CN113002803B - Multi-star and ultra-long baseline compound formation method - Google Patents

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

Multi-star and ultra-long baseline compound formation method

technical field

The invention belongs to the field of aerospace technology, and relates to a multi-star flight formation control method, in particular to a multi-star ultra-long baseline composite formation method.

Background technique

Due to the difficulty of isolating low-frequency disturbances in ground-based gravitational wave experiments, the demand for space-based multi-star gravitational wave detection has been increasing in recent years. Space-based gravitational wave detection is a formation flight of multiple stars at an orbital altitude of 100,000 kilometers. The baseline between stars is between 100,000 and millions of kilometers. Two satellites form an interferometer, and the laser rangefinder is used to accurately measure the satellites carried by adjacent satellites. The distance between the inspection masses varies. In order to ensure the laser ranging performance of the ultra-long baseline, the inter-satellite pointing accuracy needs to be controlled at the nanoradian level, which requires extremely high requirements for the ultra-static and ultra-stable design and high-precision control of the satellite.

In order to meet the ultra-precise pointing and ultra-static and ultra-stable requirements of formations for such ultra-long baseline space science missions, traditional gravitational wave detection satellites use drag-free control technology, and the satellite body requires ultra-static and ultra-stable design, including the use of high-rigidity structures to To prevent the coupling vibration of the flexible accessories and avoid the micro-vibration of the rotating parts, the satellite attitude control is controlled by a Webull-level micro-thrust, which increases the complexity and difficulty of the satellite system design. In order to reduce the design complexity of the satellite platform and further improve the formation control accuracy, the present invention proposes a multi-satellite composite formation method using separate satellites for the application background of the ultra-long baseline multi-satellite high-precision formation. The separate design can isolate the disturbance of the satellite platform, and the satellite body does not need to be designed for ultra-static and ultra-stable.

The document "Chinese invention patent with the application publication number of CN107554817A" discloses a satellite composite formation method, aiming at the problem of attitude and orbit control of the master-slave formation satellite, respectively through the micro thruster of the satellite body and the design between the satellite body and the load. The separate active vibration isolation and six-degree-of-freedom control system conducts a two-stage rough-fine composite formation to precisely control the six-degree-of-freedom of the satellite body and the payload. This formation method is only suitable for short-baseline distributed optical imaging and other fields, and its application under ultra-long baselines is limited.

SUMMARY OF THE INVENTION

technical problem to be solved

In order to overcome the deficiencies in the ultra-quiet and ultra-stable formation design of the existing gravitational wave interference satellites, which cannot realize the composite formation under the ultra-long baseline, the invention provides a multi-satellite ultra-long baseline composite formation method.

Technical solutions

The satellite body and the load module of the present invention are connected by a separate drive platform, two load modules and one service module form an ultra-static and ultra-stable satellite, and the ultra-long baseline composite formation is realized through two-level formation of rough and fine. First, the satellite body and the load module are electromagnetically locked or actively controlled to realize a rigid connection. The satellite body carries the load module into a predetermined orbit and performs relative motion and pointing control between the satellite bodies. After that, the separate electromagnetic actuator is released, and cooperates as an actuator to generate a three-axis control force and a three-axis control torque to control the relative movement and orientation of the load modules. Finally, the non-drag inspection mass is released. The load module realizes non-drag control through the separate electromagnetic actuator. The satellite body follows the load module through the on-board actuator, and multiple satellites perform the above control at the same time to realize the ultra-long baseline composite formation system.

A multi-satellite ultra-long baseline composite formation method, which is characterized in that: a satellite is equipped with two load modules, each load module is equipped with a gravitational wave detection mass block, and the satellite body and the two load modules are separated by eight separations respectively. connected with electromagnetic actuators; the separate connection interface adopts an eight-bar isotropic configuration; the steps are as follows:

Step 1: In the rough formation stage, the separate electromagnetic actuators between the satellite body and the two load modules are rigidly connected respectively; a local reference orbit coordinate system is established between the satellites to describe the relative motion of the double star body and the payload, and establish an inter-satellite coordinate system. Pointing to the kinematic model; the local reference orbit coordinate system is:

