CN115610713A - In-orbit docking device and docking system for spacecraft - Google Patents
In-orbit docking device and docking system for spacecraft Download PDFInfo
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- CN115610713A CN115610713A CN202211631205.0A CN202211631205A CN115610713A CN 115610713 A CN115610713 A CN 115610713A CN 202211631205 A CN202211631205 A CN 202211631205A CN 115610713 A CN115610713 A CN 115610713A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
- B64G1/646—Docking or rendezvous systems
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Abstract
The invention discloses an on-orbit docking device and a docking system for a spacecraft, relates to the technical field of spacecraft devices, and aims to improve docking tolerance and reduce power consumption. The docking device is arranged on a main spacecraft for docking a target spacecraft, is a rod-shaped mechanism and comprises a control section, a buffer section and a capturing section, wherein the capturing section comprises a limiting module and a capturing module capable of inclining towards any direction, the capturing module is used for acquiring attitude information and position information of the target spacecraft and docking or separating the attitude information and the position information with the target spacecraft through electromagnetic force, and the limiting module is used for limiting the capturing section; the control section comprises a control unit, the control unit is used for controlling the main spacecraft to approach the target spacecraft according to the attitude information and the position information and sending an electromagnetic adjusting instruction to the capturing module so as to control the capturing module to adsorb or repel the target spacecraft; the buffer section is used for absorbing impact. The butt joint device can reduce the precision requirement of butt joint and realize large-tolerance butt joint.
Description
Technical Field
The invention relates to the technical field of spacecraft devices, in particular to an in-orbit docking device and a docking system for a spacecraft.
Background
The space rendezvous and docking technology refers to a technology that two spacecrafts meet on a space track and are structurally connected into a whole, and can be widely applied to the fields of construction of various space facilities, on-orbit assembly, recovery, supply, maintenance, rescue and the like. In two space vehicles which are intersected and butted in space, a butted object can be an on-orbit large space vehicle or a space vehicle out of control or having a fault in space, and the whole process can be roughly divided into four stages of ground guidance, automatic searching, approaching and butting and folding.
In the prior art, the docking mechanism is mainly in a collision type docking mechanism, collision capture is realized based on relative motion between docked spacecrafts, the collision type docking mechanism has the technical problems of high requirements on attitude orbit control precision of the spacecrafts and large docking impact force, and meanwhile, a large amount of fuel is consumed during attitude orbit adjustment and plume pollution is possibly caused.
In the other docking form in the prior art, docking of the spacecraft is realized through magnetic force capture, wherein the magnetic force capture means that attitude adjustment and distance approaching of the docking tail section of the spacecraft are realized through a magnetic force adsorption mode, and finally docking is realized. The main forms of the magnetic force comprise electromagnetic force butt joint and permanent magnetic force butt joint, and the electromagnetic force butt joint has the main problems that the butt joint needs high power and high power consumption, and long-time power supply or mechanical structure locking is needed if long-time connection is realized; the main problems of the permanent magnet butt joint are that the permanent magnet is difficult to release after butt joint, and meanwhile, the electromagnetic force in the butt joint process cannot be adjusted. The magnetic docking method has higher requirements on attitude control precision.
Disclosure of Invention
In order to solve the above technical problems, embodiments of the present invention are intended to provide an in-orbit docking device and a docking system for a spacecraft, and the technical solution of the present invention is implemented as follows:
in a first aspect, the present invention provides an in-orbit docking device for a spacecraft, the docking device being mounted on a main spacecraft for docking a target spacecraft, the docking device being a rod-shaped mechanism, the docking device being divided into a control section, a buffer section, and a capture section, wherein the capture section includes a position limiting module and a capture module capable of tilting in any direction, the capture module being configured to acquire attitude information and position information of the target spacecraft and dock with or separate from the target spacecraft by electromagnetic force, the position limiting module being configured to limit the capture section; the control section comprises a control unit which is configured to control the main spacecraft to approach the target spacecraft according to the attitude information and the position information and send an electromagnetic adjusting instruction to the capture module to control the capture module to adsorb or repel the target spacecraft; the buffer section is configured to absorb an impact when the master spacecraft is docked with the target spacecraft.
