Experimental Method for Separated Satellite Pointing Control Composite Formation

Abstract

The present invention relates to a method for an experiment on a separated satellite pointing control composite formation. First, the satellite body and the payload module are electromagnetically locked and connected, or the payload module is actively controlled to follow the satellite body. The satellite body uses a conventional attitude actuator to perform coarse-precision pointing control of the target and coarse formation control between satellite bodies. Subsequently, the separated electromagnetic actuator is released to generate two-axis translation control force and single-axis pointing control torque, and perform high-precision pointing of the payload module to the target and relative motion control between payload modules. Finally, after the formation is stable, the satellite body tracks the pointing of the payload module to maintain the relative position between the modules and prevent the modules from colliding. Ultimately, multiple separated experimental satellites cooperate together to achieve pointing control composite formation. This method adopts a separated satellite design, designs a separated multi-degree-of-freedom control system between the satellite body and the payload module, and performs high-precision and high-stability control of the payload module pointing through secondary formation.

Description

Technical Field

[0001] The present invention relates to a high-precision control technology for a super-quiet and super-stable separated satellite, and particularly to an experimental method for the pointing control composite formation of a separated satellite. Background Art

[0002] The space high-precision distributed optical system poses high requirements for the formation pointing and the super-quiet and super-stable mechanical environment of the optoelectronic payload. A multi-satellite composite formation system is composed of separated satellites. In the separated design, the satellite body and the payload module are connected by a separated drive platform, which can isolate the satellite platform disturbance. The satellite body does not need to be designed for super-quiet and super-stable performance. By performing composite formation control between the satellite body and the payload module respectively, the high-precision relative motion requirements between satellites can be met, and system solutions can be provided for fields such as ultra-long baseline gravitational wave detection and ultra-large aperture space distributed telescopes.

[0003] The composite formation method of the separated satellite first performs coarse formation control on multiple satellite bodies using traditional satellite actuators such as micro-thrusters and reaction wheels to initially form a formation configuration. After the coarse formation accuracy enters the fine formation process, the separated drive platform is unlocked, and the separated electromagnetic actuator is controlled to perform relative motion control between multiple payload modules to achieve fine formation.

[0004] In order to simulate the situation of the separated satellite forming a composite formation to perform cooperative orientation on a space target on the ground, a single-axis air-bearing turntable and a translational air-bearing table are respectively used as the separated satellite platform and the payload module for composite formation pointing control. The master-slave form is adopted to realize the composite formation process from coarse to fine and perform pointing maintenance, and at the same time, the relative motion cooperative control between the satellite platform and the payload module is studied.

[0005] The Chinese invention patent with the authorization publication number CN111099045B discloses a full physical simulation method for the dynamics and control of a double-super satellite air-bearing platform. Aiming at the ground verification method for a separated satellite with ultra-high pointing accuracy and ultra-high stability, a full physical simulation method for the air-bearing platform is proposed. By using the air-bearing method to offset gravity and simulate the space weightlessness environment, the rotation and translation control of the two modules of a single satellite can be experimentally verified.

[0006] This patent only studies the motion control of the two modules of a single separated satellite, does not verify the relative motion control between the modules of multiple separated satellites, and does not study the formation process of the formation from coarse to fine. The present invention studies the formation process of the separated satellite pointing control composite formation and verifies the composite formation pointing control ability. Summary of the Invention

[0007] Technical Problems to be Solved

[0008] To verify the separated satellite composite formation technology for space distributed systems on the ground, the present invention provides an experimental method for separated satellite pointing control composite formation.

[0009] Technical solution

[0010] The present invention provides an experimental method for separated satellite pointing control composite formation. In this method, the satellite body is installed on a single-axis air-bearing turntable to simulate the single-axis rotation of the satellite, and the turntable base is fixedly connected to the ground; the payload module is installed on a translational air-bearing stage to simulate the two-axis translation and single-axis rotation of the payload module, and the translational air-bearing stage is placed on a smooth marble platform. A separated drive platform is used to connect the satellite body and the payload module. A single formation satellite consists of a satellite body and a payload platform. High-precision pointing of the payload module is achieved through coarse formation of the satellite body and fine formation of the payload module in sequence. First, the satellite body and the payload module are electromagnetically locked and connected, or the payload module is actively controlled to follow the satellite body. The satellite body uses a conventional attitude actuator to perform coarse-precision pointing control of the target and coarse formation control between satellite bodies. Subsequently, the separated electromagnetic actuator is released to generate two-axis translation control force and single-axis pointing control torque, and perform high-precision pointing of the payload module to the target and relative motion control between payload modules. Finally, after the formation is stable, the satellite body tracks the pointing of the payload module to maintain the relative position between modules and prevent collisions between modules. Ultimately, multiple separated experimental satellites cooperate together to achieve pointing control composite formation.