Among them, Δr ₁₂ =r ₁ -r ₂ , r ₁ and r ₂ are the position vectors of the two satellites in the inertial coordinate system, respectively, x _F , y _F , and z _F are the three axes of the local reference orbit coordinate system, respectively. The axis obeys the right-hand rule;

Step 2: Taking the formation satellite running in the gravitational field at the center of the earth as an example, the orbital dynamics equation of satellite i in the coordinate system of the earth's center orbit is:

Among them, ri and vi are the displacement and velocity vector of satellite _i , respectively _{, and d i is} the disturbance;

The attitude dynamics equation of satellite i is:

Among them, T _ic and T _id are the control torque and disturbance torque of the satellite body, respectively;

The relative kinematics equation of the satellite can be obtained through the transformation of the relative coordinate system, in which the relative change rate of the 6-DOF vector can be obtained by formula (4); the change rate of the known vector relative to the fixed reference coordinate is equal to the change rate of the vector in the moving coordinate and The sum of the speed vector ω of the moving coordinate relative to the reference coordinate and the product of this vector:

By measuring the relative positions and relative attitudes of the formation satellites, the kinematics and dynamics models derived from the above are used to control the relative motion between the formation satellites through the conventional pose control method, determine the local reference orbit coordinate system, and control the pointing of the satellites Preliminarily track the origin of the local reference orbit coordinate system;

Step 3: When the relative motion of the satellite body is stable and the coarse pointing accuracy between the satellites meets the requirements of the working domain of the load module, the separate electromagnetic actuator starts to work and enters the fine formation stage; the separate electromagnetic actuator is unlocked, and the load module and the satellite body are unlocked. The rigid connection state is no longer maintained, and the satellite body and the two load modules are converted from the connected state to the separated state. At this time, the micro-vibration transmission path from the satellite platform to the load module is physically isolated, and the load module realizes ultra-static and ultra-stable; The two load modules of i point to the load modules of the adjacent satellites respectively;

The relative attitude deviations of different load modules of adjacent satellites and their local orbit coordinate systems are obtained through the calculation of the measurement system, and the attitude of the load modules is controlled by a separate electromagnetic actuator to track the local orbit coordinate system;

According to the structure and layout of the separate electromagnetic actuators, each separate electromagnetic actuator outputs force at different positions of the load module, and finally obtains the control force and control torque of each degree of freedom;

The dynamic equation of the load module is

Among them, χ _B = (x _B , y _B , z _B , α _B , β _B , γ _B ) ^T is the generalized coordinate of the satellite ontology, χ _P = (x _P , y _P , z _P , α _P , β _P ,γ _P ) ^T is the generalized coordinate of the separate load module, M _P is the inertia matrix of the load module, C and K represent the equivalent damping and equivalent stiffness matrices of the actuator, respectively, where the subscripts B and P represent satellites, respectively Body and load module; the relative motion between load modules is obtained by transforming the relative coordinate system of the absolute motion equation of the load module, F _ic and T _ic are the control force and control torque, respectively, F _iPd , T _iPd are the disturbance force and disturbance torque, respectively ;Have

Among them, F is the vector composed of the output forces of 8 separate electromagnetic actuators; the relationship between the output force F of a single separate electromagnetic actuator and the current I is given by the following formula:

F=kBLI (7)

Among them, L is the effective length of the wire, B is the magnetic field strength, and k is the inductive coefficient of the wire;

Step 4: After the motion of the load module and the direction between the stars meet the gravitational wave measurement conditions, unlock the detection mass, measure the non-gravitational acceleration of the load module in the displacement mode or the acceleration mode, and decouple the six-degree-of-freedom motion of the load module , select 6 degrees of freedom as sensitive degrees of freedom for a total of 12 degrees of freedom of the two load modules, and the remaining 6 degrees of freedom as non-sensitive degrees of freedom; at this time, the control of the load module is divided into sensitive degrees of freedom without drag control and pointing control, Stable control of non-sensitive degrees of freedom; non-drag control and inter-satellite pointing control are simultaneously performed for the sensitive degrees of freedom of the load module through separate electromagnetic actuators. The non-gravitational acceleration; the pointing control is used to control the pointing between the load modules of adjacent satellites in the non-measurement frequency band to ensure the accuracy of laser measurement; the six non-sensitive degrees of freedom of the load module are stably controlled by separate electromagnetic actuators; Under the simultaneous action of non-drag control, pointing control and stability control, the payload module realizes the ultra-static environment and ultra-high-precision inter-satellite pointing required for scientific exploration; the satellite body tracks the two payload modules through the conventional attitude and attitude control method, and conducts satellite tracking. The relative movement between the body and the load module is kept under control to avoid collision with the load module and interfere with the measurement results; multiple satellites form an ultra-long baseline composite formation system.