In a second aspect, the invention provides a docking system for a spacecraft, the docking system comprising a docking device as described above, the docking system comprising a spacecraft body and at least one of the docking devices, the spacecraft body being configured as a cube structure, the docking system being configured to: the docking device is rotatably connected with the spacecraft main body and can rotate within the space by any angle within the range of 0-180 degrees.
According to the on-orbit docking device and the docking system for the spacecraft, disclosed by the invention, the structural stability of the main spacecraft in the launching and flying processes is ensured through the rod-shaped docking device, the installation envelope of the main spacecraft is reduced, and the influence of a magnetic field at the tail end of the docking device on the main spacecraft body is reduced; the controllability of the docking device is enhanced, and the angle of the capturing section is adjusted according to the position information and the attitude information of the target spacecraft, so that the docking precision is improved, and the requirement on the attitude control precision of the main spacecraft is reduced; the electromagnetic force of the capturing section is adjustably configured, so that stable connection can be realized without electrifying, and the power consumption is reduced; the butt joint device is strong in reproducibility, and can be developed to realize the joint capture of multiple spacecrafts to multiple spacecrafts.
Drawings
Fig. 1 is a schematic structural diagram of an in-orbit docking device for a spacecraft, which is disclosed by an embodiment of the invention;
FIG. 2 is an exploded view of a capture section of an in-orbit docking device for a spacecraft as disclosed in an embodiment of the present invention;
fig. 3 is a schematic structural diagram of a capture module of an in-orbit docking device for a spacecraft according to an embodiment of the present invention;
FIG. 4 is a schematic view of a first plate rotating relative to a second plate in a capture module of an in-orbit docking apparatus for a spacecraft in accordance with an embodiment of the present invention;
fig. 5 is a schematic structural diagram of a position limiting module of an in-orbit docking device for a spacecraft, which is disclosed in an embodiment of the invention;
fig. 6 is an assembly diagram of a capture module and a limit module of an in-orbit docking device for a spacecraft, according to an embodiment of the present invention;
FIG. 7 is a force analysis diagram of a rigid ball of an in-orbit docking device for a spacecraft according to an embodiment of the present invention;
fig. 8 is a schematic structural diagram of a first plate-like member of an in-orbit docking device for a spacecraft according to an embodiment of the present invention;
fig. 9 is a schematic structural diagram of a tracking unit of an in-orbit docking device for a spacecraft according to an embodiment of the present invention;
fig. 10 is a schematic structural diagram of an electromagnetic capture unit of an in-orbit docking device for a spacecraft according to an embodiment of the present invention;
FIG. 11 is a flowchart of a process for docking a host spacecraft with a target spacecraft in an embodiment of the present invention;
FIG. 12 is a schematic structural diagram of a docking system for a spacecraft in accordance with an exemplary embodiment of the present invention;
fig. 13 is a schematic view of the rotary deployment of all docking devices in the docking system for a spacecraft according to the embodiment of the present invention.
Wherein the reference numerals are: 1. a docking device; 10. a control section; 20. a buffer section; 30. a capture section; 31. a capture module; 311. a first plate-like member; 3111. a tracking unit; 31111. a signal light emitter; 31112. a signal light receiver; 3112. an electromagnetic capture unit; 31121. a permanent magnet; 31122. an electromagnet; 31123. a resin interlayer; 312. a rod-like member; 3121. a spherical bulge; 3122. a spring; 3123. a rigid ball; 313. a second plate-like member; 32. a limiting module; 321. spherical grooving is conducted; 33. a cover plate; 2. a docking system; 3. a cylindrical rotary joint.
Detailed Description
The technical solution in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention.
The existing docking modes of the spacecraft through electromagnetic capture can be divided into electromagnetic docking and permanent magnet docking, wherein the electromagnetic docking is realized by electrifying an electromagnet to attract an armature, the main problems of the electromagnet docking are that large power is needed during docking, the consumed electric quantity is large, continuous power supply is needed or an additional mechanical device is added to provide mechanical structure locking if long-time connection is realized, and the requirement on the attitude control precision of the main spacecraft is high in the docking process. The permanent magnet type butt joint is realized by utilizing mutual attraction between the permanent magnet and the permanent magnet or between the permanent magnet and the armature, is difficult to release after butt joint, and has unadjustable electromagnetic force in the butt joint process.