[0011] An experimental method for separated satellite pointing control composite formation, characterized by the following steps:

[0012] Step 1: One satellite body carries one payload module. The satellite body is installed on a ground single-axis air-bearing turntable to simulate the single-axis rotation of the satellite, and the payload module is installed on a translational air-bearing stage placed on a smooth marble table to simulate the two-axis translation and single-axis rotation of the payload module; the satellite body and the two payload modules are respectively connected by 3 separated electromagnetic actuators. The movers and stators of the 3 separated electromagnetic actuators are respectively connected to the satellite body and the payload module by bolts; the separated formation satellites are divided into 1 main satellite and n slave satellites. Among them, the main satellite controls the pointing to a fixed target to ensure the overall pointing of the formation, and the slave satellites perform relative motion control relative to the main satellite to maintain the formation configuration;

[0013] Step 2: In the rough formation stage, the satellite body and the payload module are electromagnetically locked or the payload module is actively controlled to follow the satellite body; when the active control method is used to control the payload module to follow the body for relative motion maintenance and achieve low-stiffness connection between the body and the payload module, the movement of the separable electromagnetic actuators is measured by multiple separable displacement sensors installed parallel to the axes of the separable actuators, and displacement maintenance control is performed on the separable electromagnetic actuators to avoid collision between the satellite body and the payload module or the actuator from coming off, and indirect relative motion maintenance control is formed for the body and the payload module;

[0014] By measuring the pointing of the main satellite body to the target and the relative pointing of the slave satellite body to the main satellite body, the absolute pointing of the main satellite body and the relative pointing between the slave satellite body and the main satellite body are controlled by conventional attitude control methods;

[0015] Step 3: When the relative motion of the satellite body is stable and the inter-satellite rough pointing accuracy meets the working range requirements of the payload module, the separable electromagnetic actuators start to work and enter the fine formation stage; the separable electromagnetic actuators are unlocked, and the payload module and the satellite body no longer maintain the connection state, and the satellite body and the two payload modules are converted from the connected state to the separated state. At this time, the micro-vibration transfer path from the satellite platform to the payload module is physically isolated, and the payload module achieves ultra-quiet and ultra-stable; the payload module of the slave satellite i performs relative motion control with the payload module of the main satellite;

[0016] By measuring the pointing of the main satellite payload to the target, the relative pointing of the slave satellite payload to the main satellite payload, and the displacement of the separable electromagnetic actuators, the measurement information obtained by the main satellite payload and the slave satellite payload is respectively sent to the main satellite body and the slave satellite body by using the wireless communication system, a control force and control torque model of the separable electromagnetic actuators is established, the control problem of the payload module with the displacement constraint of the separable electromagnetic actuators is solved to obtain the control command, and the two-axis translation control force and the single-axis rotation control torque generated according to the control command are distributed to each separable electromagnetic actuator, and the absolute pointing of the main satellite payload and the relative pointing between the slave satellite payload and the main satellite payload are controlled by the separable electromagnetic actuators to achieve high-precision inter-satellite composite formation pointing control and avoid collision between modules at the same time;

[0017] Step 4: After the absolute pointing of the main satellite payload module and the relative pointing of the slave satellite payload module meet the composite formation requirements, enter the composite formation maintenance stage; the movement of the separable electromagnetic actuators is measured by multiple displacement sensors installed parallel to the axes of the separable electromagnetic actuators. When the pointing of the payload module remains unchanged, i.e., T zSolve for the control forces of the three separated electromagnetic actuators under the constraint of = 0, control the displacements of the three separated electromagnetic actuators, and avoid collisions, that is, make max{|Δl1|, Δl2|, |Δl3|} < ΔL; at the same time, calculate the relative orientations of the main satellite body and the payload module, and the slave satellite body and the payload module respectively according to the measurement results of the displacement sensors, and control the orientations of the main satellite body and the slave satellite body to track the orientations of the main satellite payload module and the slave satellite payload module respectively. Finally, form a separated satellite orientation control composite formation to obtain a high-precision payload orientation.

[0018] A further technical solution of the present invention: The step of obtaining the orientation of the main satellite body to the target in step 2 is as follows:

[0019] A laser fixed on the ground is used as the orientation target, and it is irradiated on the photodetector through the angle measurement optical path installed on the main satellite body. Calculate the deflection angle of the main satellite body according to the spot position; the calculation formula is:

[0020] θ SL = Δx SL / f SL (1)

[0021] Where, θ SL is the orientation angle of the main satellite body, Δx SL is the coordinate of the laser spot on the photodetector of the main satellite body, and f SL is the focal length of the angle measurement optical path of the main satellite body.