The technical solution of the present invention further states that the separate electromagnetic actuator uses a voice coil actuator, which realizes the output of force by passing a DC current into the coil, and changes the output force by adjusting the magnitude and direction of the input current. size and orientation.

The technical solution of the present invention further states that in step 1, the rigid connection mode of the separate electromagnetic actuator between the satellite body and the two load modules is electromagnetic locking or active control.

The technical solution of the present invention further states that the movers and stators of the eight separate electromagnetic actuators are respectively connected with the satellite body and the load module through bolts.

beneficial effect

The invention proposes a multi-satellite ultra-long baseline composite formation method. The method designs a separate active vibration isolation and six-degree-of-freedom control system between the bodies and load modules of multiple satellites. Carry out electromagnetic locking or active control to achieve rigid connection, and use conventional methods to perform rough formation and rough pointing between multiple satellites. For the attitude and orbit control problems of ultra-long baseline formation satellites, the relative relationship between satellite bodies is derived according to the absolute motion equation. The motion and the local reference orbit coordinate system are used to control the relative motion between satellite bodies and the inter-satellite pointing using conventional pose control methods. After that, the separate system is released, and the relative motion and orientation information between the load modules are derived according to the absolute motion equation of the load modules, and 8 separate electromagnetic actuators are used to cooperate to generate the control force and control torque to stabilize the three-degree-of-freedom displacement of the load. Control the attitude of the payload module to achieve high-precision pointing between satellites. Combined with the no-drag control technology, the payload module is controlled without drag and inter-satellite pointing, and the satellite body tracks the motion of the payload module to meet the ultra-static environment and ultra-high-precision inter-satellite pointing requirements required for scientific exploration, and to control multiple satellites to achieve ultra-high accuracy. Long Baseline Compound Formation System.

A separate active vibration isolation and six-degree-of-freedom control system is designed between the satellite body and the load module. The separate interface can physically isolate the interference of various micro-vibrations of the satellite body to the precision load in the full frequency band. The satellite body can adopt traditional satellite design. , reduce the difficulty and complexity of satellite design.

The separate active vibration isolation and six-degree-of-freedom control system can provide ultra-precise load module control, perform secondary formation stabilization and pointing control of the load on the basis of the rough formation and rough pointing of the satellite body, and further improve the formation and pointing accuracy. The ultra-precise pointing requirement of scientific exploration.

The separated interface can provide a hyperstatic environment for the load module, and combined with the drag-free control, it can further meet the hyperstatic requirements of scientific exploration.

The invention can break through the limitation of the baseline, and multiple satellites can form a composite formation on the baseline of the order of 100,000 to one million kilometers.

Description of drawings

FIG. 1 is a flow chart of the multi-star and ultra-long baseline composite formation method of the present invention.

FIG. 2 is a schematic diagram of the multi-star and ultra-long baseline composite formation method of the present invention.

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.

4 is a top view of a separate active vibration isolation and six-degree-of-freedom control system in the method of the present invention.

FIG. 5 is a block diagram of the coordinated control of the load platform without drag and the inter-satellite pointing in the method of the present invention.

In the figure, 1-satellite body, 2-first load module, 3-second load module, 4-detection mass, 5-load module connection interface, 6-satellite body connection interface, 7-first separate electromagnetic actuator Actuator, 8-second separate electromagnetic actuator, 9-third separate electromagnetic actuator, 10-fourth separate electromagnetic actuator, 11-fifth separate electromagnetic actuator, 12-th Six separate electromagnetic actuators, 13 - seventh separate electromagnetic actuator, 14 - eighth separate electromagnetic actuator.