Based on this, referring to fig. 1, a schematic diagram of an in-orbit docking device for a spacecraft, provided by an embodiment of the present invention, is shown, wherein the docking device 1 is installed on a main spacecraft for docking a target spacecraft. Accordingly, the target spacecraft is provided with a magnetic device with fixed magnetic field intensity and fixed magnetic field direction, such as a permanent magnet, so that the docking device 1 can generate force on the target spacecraft through magnetic force, and the docking device 1 adsorbs or repels the target spacecraft to realize docking or separation of the main spacecraft and the target spacecraft. The docking device 1 is configured to be rod-shaped, and the docking device 1 is sequentially divided into: the device comprises a control section 10, a buffer section 20 and a capturing section 30, wherein the capturing section 30 is used for adsorbing or repelling the target spacecraft through electromagnetic force, the control section 10 controls the capturing section 30 to be adjusted according to different conditions, so that the weak impact butt joint or separation of the main spacecraft and the target spacecraft is realized, and the buffer section 20 is used for absorbing impact energy in impact to protect the spacecraft and devices on the spacecraft.
Referring to fig. 2, the capturing section 30 includes a capturing module 31 and a limiting module 32, the capturing module 31 is configured to obtain position information and attitude information of the target spacecraft, the capturing module 31 is configured to compensate an angular difference between itself and the target spacecraft through tilting, and to attract or repel the target spacecraft through changing a magnetic field strength and/or a magnetic field direction of itself under the control of the control section 10, and the limiting module 32 is configured to limit the capturing section 30, specifically, the limiting module 32 limits tilting of the capturing module 31.
The capturing module 31 includes a first plate-like member 311, a rod-like member 312, and a second plate-like member 313, the first plate-like member 311 is provided with a tracking unit 3111 for acquiring the position information and the posture information, and an electromagnetic capturing unit 3112 for generating a magnetic field, and the rod-like member 312 and the second plate-like member 313 together constitute a mechanical mechanism for driving the first plate-like member 311 to tilt. Wherein, referring to fig. 2 to 4, the rod 312 is fixed to the bottom of the first plate 311, and preferably, the rod 312 and the first plate 311 are integrally formed. The central portion of the rod 312 is provided with a spherical protuberance 3121, the second plate 313 is provided with a spherical through slot adapted to the spherical protuberance 3121, when the rod 312 passes through the spherical through slot and is connected with the second plate 313, the spherical protuberance 3121 is fitted in the spherical through slot and prevents the rod 312 from continuously passing through the spherical through slot, see fig. 3 and 4, the spherical protuberance 3121 is rotated in the spherical through slot relative to the second plate 313 so that the first plate 311 is tilted from 0 degree to a first angle relative to the second plate 313, the first angle refers to a maximum value of the angle at which the first plate 311 is tilted. Under the above configuration, the size of the first angle is defined by the spherical protrusions 3121 and the spherical through grooves, when the spherical protrusions 3121 are assembled in the spherical through grooves, a solid lubricant is applied between the spherical protrusions 3121 and the spherical through grooves to reduce the friction force between the spherical protrusions 3121 and the surfaces of the spherical through grooves, and meanwhile, in order to prevent the leakage of the solid lubricant, an organic lamellar crystalline compound is selected as a lubricant, such as an organic polymer compound of polytetrafluoroethylene, polyimide, and the like, so that a good lubricating effect can be achieved.