[0022] A further technical solution of the present invention: The step of obtaining the relative orientation between the slave satellite body and the main satellite body in step 2 is as follows:

[0023] A laser installed on the main satellite body is used as the reference light, and it is irradiated on the photodetector through the angle measurement optical path installed on the slave satellite body. Calculate the deflection angle of the slave satellite body according to the spot position; the calculation formula is:

[0024] θ SF = Δx SF / f SF (2)

[0025] Where, θ SF is the orientation angle of the slave satellite body, Δx SF is the coordinate of the laser spot on the photodetector of the slave satellite body, and f SF is the focal length of the angle measurement optical path of the slave satellite body.

[0026] A further technical solution of the present invention: The step of obtaining the orientation of the main satellite payload to the target in step 3 is as follows:

[0027] The laser fixed on the ground serves as a pointing target. It irradiates on the photodetector through the angle-measuring optical path installed on the payload of the main satellite. The deflection angle of the payload of the main satellite is calculated based on the position of the light spot. The calculation formula is as follows:

[0028] θ PL = Δx PL / f PL (3)

[0029] Where, θ PL is the pointing angle of the payload of the main satellite, Δx PL is the coordinate of the laser light spot on the photodetector of the payload of the main satellite, and f PL is the focal length of the angle-measuring optical path of the payload of the main satellite.

[0030] The further technical solution of the present invention: The obtaining step of the relative pointing between the slave satellite payload and the main satellite payload in step 3 is as follows:

[0031] The laser installed on the payload of the main satellite serves as the reference light. It irradiates on the photodetector through the angle-measuring optical path installed on the slave satellite payload. The deflection angle of the slave satellite payload is calculated based on the position of the light spot. The calculation formula is as follows:

[0032] θ PF = Δx PF / f PF (4)

[0033] Where, θ PF is the pointing angle of the slave satellite payload, Δx PF is the coordinate of the laser light spot on the photodetector of the slave satellite payload, and f PF is the focal length of the angle-measuring optical path of the slave satellite payload.

[0034] The further technical solution of the present invention: The displacement acquisition of the separable electromagnetic actuator in step 3 is specifically as follows:

[0035] The separable electromagnetic actuator generally can select a voice coil actuator. The force output is realized by passing a direct current through the coil. The magnitude and direction of the output force are changed by adjusting the magnitude and direction of the input current. According to the structural layout mode of the actuator, each actuator outputs force at different positions of the payload module, and finally the control force and control torque of each degree of freedom are obtained;

[0036] The relationship between the output force magnitude F of a single voice coil actuator and the current I is given by the following formula:

[0037] F = kBLI (5)

[0038] Where, L is the effective length of the wire, B is the magnetic field strength, and k is the inductive coefficient of the wire;

[0039] The control solution problem of the payload module with the displacement constraint of the separable electromagnetic actuator is:

[0040]

[0041] Among them, u is the generated two-axis translation control force and single-axis rotation control torque, and θ e is the attitude error of the load module, Δl1, Δl2, and Δl3 are the displacements of the split electromagnetic actuators 1, 2, and 3 relative to the zero point respectively, and ΔL is 1 / 2 of the stroke of the split electromagnetic actuator;

[0042] The control force and control torque model of the split electromagnetic actuator is:

[0043] Assume that the two-axis translation control force and single-axis rotation control torque are [F x F y T z T , which is generated by the cooperation of 3 split electromagnetic actuators, and its output force is [F1 F2 F3] T , according to the installation method, there is

[0044]

[0045] That is

[0046]

[0047] Among them, d1 is the perpendicular distance between the acting force axis of the split electromagnetic actuator 1 and the x-axis of the load module body coordinate system, d2 is the perpendicular distance between the acting force axis of the split electromagnetic actuator 2 and the y-axis of the load module body coordinate system, d3 is the perpendicular distance between the acting force axis of the split electromagnetic actuator 3 and the z-axis of the load module body coordinate system, and the body coordinate system is fixedly connected to the center of mass of the load module.

[0048] A computer system, characterized in that it includes: one or more processors, a computer-readable storage medium for storing one or more programs, wherein, when the one or more programs are executed by the one or more processors, the one or more processors are caused to implement the above method.

[0049] A computer-readable storage medium, characterized in that it stores computer-executable instructions, and the instructions are used to implement the above method when executed.

