
And step two, performing coarse formation relative motion of the satellite bodies, and performing pointing modeling and coarse control among satellites. In the coarse formation stage, the separated electromagnetic actuators between the satellite body and the two load modules are respectively electromagnetically locked or actively controlled to realize rigid connection. A local reference orbit coordinate system is established between every two satellites, relative motion of the double-satellite body and the load is described, and an inter-satellite pointing motion model is established. The local reference orbit coordinate system is obtained by the following formula:
wherein, Δ r 12 ＝r 1 -r 2 ，r 1 、r 2 Are respectively the position vectors, x, of two satellites in an inertial coordinate system F 、y F 、z F Three axes of the local reference orbital coordinate system, respectively, follow the right-hand rule.
Taking the formation satellites operating in the earth center gravitational field as an example, the orbital dynamics equation of the satellite i in the earth center orbit coordinate system is as follows:
wherein r is i 、v i Respectively displacement and velocity vector of satellite i, d i It is disturbed.
The attitude dynamics equation of the satellite i is
Wherein, T ic 、T id Respectively the control moment and the disturbance moment borne by the satellite body.
The satellite relative kinematics equation can be obtained by transformation of a relative coordinate system. The relative rate of change of the 6-degree-of-freedom vector can be obtained from equation (4). The rate of change of the known vector with respect to the fixed reference coordinate is equal to the sum of the rate of change of the vector at the moving coordinate and the product of the rotation speed vector ω of the moving coordinate with respect to the reference coordinate and x of the vector:
by measuring the relative position and the relative attitude among the formation satellites, the kinematics and the dynamics model obtained by the derivation control the relative motion among the formation satellite bodies by a conventional pose control method, a local reference orbit coordinate system is determined, and the pointing direction of the satellites is controlled to primarily track the origin of the local reference orbit coordinate system.
And fourthly, accurately forming relative motion and inter-satellite pointing modeling and accurately controlling the load modules. When the relative motion of the satellite body is stable and the inter-satellite coarse pointing precision meets the requirement of a load module working domain, the separated electromagnetic actuator starts to work and enters a fine formation stage. The separating electromagnetic actuator unlocks or does not control the length of each supporting rod to be unchanged any more, the satellite body and the two load modules are converted into a separating state from a connecting state, the micro-vibration transmission path from the satellite platform to the load modules is physically isolated at the moment, and the load modules realize hyperstatic and hyperstable effects. The two load modules of satellite i are each directed to a load module of an adjacent satellite.
The relative attitude deviation between different load modules of adjacent satellites and a local orbit coordinate system of the adjacent satellites is obtained through calculation of a measuring system, the attitude of the load module is controlled by a separated voice coil actuator to track the local orbit coordinate system, high-precision pointing between the satellites is achieved, and three-degree-of-freedom displacement motion stability of the load module is kept.
The separated electromagnetic actuator can be a voice coil actuator generally, the output of force is realized by introducing direct current into a coil, and the magnitude and the direction of output force are changed by adjusting the magnitude and the direction of input current. And (4) outputting force by each actuator at different positions of the load module according to the structural layout mode of the actuators in the step one, and finally obtaining the control force and the control moment of each degree of freedom.
The dynamic equation of the load module is
Wherein, χ B ＝(x B ,y B ,z B ,α B ,β B ,γ B ) T Is a generalized coordinate of the satellite body, χ P ＝(x P ,y P ,z P ,α P ,β P ,γ P ) T As a generalized coordinate of the separate load module, M P And C and K represent equivalent damping and equivalent stiffness matrixes of the actuator respectively, wherein subscripts B and P represent the satellite body and the load module respectively. The relative motion between the load modules is obtained by transforming the relative coordinate system through the absolute motion equation of the load modules, F ic 、T ic Respectively a control force and a control moment, F iPd 、T iPd Respectively disturbance force and disturbance torque. Is provided with
Wherein, F is the vector that 8 voice coil actuator output power are constituteed. The magnitude of the output force F of a single voice coil actuator is related to the current I by the following equation:
F＝kBLI (7)
wherein, L is the effective length of the wire, B is the magnetic field intensity, and k is the inductive coefficient of the wire.