Preferably, the inclination of the first plate-shaped element 311 is limited by the limiting module 32, i.e. the size of the first angle is further controlled, referring to fig. 5, which shows a schematic structural diagram of the limiting module 32, wherein the limiting module 32 includes a spherical recess 321, specifically, the spherical recess 321 refers to a recess disposed inside the limiting module 32, the recess is configured to have a hemispherical bottom, referring to fig. 3 to 7, the second plate-shaped element 313 is fixedly mounted on the top of the limiting module 32 to cover the spherical recess 321, after the rod-shaped element 312 passes through the second plate-shaped element 313, the rod-shaped element 312 extends into the spherical recess 321, the bottom end of the rod-shaped element 312 is connected with a rigid ball 3123 through a spring 3122, and the rod-shaped element 312 is configured to: the rigid ball 3123 is in contact with the surface of the spherical recess 321, and the rigid ball 3123 and the spherical recess 321 are capable of sliding relative to each other. Referring to fig. 6, when the first plate 311 is not inclined, that is, the rod 312 is in an upright position relative to the limiting module 32, the length of the spring 3122 is a natural length thereof, and the spring 3122 is not pressed by the bottom of the spherical recess 321; when the first plate 311 tilts, the rod 312 tilts as well, the bottom end of the rod 312 rises to bring the rigid ball 3123 to slide upward along the surface of the spherical groove 321, and the length of the rod 312 extending into the spherical groove 321 remains constant because the spherical protrusion 3121 fits into the spherical through groove, so as the rigid ball 3123 slides upward, the rigid ball 3123 is pressed by the surface of the spherical groove 321, further, the spring 3122 is pressed by the surface of the spherical groove 321, so that the spring 3122 is compressed to accumulate elastic potential energy, accordingly, the rigid ball 3123 is also pressed by the spring 3122 in the direction of the length of the spring 3122, see fig. 7, the pressing force F is a force Fx along the radius of the spherical groove 321, and a force Fy downward in the tangential direction, wherein the force Fy causes the rigid ball 3123 to have a tendency to move downward along the spherical groove 321, so as to return to the lowest point of the spherical groove 321, so that the rigid ball 3123 returns to the vertical position of the rigid ball 3123, and the rigid ball 3123 returns to the vertical position after the rigid ball 3123 has a natural tendency to move downward, so that the rigid ball 3123 returns to the upright position, and the rigid ball 3123 is returned to the upright position, thereby restoring the rigid ball 3123 to the upright position after the rigid ball upright position of the rigid ball 3123, and the rigid ball upright position to the upright position, and the rigid ball is reached the upright position; referring to fig. 6, the capturing section 30 is configured to: when the first plate 311 tilts until the rigid ball 3123 is located at the hemispherical edge of the spherical recess 321, the rigid ball 3123 contacts the bottom end of the rod 312, i.e., the spring 3122 can no longer be compressed, and the surface of the spherical recess 321 restricts the rigid ball 3123 from further sliding upward, thereby restricting the rod 312 and the first plate 311 from further tilting, and achieving the purpose of adjusting the first angle. Specifically, adjustment of the magnitude of the first angle is achieved by varying any one or more of the length of the rod-like member 312 extending into the spherical recess 321, the diameter of the rigid ball 3123, and the diameter of the hemispherical bottom of the spherical recess 321, preferably, the capturing section 30 is configured such that when the rigid ball 3123 is located at the hemispherical edge of the spherical recess 321, the first plate-like member 311 is inclined at an angle of 30 °, i.e., the first angle is 30 °. Preferably, the surface of the spherical groove 321 is also coated with the solid lubricant.
In the docking apparatus 1 provided by the present invention, the target spacecraft is attracted or repelled by the capturing section 30, the capturing section 30 includes a first plate 311 provided with an electromagnetic capturing unit 3112, see fig. 8, which shows a schematic diagram of the first plate 311, the first plate 311 is provided with a tracking unit 3111 for acquiring the position information and the attitude information and an electromagnetic capturing unit 3112 for generating a magnetic field, see fig. 9, which shows a structural schematic diagram of the tracking unit 3111, the tracking unit 3111 includes a signal light emitter 31111 and a signal light receiver 31112, the signal light emitter 31111 is used for emitting signal light to the target spacecraft, and preferably, the signal light emitter 31111 is composed of a group of LED lights. The signal light receiver 31112 is configured to receive the signal light processed by the target spacecraft, and analyze a current position of the target spacecraft and an attitude of the target spacecraft from the processed signal light, and preferably, the signal light receiver 31112 is configured by a binocular camera, and the binocular camera can recognize an attitude angle of a target object (the target spacecraft) according to a parallax principle and transmit information back to the control section 10. Preferably, in order to reduce an error in which the signal light receiver 31112 recognizes the signal light, the signal light receiver 31112 is disposed at the center of the first plate-like member 311.