[0050] Beneficial effects

[0051] ​A separated satellite pointing control composite formation experiment method provided by the present invention. This method uses an air-bearing platform to form multiple separated satellites, a single-axis air-bearing turntable to form the separated satellite body, and a translational air-bearing table to form the three-degree-of-freedom payload module of the separated satellite, which can well simulate the rotation of the separated satellite in space, realize the ground experimental verification of the dynamics and control of the separated satellite composite formation, and provide research references and experimental methods for the system structure and control design of the composite formation technology in actual tasks. This method uses lasers and high-precision photodetectors to measure the absolute pointing of the separated satellite main body and payload module, the relative pointing between the main and slave satellite bodies, and between the payloads. The separated measurement method does not interfere with the controlled object and provides technical references for the absolute and relative motion measurement of the space composite formation. By measuring the displacement of the separated electromagnetic actuator through the displacement sensor of the separated interface and designing the displacement holding method of the separated electromagnetic actuator, it can control the payload module to follow the satellite body during the coarse formation stage and cooperate with the payload module to form and maintain the precise formation control, avoiding collisions between modules or the actuator from coming out, and improving the reliability of the composite formation experiment.

[0052] Compared with the traditional formation ground verification technology, this method adopts a separated satellite design, designs a separated multi-degree-of-freedom control system between the satellite body and the payload module, and the separated interface physically isolates the interference of the satellite body on the payload. The satellite body can be conventionally designed with moving parts such as solar panels and flywheels, and the pointing of the payload module is controlled with high precision and high stability through secondary formation. Brief Description of the Drawings

[0053] The drawings are only for the purpose of showing specific embodiments and are not considered as limitations of the present invention. Throughout the drawings, the same reference signs represent the same components.

[0054] Figure 1 It is a schematic diagram of the separated satellite pointing control composite formation experiment system of the present invention.

[0055] Figure 2 It is a top view of the separated satellite pointing control composite formation experiment system of the present invention.

[0056] Figure 3 It is a schematic diagram of the separated satellite pointing measurement system of the present invention. The optical path of the measurement system used in the detailed description is the same, the meanings of the variables are the same, and different subscripts in the measurement calculation formula only represent the same measurement parameter of different modules.

[0057] Figure 4 It is a schematic cross-sectional view of the separated electromagnetic actuator of the present invention. The separated electromagnetic actuators used in the detailed description are the same, the meanings of the variables are the same, and different subscripts in the symbols only represent the same parameter of different separated electromagnetic actuators.

[0058] Figure 5 It is a schematic diagram of the control system for the separated satellite pointing control composite formation experiment method of the present invention.

[0059] In the figure, 1 - Separated electromagnetic actuator 1, 2 - Separated electromagnetic actuator 2, 3 - Separated electromagnetic actuator 3, 4 - Separated displacement sensor 1, 5 - Separated displacement sensor 2, 6 - Separated displacement sensor 3, 7 - Photoelectric detection system, 8 - Single-axis air-bearing, 9 - Reaction wheel. Specific implementation mode

[0060] In order to make the objectives, technical solutions and advantages of the present invention clearer and more understandable, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0061] Referring to Figures 1-3 As shown, the present invention provides a separated satellite pointing control composite formation experiment method. In this method, the satellite body is installed on a single-axis air-bearing turntable to simulate the rotation of the satellite around the z-axis, and the turntable base is fixedly connected to the ground; the payload module is installed on a translation air-bearing platform to simulate the translation of the payload module along the x and y axes and the rotation around the z-axis, and the translation air-bearing platform is placed on a smooth marble platform. A separated drive platform is used to connect the satellite body and the payload module. Each formation satellite consists of a satellite body and a payload platform. High-precision pointing of the payload module is achieved through the coarse formation of the satellite body and the fine formation of the payload module in sequence. First, the satellite body and the payload module are electromagnetically locked and connected, or the payload module is actively controlled to follow the satellite body. The satellite body uses a reaction wheel to perform coarse-precision pointing control of the target and coarse formation control between satellite bodies. Subsequently, the separated electromagnetic actuator is released to generate two-axis translation control force and single-axis pointing control torque for high-precision pointing of the payload module to the target and relative motion control between payload modules. Finally, after the formation is stable, the satellite body tracks the pointing of the payload module to maintain the relative position between the modules and prevent the modules from colliding. Ultimately, multiple separated experimental satellites cooperate with each other to achieve pointing control composite formation.