Detailed ways

The present invention will now be further described in conjunction with the embodiments and accompanying drawings:

Refer to Figures 1-5.

Step 1. Design of separate active vibration isolation and six-degree-of-freedom control system. A satellite is equipped with two load modules, including two separate active vibration isolation and six-degree-of-freedom control systems. The system consists of eight separate electromagnetic actuators. The system is shown in Figure 3. Each load module is equipped with a gravitational wave detection mass block, and the satellite body and the two load modules are respectively connected by 8 separate electromagnetic actuators. The movers and stators of the eight separate electromagnetic actuators are respectively connected with the satellite body and the load module through bolts. The split connection interface adopts an eight-bar isotropic configuration.

Step 2: Modeling and rough control of relative motion of satellite body rough formation and inter-satellite pointing. In the rough formation stage, the separate electromagnetic actuators between the satellite body and the two load modules are respectively electromagnetically locked or actively controlled to achieve a rigid connection. A local reference orbit coordinate system is established between the satellites to describe the relative motion of the binary star itself and its payload, and a model of inter-satellite pointing motion is established. The local reference orbit coordinate system is obtained by the following formula:

Among them, Δr ₁₂ =r ₁ -r ₂ , r ₁ and r ₂ are the position vectors of the two satellites in the inertial coordinate system, respectively, x _F , y _F , and z _F are the three axes of the local reference orbit coordinate system, respectively. Axes obey the right-hand rule.

Step 3. Taking the formation satellite running in the gravitational field of the center of the earth as an example, the orbital dynamics equation of satellite i in the coordinate system of the earth's center orbit is:

Among them, ri and vi are the displacement and velocity vector of satellite _i , respectively _{, and d i is} the disturbance.

The attitude dynamics equation of satellite i is:

Among them, T _ic and T _id are the control torque and disturbance torque of the satellite body, respectively.

The satellite relative kinematics equation can be obtained by transforming the relative coordinate system. Among them, the relative change rate of the 6-DOF vector can be obtained by equation (4). The rate of change of the known vector relative to the fixed reference coordinate is equal to the sum of the rate of change of the vector in the moving coordinate and the product of the speed vector ω of the moving coordinate relative to the reference coordinate and this vector:

By measuring the relative positions and relative attitudes of the formation satellites, the kinematics and dynamics models derived from the above are used to control the relative motion between the formation satellites through the conventional pose control method, determine the local reference orbit coordinate system, and control the pointing of the satellites Initially track the origin of the local reference orbit coordinate system.

Step 4: Modeling and precise control of the relative motion and inter-satellite pointing of the fine formation of the load module. When the relative motion of the satellite body is stable and the coarse pointing accuracy between the satellites meets the requirements of the working domain of the load module, the separate electromagnetic actuator starts to work and enters the fine formation stage. The separate electromagnetic actuator is unlocked or the length of each strut is no longer controlled, and the satellite body and the two load modules are converted from the connected state to the separated state. At this time, the micro-vibration transmission path from the satellite platform to the load module is physically blocked. Isolation, the load module achieves ultra-static and ultra-stable. The two payload modules of satellite i respectively point to the payload modules of the adjacent satellites.

The relative attitude deviation of different load modules of adjacent satellites and their local orbital coordinate system is obtained through the calculation of the measurement system, and the attitude of the load module is controlled by the separate voice coil actuator to track the local orbital coordinate system to achieve high-precision pointing between satellites. And keep the three-degree-of-freedom displacement motion of the load module stable.

Separable electromagnetic actuators generally use voice coil actuators, which can realize the output of force by passing a DC current into the coil, and change the magnitude and direction of the output force by adjusting the magnitude and direction of the input current. According to the structure and layout of the actuators in step 1, each actuator outputs force at different positions of the load module, and finally obtains the control force and control torque of each degree of freedom.