In another embodiment of the present invention, in order to protect the precise optical components in the signal light receiver 31112, the capture module 31 further includes a cover plate 33 for protecting the first plate 311, and the cover plate 33 is provided with a hole for passing the signal light. The cover plate 33 is made of an energy absorbing material and can absorb impact energy at the first time when the butt joint is established. In addition, if the main spacecraft needs to be stably connected to the target spacecraft for a long time without separation, or the suction force of the docking device 1 cannot provide stable connection to the target spacecraft, an adhesive substance may be applied to the cover plate 33, and stable connection after docking may be further ensured by adhesion.
The first plate 311 is further provided with at least one electromagnetic capture unit 3112, the at least one electromagnetic capture unit 3112 is circumferentially and uniformly arranged on the surface of the first plate 311, each electromagnetic capture unit 3112 of the at least one electromagnetic capture unit 3112 can work independently to enhance the flexibility of docking, and when the tracking unit 3111 is located at the center of the first plate 311, the at least one electromagnetic capture unit 3112 is arranged in a manner surrounding the tracking unit 3111. Referring to fig. 10, a schematic structural diagram of an electromagnetic capture unit 3112 is shown, wherein the electromagnetic capture unit 3112 includes a permanent magnet 31121 and an electromagnet 31122, the permanent magnet 31121 and the electromagnet 31122 are connected and fixed to each other as a whole, and preferably, the electromagnetic capture unit 3112 further includes a resin interlayer 31123 disposed between the permanent magnet 31121 and the electromagnet 31122. Permanent magnet 31121 adopts the neodymium iron boron material, and the upper and lower surface is the S utmost point and the N utmost point respectively, can produce the perpendicular to first plate-like piece 311' S first magnetic field permanent magnet 31121 with on the target spacecraft produce first adsorption affinity between the device of taking magnetism, for the convenience of explanation, electromagnetism capture unit 3112 with electromagnetic action between the target spacecraft all indicates electromagnetism capture unit 3112 with take electromagnetic action between the device of magnetism. The neodymium iron boron has strong magnetism and coercive force, can provide larger basic attractive force when adsorbing the target spacecraft, and does not cause larger change of the strength of the first magnetic field because the second magnetic field generated after the electromagnet 31122 is electrified, preferably, the neodymium iron boron magnet with coercive force larger than 1000KA/m is selected, and the residual magnetism is about 1.3-1.5T.
The electromagnet 31122 includes a coil and an iron core, the coil is wound around the iron core, and the iron core is separated from the permanent magnet 31121 by resin or other non-metallic substances, so that it is avoided that the permanent magnet 31121 magnetizes the iron core by a large margin, and when a metal substance is used as an interlayer, eddy current is generated due to high intensity change of a magnetic field, which causes energy loss, heat generation, and the like. The coil may be energized to generate a second magnetic field, which is also perpendicular to the first plate 311, the direction of the second magnetic field being determined by the direction of the current passed into the coil. The electromagnet 31122 has a first magnetic pole state and a second magnetic pole state, and in the first magnetic pole state, forward current is supplied to the coil, the direction of the magnetic field generated by the electromagnet 31122 is the same as that of the permanent magnet 31121, and the second magnetic field generates a second attractive force between the electromagnet 31122 and the target spacecraft; in the second magnetic pole state, the coil is energized with a reverse current, the direction of the magnetic field generated by the electromagnet 31122 is opposite to the direction of the magnetic field of the permanent magnet 31121, and at this time, the second magnetic field generates a first repulsive force between the electromagnet 31122 and the target spacecraft. Preferably, referring to fig. 8, 12 electromagnetic capture units 3112 are arranged on the first plate 311, the 12 electromagnetic capture units 3112 are uniformly arranged on the surface of the first plate 311 along the circumferential direction, and each electromagnetic capture unit 3112 can be controlled individually, so as to implement different docking strategies and enhance the flexibility of docking.
Specifically, when current is applied to the coil, the magnetic field strength of the second magnetic field can be adjusted by changing the magnitude of the current, so as to adjust the magnitude of the first repulsive force or the second attractive force, illustratively, for convenience of control, the current applied to the coil is divided into 0 gear, first gear and second gear, which are respectively 0%,50% and 100% of the peak current, when the coil is applied with a forward current of the first gear, the coil and the iron core can generate a magnetic field with the same strength and the same direction as the permanent magnet 31121, and at this time, the magnitude of the second attractive force is equal to the magnitude of the first attractive force; when the coil is electrified with reverse second-gear current, the coil and the iron core can generate a magnetic field which is opposite to the direction of the magnetic field of the permanent magnet 31121 and has the strength larger than the first magnetic field strength, and at the moment, the first repulsive force is larger than the first adsorption force.