[0062] Specifically, as Figure 1As shown in the figure, the present invention is provided with a separated main satellite, multiple separated slave satellites and a smooth marble table to form a separated satellite pointing control composite formation system. Each separated experimental satellite includes a single-axis air-bearing turntable, a satellite body, a translational air-bearing table and a payload module. The satellite body is installed on the single-axis air-bearing turntable to simulate the single-axis rotation of the satellite, and the payload module is installed on the translational air-bearing table to simulate the translational motion of the payload module in the x and y axes and the rotational motion around the z axis. The satellite body and the payload module are connected through a separated interface to jointly form a separated experimental satellite. Taking the satellite body driven by a flywheel as an example, as Figure 2 shown, the separated satellite body includes power supply, data acquisition, wireless communication, real-time control, computer, laser, photoelectric detection system and separated electromagnetic actuator, separated displacement sensor, reaction flywheel (including drive system). The separated satellite payload module includes gas supply, power supply, data acquisition, wireless communication, laser, photoelectric detection system. The single-axis air-bearing turntable used by the satellite body is externally supplied with gas, and both the satellite body and the payload module are affected by environmental disturbances. The laser fixed on the ground is used as the reference target for the separated main satellite body and the payload module. The relative pointing of the slave satellite body and the payload module to the main satellite body and the payload module respectively is measured by the laser installed on the main satellite and the photoelectric detection system installed on the slave satellite. The information between the separated satellite body and the payload is transmitted through the wireless communication system. The separated satellite pointing control composite formation is realized through the coarse formation of the body, the fine formation of the payload, and the maintenance of the composite formation in sequence.

[0063] More specifically, the experimental method for the separated satellite pointing control composite formation is divided into the following steps:

[0064] Step 1: Design of the separated satellite pointing control composite formation system. One satellite body carries one payload module. The satellite body is installed on the ground single-axis air-bearing turntable to simulate the single-axis rotation of the satellite, and the payload module is installed on the translational air-bearing table placed on the smooth marble table to simulate the translational motion of the payload module in the x and y axes and the rotational motion around the z axis. The satellite body and the two payload modules are respectively connected through three separated electromagnetic actuators. The movers and stators of the three separated electromagnetic actuators are respectively connected to the satellite body and the payload module through bolts. The separated formation satellites are divided into one main satellite and n slave satellites. The pointing control of the main satellite to the fixed target is used to ensure the overall pointing of the formation, and the relative motion control of the slave satellite relative to the main satellite is used to maintain the configuration of the formation. The separated optoelectronic system is used to measure the pointing motion of the separated satellite.

[0065] Step 2: Coarse formation control of the separated satellite body pointing. During the coarse formation stage, the satellite body and the payload module are electromagnetically locked or the payload module is actively controlled to follow the satellite body. When the active control method is used to control the payload module to follow the body for relative motion maintenance and achieve a low-rigidity connection between the body and the payload module, the motion of the separated electromagnetic actuators is measured by multiple separated displacement sensors installed parallel to the axes of the separated electromagnetic actuators, and displacement maintenance control is performed on the separated electromagnetic actuators to avoid collisions between the satellite body and the payload module or the actuator from disengaging, thereby forming an indirect relative motion maintenance control for the body and the payload module.

[0066] By measuring the pointing of the main satellite body to the target and the relative pointing of the slave satellite body to the main satellite body, the satellite body generates a control torque to drive the reaction wheel for the absolute pointing of the main satellite body and the relative pointing control between the slave satellite body and the main satellite body.

[0067] The steps for obtaining the pointing of the main satellite body to the target are as follows:

[0068] The laser fixed on the ground is used as the pointing target. The angular measurement optical path installed on the main satellite body irradiates on the photodetector, and the deflection angle of the main satellite body is calculated according to the spot position. The calculation formula is:

[0069] θ SL =Δx SL / f SL (1)

[0070] Where, θ SL is the pointing angle of the main satellite body, Δx SL is the coordinate of the laser spot on the photodetector of the main satellite body, and f SL is the focal length of the angular measurement optical path of the main satellite body.

[0071] The steps for obtaining the relative pointing between the slave satellite body and the main satellite body are as follows:

[0072] The laser installed on the main satellite body is used as the reference light. The angular measurement optical path installed on the slave satellite body irradiates on the photodetector, and the deflection angle of the slave satellite body is calculated according to the spot position. The calculation formula is:

[0073] θ SF =Δx SF / f SF (2)

[0074] Where, θ SF is the pointing angle of the slave satellite body, Δx SF is the coordinate of the laser spot on the photodetector of the slave satellite body, and f SF is the focal length of the angular measurement optical path of the slave satellite body.

[0075] Step 3: Precise formation control for the separable satellite payload module. When the relative motion of the satellite body is stable and the coarse inter-satellite pointing accuracy meets the working range requirements of the payload module, the separable electromagnetic actuator starts to work and enters the precise formation stage. The separable electromagnetic actuator unlocks, and the payload module no longer maintains a connection with the satellite body. The satellite body and the two payload modules change from a connected state to a separated state. At this time, the micro-vibration transfer path from the satellite platform to the payload module is physically isolated, and the payload module achieves ultra-low vibration and ultra-high stability. Perform relative motion control of the payload module of satellite i with respect to the payload module of the main satellite.