And step five, controlling the load module without dragging and pointing and controlling the satellite body to follow. After the motion and the inter-satellite pointing of the load module meet the gravitational wave measurement condition, the detection mass block is unlocked, the non-gravitational acceleration borne by the load module is measured in a displacement mode or an acceleration mode, the six-degree-of-freedom motion of the load module is decoupled, 6 degrees of freedom are selected as sensitive degrees of freedom for 12 degrees of freedom of the two load modules, and the rest 6 degrees of freedom are non-sensitive degrees of freedom. At the moment, the control of the load module comprises sensitive freedom degree drag-free control, direction control and non-sensitive freedom degree stable control. The method comprises the steps that drag-free control and inter-satellite pointing control are simultaneously carried out on the sensitive freedom degree of a load module through a separated electromagnetic actuator, the drag-free control is used for following detection quality, and non-gravitational acceleration borne by the load module in a scientific task measurement frequency range is eliminated; and the pointing control is used for controlling the pointing between the load modules of the adjacent satellites in a non-measurement frequency band, so that the laser measurement precision is ensured. The 6 insensitive degrees of freedom of the load module are stably controlled by a separated electromagnetic actuator. The load module realizes the ultra-static environment and ultra-high precision inter-satellite pointing required by scientific detection under the simultaneous action of non-dragging control, pointing control and stable control. The satellite body tracks the two load modules through a conventional pose control method, relative motion between the satellite body and the load modules is kept and controlled, and collision between the satellite body and the load modules is avoided, so that a measurement result is prevented from being interfered.
Claims (4)
1. A multi-satellite overlong baseline composite formation method is characterized by comprising the following steps: the satellite carries two load modules, each load module is provided with a gravitational wave detection mass block, and the satellite body is connected with the two load modules through 8 separated electromagnetic actuators respectively; the separated connecting interface adopts an eight-rod isotropic configuration; the method comprises the following steps:
step 1: in the coarse formation stage, the satellite body is respectively and rigidly connected with the separated electromagnetic actuators between the two load modules; establishing a local reference orbit coordinate system between every two satellites, describing the relative motion of the double-satellite body and the load, and establishing an inter-satellite directional motion model; the local reference orbit coordinate system is as follows:
wherein, Δ r 12 ＝r 1 -r 2 ，r 1 、r 2 Respectively the position vectors, x, of two satellites in an inertial frame F 、y F 、z F Three axes of a local reference orbit coordinate system are respectively, and the three axes follow a right-hand rule;
step 2: taking a formation satellite operating in an earth center gravitational field as an example, an orbital dynamics equation of a satellite i in an earth center orbital coordinate system is as follows:
wherein r is i 、v i Position and velocity vectors, d, of satellite i, respectively i Is subject to disturbance;
the attitude dynamics equation of the satellite i is
Wherein, T ic 、T id Respectively the control moment and the disturbance moment borne by the satellite body;
the equations of the satellite relative kinematics can be obtained by transformation of a relative coordinate system, wherein,the relative rate of change of the 6 degree-of-freedom vectors can be obtained from equation (4); the rate of change of the known vector relative to a fixed reference coordinate is equal to the rate of change of the vector in the moving coordinate and the speed vector omega of the moving coordinate relative to the reference coordinate i Sum of cross products with this vector:
by measuring the relative position and the relative attitude among the formation satellites, controlling the relative motion among the formation satellite bodies according to the kinematics and the dynamics model obtained by derivation and a conventional pose control method, determining a local reference orbit coordinate system, and controlling the pointing direction of the satellites to preliminarily track the origin of the local reference orbit coordinate system;
and 3, step 3: when the relative motion of the satellite body is stable and the inter-satellite coarse pointing accuracy meets the requirement of a load module working domain, the separated electromagnetic actuator starts to work and enters a fine formation stage; the separated electromagnetic actuator is unlocked, the load module and the satellite body do not maintain a rigid connection state any more, the satellite body and the two load modules are converted from the connection state to the separation state, at the moment, a micro-vibration transmission path from the satellite platform to the load module is physically isolated, and the load module realizes hyperstatic hyperstability; two load modules of the satellite i point to load modules of adjacent satellites respectively;