The dynamic equation of the load module is

Among them, χ _B = (x _B , y _B , z _B , α _B , β _B , γ _B ) ^T is the generalized coordinate of the satellite ontology, χ _P = (x _P , y _P , z _P , α _P , β _P ,γ _P ) ^T is the generalized coordinate of the separate load module, M _P is the inertia matrix of the load module, C and K represent the equivalent damping and equivalent stiffness matrices of the actuator, respectively, where the subscripts B and P represent satellites, respectively Body and load modules. The relative motion between the load modules is obtained by transforming the relative coordinate system of the absolute motion equation of the load modules. F _ic and T _ic are the control force and the control torque, respectively, and F _iPd and T _iPd are the disturbance force and the disturbance torque, respectively. Have

Among them, F is the vector composed of the output force of 8 voice coil actuators. The relationship between the output force F of a single voice coil actuator and the current I is given by the following formula:

F=kBLI (7)

Among them, L is the effective length of the wire, B is the magnetic field strength, and k is the inductive coefficient of the wire.

Step 5. The load module has no drag and pointing control and satellite body follow control. After the motion of the load module and the inter-satellite pointing meet the gravitational wave measurement conditions, unlock the detection mass, measure the non-gravitational acceleration of the load module in the displacement mode or the acceleration mode, and decouple the six-degree-of-freedom motion of the load module, so that the two There are a total of 12 degrees of freedom in each load module, and 6 degrees of freedom are selected as sensitive degrees of freedom, and the remaining 6 degrees of freedom are non-sensitive degrees of freedom. At this time, the control of the load module is divided into the non-drag control of the sensitive degree of freedom, the pointing control, and the stable control of the non-sensitive degree of freedom. Sensitive degrees of freedom of the load module. Simultaneous non-drag control and inter-satellite pointing control through separate electromagnetic actuators. The non-drag control is used to follow the detection mass and eliminate the non-gravitational acceleration of the load module in the measurement frequency band of scientific missions; pointing control It is used to control the pointing between the load modules of adjacent satellites in the non-measurement frequency band to ensure the accuracy of laser measurement. The six insensitive degrees of freedom of the load module are stably controlled by separate electromagnetic actuators. Under the simultaneous action of no-drag control, pointing control and stability control, the payload module realizes the ultra-static environment and ultra-high-precision inter-satellite pointing required for scientific exploration. The satellite body tracks the two load modules through the conventional pose control method, and controls the relative motion between the satellite body and the load module to avoid collision with the load module and interfere with the measurement results.

Claims (4)

1. A multi-star super-long baseline composite formation method is characterized in that: a satellite is equipped with two load modules, each load module is equipped with a gravitational wave detection mass block, and the satellite body and the two load modules pass through 8 The separate electromagnetic actuators are connected; the separate connection interface adopts an eight-bar isotropic configuration; the steps are as follows: Step 1: In the rough formation stage, the separate electromagnetic actuators between the satellite body and the two load modules are rigidly connected respectively; a local reference orbit coordinate system is established between the satellites to describe the relative motion of the double star body and the payload, and establish an inter-satellite coordinate system. Pointing to the kinematic model; the local reference orbit coordinate system is:

Among them, Δr ₁₂ =r ₁ -r ₂ , r ₁ and r ₂ are the position vectors of the two satellites in the inertial coordinate system, respectively, x _F , y _F , and z _F are the three axes of the local reference orbit coordinate system, respectively. The axis obeys the right-hand rule; Step 2: Taking the formation satellite running in the gravitational field at the center of the earth as an example, the orbital dynamics equation of satellite i in the coordinate system of the earth's center orbit is:

Among them, ri and vi are the position vector and velocity vector of satellite _i , respectively _{, and d i is} the disturbance; The attitude dynamics equation of satellite i is:

Among them, T _ic and T _id are the control torque and disturbance torque of the satellite body, respectively; The relative kinematics equation of the satellite can be obtained through the transformation of the relative coordinate system, in which the relative rate of change of the six degree-of-freedom vectors can be obtained from equation (4); the rate of change of the known vector relative to the fixed reference coordinate is equal to the rate of change of the vector in the moving coordinate The sum of the cross product of the rotational speed vector ω _i of the moving coordinate relative to the reference coordinate and the vector:

By measuring the relative positions and relative attitudes of the formation satellites, the kinematics and dynamics models derived from the above are used to control the relative motion between the formation satellites through the conventional pose control method, determine the local reference orbit coordinate system, and control the pointing of the satellites Preliminarily track the origin of the local reference orbit coordinate system; Step 3: When the relative motion of the satellite body is stable and the coarse pointing accuracy between the satellites meets the requirements of the working domain of the load module, the separate electromagnetic actuator starts to work and enters the fine formation stage; the separate electromagnetic actuator is unlocked, and the load module and the satellite body are unlocked. The rigid connection state is no longer maintained, and the satellite body and the two load modules are converted from the connected state to the separated state. At this time, the micro-vibration transmission path from the satellite platform to the load module is physically isolated, and the load module realizes ultra-static and ultra-stable; The two load modules of i point to the load modules of the adjacent satellites respectively; The relative attitude deviations of different load modules of adjacent satellites and their local orbit coordinate systems are obtained through the calculation of the measurement system, and the attitude of the load modules is controlled by a separate electromagnetic actuator to track the local orbit coordinate system; According to the structure and layout of the separate electromagnetic actuators, each separate electromagnetic actuator outputs force at different positions of the load module, and finally obtains the control force and control torque of each degree of freedom; The dynamic equation of the load module is

Among them, χ _B = (x _B , y _B , z _B , α _B , β _B , γ _B ) ^T is the generalized coordinate of the satellite ontology, χ _P = (x _P , y _P , z _P , α _P , β _P ,γ _P ) ^T is the generalized coordinate of the separate load module, M _P is the inertia matrix of the load module, C and K represent the equivalent damping and equivalent stiffness matrices of the actuator, respectively, where the subscripts B and P represent satellites, respectively Body and load module; the relative motion between load modules is obtained by transforming the relative coordinate system of the absolute motion equation of the load module, F _ic and T _ic are the control force and control torque, respectively, F _iPd , T _iPd are the disturbance force and disturbance torque, respectively ;Have

Among them, F is the vector composed of the output forces of 8 separate electromagnetic actuators; the relationship between the output force F of a single separate electromagnetic actuator and the current I is given by the following formula: F=kBLI (7) Among them, L is the effective length of the wire, B is the magnetic field strength, and k is the inductive coefficient of the wire; Step 4: After the motion of the load module and the direction between the stars meet the gravitational wave measurement conditions, unlock the detection mass, measure the non-gravitational acceleration of the load module in the displacement mode or the acceleration mode, and decouple the six-degree-of-freedom motion of the load module , for the total 12 degrees of freedom of the two load modules, 6 degrees of freedom are selected as sensitive degrees of freedom, and the remaining 6 degrees of freedom are non-sensitive degrees of freedom; at this time, the control of the load module is divided into sensitive degrees of freedom without drag control and pointing control, Stable control of non-sensitive degrees of freedom; non-drag control and inter-satellite pointing control are simultaneously performed for the sensitive degrees of freedom of the load module through separate electromagnetic actuators. The non-gravitational acceleration; the pointing control is used to control the pointing between the load modules of adjacent satellites in the non-measurement frequency band to ensure the accuracy of laser measurement; the six non-sensitive degrees of freedom of the load module are stably controlled by separate electromagnetic actuators; Under the simultaneous action of non-drag control, pointing control and stability control, the payload module realizes the ultra-static environment and ultra-high-precision inter-satellite pointing required for scientific exploration; the satellite body tracks the two payload modules through the conventional attitude and attitude control method, and conducts satellite tracking. The relative movement between the body and the load module is kept under control to avoid collision with the load module and interfere with the measurement results; multiple satellites form an ultra-long baseline composite formation system. 2. a kind of multi-star super-long baseline compound formation method according to claim 1, it is characterized in that described separate electromagnetic actuator selects voice coil actuator, realizes the force of force by feeding direct current in coil. Output, the magnitude and direction of the output force can be changed by adjusting the magnitude and direction of the input current. 3. a kind of multi-satellite ultra-long baseline compound formation method according to claim 1, is characterized in that the rigid connection mode of the separate electromagnetic actuator between step 1 satellite body and two load modules is electromagnetic locking or active control. 4. A kind of multi-satellite ultra-long baseline composite formation method according to claim 1 is characterized in that the mover and the stator of 8 separate electromagnetic actuators are respectively connected with the satellite body and the load module through bolts.

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