[0076] By measuring the pointing of the main satellite payload to the target, the relative pointing between the slave satellite payload and the main satellite payload, and the displacement of the separable electromagnetic actuator, use the wireless communication system to send the measurement information obtained by the main satellite payload and the slave satellite payload to the main satellite body and the slave satellite body respectively, establish a control force and control torque model for the separable electromagnetic actuator, solve the control problem of the payload module with the displacement constraint of the separable electromagnetic actuator to obtain control instructions, and distribute the translational control forces along the x and y axes and the rotational control torque around the z axis generated according to the control instructions to each separable electromagnetic actuator. Control the absolute pointing of the main satellite payload and the relative pointing between the slave satellite payload and the main satellite payload through the separable electromagnetic actuator, achieve high-precision inter-satellite composite formation pointing control, and avoid collisions between modules at the same time.

[0077] The steps for obtaining the pointing of the main satellite payload to the target are as follows:

[0078] A laser fixed on the ground is used as the pointing target. It irradiates on the photodetector through the angle measurement optical path installed on the main satellite payload, and calculates the deflection angle of the main satellite payload according to the spot position. The calculation formula is:

[0079] θ PL =Δx PL / f PL (3)

[0080] Where, θ PL is the pointing angle of the main satellite payload, Δx PL is the coordinate of the laser spot on the photodetector of the main satellite payload, and f PL is the focal length of the angle measurement optical path of the main satellite payload.

[0081] The steps for obtaining the relative pointing between the slave satellite payload and the main satellite payload are as follows:

[0082] A laser installed on the main satellite payload is used as the reference light. It irradiates on the photodetector through the angle measurement optical path installed on the slave satellite payload, and calculates the deflection angle of the slave satellite payload according to the spot position. The calculation formula is:

[0083] θ PF =Δx PF / f PF (4)

[0084] Among them, θ PF is the pointing angle of the slave satellite payload, and Δx PF is the coordinate of the laser spot on the optoelectronic detector of the slave satellite payload, and f PF is the focal length of the angle measurement optical path of the slave satellite payload.

[0085] The separable electromagnetic actuator generally can select a voice coil actuator. By passing a direct current through the coil, the output of force is realized. By adjusting the magnitude and direction of the input current, the magnitude and direction of the output force are changed. According to the structural layout mode of the actuator, each actuator outputs force at different positions of the payload module, and finally the control force and control torque of each degree of freedom are obtained.

[0086] The relationship between the output force magnitude F of a single voice coil actuator and the current I is given by the following formula:

[0087] F = kBLI (5)

[0088] 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.

[0089] The problem of solving the control of the payload module with displacement constraints of the separable electromagnetic actuator is:

[0090]

[0091] Among them, u is the generated translational control force of the x and y axes and the rotational control torque about the z axis, and θ e is the attitude error of the payload module, and Δl1, Δl2, and Δl3 are the displacements of the separable electromagnetic actuators 1, 2, and 3 relative to the zero point respectively, and ΔL is 1 / 2 of the stroke of the separable electromagnetic actuator.

[0092] The control force and control torque model of the separable electromagnetic actuator is:

[0093] Assume that the translational control forces of the x and y axes and the rotational control torque about the z axis are [F x F y T z T , which is generated by the cooperation of 3 separable electromagnetic actuators, and its output force is [F1 F2 F3] T , according to the installation method, there is

[0094]

[0095] That is

[0096]

[0097] ​Among them, d1 is the perpendicular distance between the acting force axis of the separable electromagnetic actuator 1 and the x-axis of the load module body coordinate system, d2 is the perpendicular distance between the acting force axis of the separable electromagnetic actuator 2 and the y-axis of the load module body coordinate system, d3 is the perpendicular distance between the acting force axis of the separable electromagnetic actuator 3 and the z-axis of the load module body coordinate system, and the body coordinate system is fixedly connected to the center of mass of the load module.

[0098] Step Four: Maintain the composite formation of the separable satellite pointing control. After the absolute pointing of the main satellite load module and the relative pointing of the slave satellite load module meet the requirements of the composite formation, enter the stage of maintaining the composite formation. Measure the movement of the separable electromagnetic actuators through multiple displacement sensors installed parallel to the axes of the respective separable electromagnetic actuators. Under the constraint that the pointing of the load module remains unchanged, i.e., T z = 0, solve for the control forces of the three separable electromagnetic actuators, control the displacements of the three separable electromagnetic actuators, and avoid collisions, that is, make max{|Δl1|, Δl2|, |Δl3|} < ΔL; at the same time, calculate the relative pointings of the main satellite body and the load module, and the slave satellite body and the load module respectively according to the measurement results of the displacement sensors, and drive the reaction flywheels to control the pointings of the main satellite body and the slave satellite body to track the pointings of the main satellite load module and the slave satellite load module respectively. Finally, form the composite formation of the separable satellite pointing control and obtain high-precision load pointing.