calculating by a measuring system to obtain the relative attitude deviation between different load modules of adjacent satellites and a local orbit coordinate system of the adjacent satellites, and controlling the attitude of the load modules to track the local orbit coordinate system by a separated electromagnetic actuator;
according to the structural layout mode of the separated electromagnetic actuators, the separated electromagnetic actuators output forces at different positions of the load module, and finally control forces and control moments of various degrees of freedom are obtained;
the dynamic equation of the load module is
Wherein, χ B ＝(x B ,y B ,z B ,α B ,β B ,γ B ) T Is a generalized coordinate of the satellite body, χ P ＝(x P ,y P ,z P ,α P ,β P ,γ P ) T As a generalized coordinate of the separate load module, M P C and K respectively represent equivalent damping and equivalent stiffness matrixes of the actuator, wherein subscripts B and P respectively represent the satellite body and the load module; the relative motion between the load modules is obtained by transforming the relative coordinate system through the absolute motion equation of the load modules, F ic 、T ic Respectively a control force and a control moment, F iPd 、T iPd Respectively a disturbance force and a disturbance torque; is provided with
Wherein F is a vector formed by output forces of 8 separate electromagnetic actuators; the output force magnitude F of the single split electromagnetic actuator is related to the current I by the following equation:
F＝kBLI (7)
wherein L is the effective length of the wire, B is the magnetic field intensity, and k is the inductive coefficient of the wire;
and 4, step 4: after the motion and the inter-satellite pointing of the load modules meet gravitational wave measurement conditions, unlocking the detection mass block, measuring the non-gravitational acceleration borne by the load modules by adopting a displacement mode or an acceleration mode, decoupling the six-degree-of-freedom motion of the load modules, selecting 6 degrees of freedom as sensitive degrees of freedom for 12 degrees of freedom of the two load modules, and selecting the rest 6 degrees of freedom as non-sensitive degrees of freedom; at the moment, the control of the load module is divided into sensitive freedom degree drag-free control and pointing control and non-sensitive freedom degree stable control; the method comprises the following steps of simultaneously carrying out non-dragging control and inter-satellite pointing control on the sensitive degree of freedom of a load module through a separated electromagnetic actuator, wherein the non-dragging control is used for following the detection quality and eliminating the non-gravitational acceleration borne by the load module in the scientific task measurement frequency band; the pointing control is used for controlling the pointing between load modules of adjacent satellites in a non-measurement frequency band so as to ensure the laser measurement precision; the 6 non-sensitive degrees of freedom of the load module are stably controlled through a separated electromagnetic actuator; the load module realizes the ultra-static environment and ultra-high precision inter-satellite pointing required by scientific detection under the simultaneous action of non-dragging control, pointing control and stable control; the satellite body tracks the two load modules through a conventional pose control method, relative motion between the satellite body and the load modules is kept and controlled, and collision between the satellite body and the load modules is avoided, so that a measurement result is prevented from being interfered; and a plurality of satellites form an ultra-long baseline composite formation system.
2. The multi-satellite overlength baseline composite formation method according to claim 1, wherein the separate electromagnetic actuator is a voice coil actuator, force output is realized by passing direct current through a coil, and the magnitude and direction of output force are changed by adjusting the magnitude and direction of input current.
3. The multi-satellite ultra-long baseline composite formation method according to claim 1, wherein in the step 1, the rigid connection mode of the separated electromagnetic actuators between the satellite body and the two load modules is electromagnetic locking or active control.
4. The multi-satellite overlength baseline composite formation method according to claim 1, wherein the active cells and the stators of the 8 separate electromagnetic actuators are respectively connected with the satellite body and the load module through bolts.
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| CN110147115A (en) * | 2019-06-21 | 2019-08-20 | 哈尔滨工业大学 | Centered on load, the spin load satellite attitude control method that platform is servo-actuated |
| CN112506211A (en) * | 2020-12-07 | 2021-03-16 | 上海卫星工程研究所 | Future gravitational field measurement oriented separation type satellite platform drag-free control method and system |
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