[0099] As described above, it is only the specific implementation manner of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art within the technical scope disclosed by the present invention can easily think of various equivalent modifications or substitutions, and these modifications or substitutions should all be covered within the protection scope of the present invention.

Claims

1. A separation-type satellite pointing control composite formation experiment method, characterized in that The steps are as follows: Step 1: A satellite body carries a payload module. The satellite body is installed on a ground single-axis air-bearing turntable to simulate the single-axis rotation of the satellite, and the payload module is installed on a translational air-bearing stage placed on a smooth marble table to simulate the two-axis translation and single-axis rotation of the payload module. The satellite body is respectively connected to the two payload modules through three separable electromagnetic actuators. The movers and stators of the three separable electromagnetic actuators are respectively connected to the satellite body and the payload module through bolts. The separable formation satellites are divided into one master satellite and n slave satellites. The master satellite controls the pointing to a fixed target to ensure the overall pointing of the formation, and the slave satellites perform relative motion control relative to the master satellite to maintain the formation configuration; Step 2: In the coarse formation stage, the satellite body and the payload module are electromagnetically locked or the payload module is actively controlled to follow the satellite body. When the active control method is used to control the payload module to follow the body for relative motion maintenance and achieve a low-rigidity connection between the body and the payload module, the motion of the separable electromagnetic actuator is measured by a plurality of separable displacement sensors installed parallel to the axis of each separable actuator, and displacement maintenance control is performed on the separable electromagnetic actuator to avoid collisions between the satellite body and the payload module or the actuator coming out, and indirect relative motion maintenance control is formed for the body and the payload module; By measuring the pointing of the master satellite body to the target and the relative pointing of the slave satellite body to the master satellite body, the absolute pointing of the master satellite body and the relative pointing between the slave satellite body and the master satellite body are controlled by the conventional attitude control method; Step 3: When the relative motion of the satellite body is stable and the coarse inter-satellite pointing accuracy meets the working range requirements of the payload module, the separable electromagnetic actuator starts to work and enters the fine formation stage. The separable electromagnetic actuator is unlocked, and the payload module and the satellite body no longer maintain the connection state. The satellite body and the two payload modules are converted from the connected state to the separated state. At this time, the micro-vibration transfer path from the satellite platform to the payload module is physically isolated, and the payload module achieves ultra-quiet and ultra-stable. The payload module of the slave satellite i performs relative motion control with the payload module of the master satellite; By measuring the pointing of the master satellite payload to the target, the relative pointing of the slave satellite payload to the master satellite payload, and the displacement of the separable electromagnetic actuator, the measurement information obtained by the master satellite payload and the slave satellite payload is respectively sent to the master satellite body and the slave satellite body by using the wireless communication system, a control force and control torque model of the separable electromagnetic actuator is established, the control problem of the payload module with the displacement constraint of the separable electromagnetic actuator is solved to obtain the control command, the two-axis translation control force and the single-axis rotation control torque generated according to the control command are distributed to each separable electromagnetic actuator, and the absolute pointing of the master satellite payload and the relative pointing between the slave satellite payload and the master satellite payload are controlled by the separable electromagnetic actuator to achieve high-precision inter-satellite composite formation pointing control and avoid collisions between modules at the same time; Step 4: After the absolute pointing of the primary satellite payload module and the relative pointing of the secondary satellite payload module meet the requirements of the composite formation, enter the composite formation maintenance stage; measure the movement of the separable electromagnetic actuators through multiple displacement sensors installed parallel to the axes of the separable electromagnetic actuators. Under the constraint that the pointing of the payload module remains unchanged, i.e., T z = 0, solve the control forces of the three separable electromagnetic actuators, control the displacements of the three separable electromagnetic actuators, and avoid collisions, that is, make max{|Δl1|, |Δl2|, |Δl3|} < ΔL; at the same time, calculate the relative pointings of the primary satellite body and the payload module, and the secondary satellite body and the payload module respectively according to the measurement results of the displacement sensors, and control the pointings of the primary satellite body and the secondary satellite body to track the pointings of the primary satellite payload module and the secondary satellite payload module respectively; finally, form a separable satellite pointing control composite formation to obtain high-precision payload pointing.

2. The experimental method for a separated satellite pointing control composite formation according to claim 1, wherein: The steps for obtaining the pointing of the master satellite body in Step 2 are as follows: A laser fixed on the ground is used as the pointing target, and it is irradiated on the photodetector through the angle measurement optical path installed on the master satellite body, and the deflection angle of the master satellite body is calculated according to the spot position; The calculation formula is: θ SL = Δx SL / f SL (1) Among them, θ SL is the pointing angle of the main star body, Δx SL is the coordinate of the laser spot on the photodetector of the main star body, and f SL is the focal length of the angle measurement optical path of the main star body.

3. A separation-type satellite pointing control composite formation experiment method according to claim 1, characterized in that: The obtaining step of the relative direction between the slave satellite body and the master satellite body in step 2 is as follows: The laser installed on the master satellite body is used as the reference light, and is irradiated on the photodetector through the angle measurement optical path installed on the slave satellite body, and the deflection angle of the slave satellite body is calculated according to the spot position; The calculation formula is: θ SF = Δx SF / f SF (2) Among them, θ SF is the pointing angle of the slave satellite body, Δx SF is the coordinate of the laser spot on the photodetector of the slave satellite body, and f SF is the focal length of the angle measurement optical path of the slave satellite body.

4. A method for an experiment on a separated satellite pointing control composite formation according to claim 1, characterized in that: The obtaining step of the direction of the master satellite payload to the target in step 3 is as follows: The laser fixed on the ground is used as the pointing target, and is irradiated on the photodetector through the angle measurement optical path installed on the master satellite payload, and the deflection angle of the master satellite payload is calculated according to the spot position; The calculation formula is: θ PL = Δx PL / f PL (3) Among them, θ PL is the pointing angle of the main satellite payload, Δx PL is the coordinate of the laser spot on the photoelectric detector of the main satellite payload, and f PL is the focal length of the angle measurement optical path of the main satellite payload.

5. A separated satellite pointing control composite formation experiment method according to claim 1, characterized in that: The obtaining step of the relative direction between the slave satellite payload and the master satellite payload in step 3 is as follows: The laser installed on the master satellite payload is used as the reference light, and is irradiated on the photodetector through the angle measurement optical path installed on the slave satellite payload, and the deflection angle of the slave satellite payload is calculated according to the spot position; The calculation formula is: θ PF = Δx PF / f PF (4) Among them, θ PF is the pointing angle of the slave satellite payload, Δx PF is the coordinate of the laser spot on the optoelectronic detector of the slave satellite payload, and f PF is the focal length of the angle measurement optical path of the slave satellite payload.

6. The experimental method for a separated satellite pointing control composite formation according to claim 1, characterized in that: The obtaining of the displacement of the separable electromagnetic actuator in step 3 is specifically as follows: The separable electromagnetic actuator generally can select a voice coil actuator, and the force output is realized by passing a direct current through the coil, and the magnitude and direction of the output force are changed by adjusting the magnitude and direction of the input current; according to the structural layout mode of the actuator, each actuator outputs force at different positions of the payload module, and finally the control force and control torque of each degree of freedom are obtained; The relationship between the output force magnitude F of a single voice coil actuator and the current I is given by the following formula: F = kBLI (5) where L is the effective length of the wire, B is the magnetic field strength, and k is the inductive coefficient of the wire; The control solution problem of the payload module with the displacement constraint of the separable electromagnetic actuator is: where u is the generated two-axis translation control force and single-axis rotation control torque, and θ e is the attitude error of the load module, Δl1, Δl2, and Δl3 are the displacements of the separable electromagnetic actuators 1, 2, and 3 relative to the zero point, and ΔL is 1 / 2 of the stroke of the separable electromagnetic actuator; The control force and control torque model of the separable electromagnetic actuator is: Assume that the two-axis translation control force and the single-axis rotation control torque are [F x F y T z T , which are generated by the cooperation of three separate electromagnetic actuators, and their output forces are [F1 F2 F3] T , according to the installation method, there are​ That is where d1 is the perpendicular distance between the action axis of the separable electromagnetic actuator 1 and the x-axis of the body coordinate system of the payload module, d2 is the perpendicular distance between the action axis of the separable electromagnetic actuator 2 and the y-axis of the body coordinate system of the payload module, d3 is the perpendicular distance between the action axis of the separable electromagnetic actuator 3 and the z-axis of the body coordinate system of the payload module, and the body coordinate system is fixedly connected to the center of mass of the payload module.

7. A computer system, characterized in that Including: One or more processors, a computer-readable storage medium for storing one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors implement the method described in claim 1.

8. A computer-readable storage medium, characterized in that Stored with computer-executable instructions, the instructions are used to implement the method described in claim 1 when executed.

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