The docking apparatus 1 includes a control section 10 configured to control the main spacecraft to approach the target spacecraft according to the attitude information and the position information and issue an electromagnetic adjustment instruction to the capture module 31, where the control section 10 includes a control unit configured to receive the position information and the attitude information and issue an electromagnetic adjustment instruction to the capture module 31 based on the position information and the attitude information, and the capture module 31 controls the electromagnet 31122 to switch between the first magnetic pole state and the second magnetic pole state, selects the amount of current supplied to the coil, and tilts the first plate 311 according to the electromagnetic adjustment instruction, thereby controlling the electromagnet 31122 to adsorb or repel the target spacecraft. The control unit is electrically connected with the capture module 31 through a cable, so that the control unit receives the position information and the posture information and sends out the electromagnetic adjusting instruction. The control unit also controls the main spacecraft to approach the target spacecraft and adjusts the attitude of the main spacecraft according to the position information and the attitude information.
When the capture module 31 receives the electromagnetic adjustment command, the first plate 311 tilts to compensate for the angular difference between the first plate 311 and the target spacecraft, so as to avoid collision or unstable connection caused by the angular difference, thereby achieving docking with a larger tolerance. An electromagnet 31122 in the at least one electromagnetic capture unit 3112 is switched between a first magnetic pole state and a second magnetic pole state, so as to realize docking or undocking of the main spacecraft and the target spacecraft, specifically, when the main spacecraft is docked with the target spacecraft, the electromagnetic capture unit 3112 is supplied with a second gear forward current, and the capture section 30 adsorbs the target spacecraft through strong adsorption force; when the main spacecraft and the target spacecraft are docked, the electromagnetic capture unit 3112 is powered off, and the target spacecraft is held only by the permanent magnets 31121, so that energy can be saved; when the main spacecraft is separated from the target spacecraft, the electromagnetic capture unit 3112 passes a second gear reverse current, and the capture section 30 repels the target spacecraft to achieve separation of the main spacecraft from the target spacecraft.
In order to prevent damage to the spacecraft or devices on the spacecraft during docking, which may be caused by a weak impact of the main spacecraft with the target spacecraft by the action of electromagnetic force, the docking device 1 is further provided with a buffer section 20 for absorbing impact, the buffer section 20 being composed of a bellows configured to absorb impact when the capturing section 30 is docked with the target spacecraft, the bellows being configured to enclose a flexible cable for connecting the capturing section 30 with the control section 10.
In the process of establishing the docking between the main spacecraft and the target spacecraft by the docking device 1, the docking is mainly divided into target identification, target approach, target capture, target docking and target release. Referring to fig. 11, there is shown a flow chart of a process for docking the master spacecraft with the target spacecraft, wherein the docking process comprises:
(111) Target identification: position and attitude recognition is carried out on the target spacecraft through a tracking unit (a binocular camera) in a capturing module in the aligning device, and recognized information is fed back to the main spacecraft;
(112) The target approaches: the main spacecraft uses a posture and orbit control single machine of the main spacecraft to approach the target spacecraft and carries out posture adjustment;
(113) Target capture: the electromagnetic capture unit in the capture module is electrified with forward current to carry out adsorption capture on the target spacecraft so as to establish butt joint;
(114) Target butt joint: after the main spacecraft is in butt joint with the target spacecraft, the electromagnetic capturing unit is powered off and is stably connected by means of a permanent magnet (and viscous substances on a cover plate) in the electromagnetic capturing unit;
(115) Target release: when the target spacecraft needs to be released, the electromagnetic capturing unit is electrified with reverse current, an electromagnet in the electromagnetic capturing unit generates a magnetic field opposite to the direction of the permanent magnet, the magnetic field strength of the electromagnetic capturing unit is enabled to be zero so as to stop adsorbing the target spacecraft, or the magnetic field direction of the electromagnetic capturing unit is enabled to be changed so as to repel the target spacecraft, and meanwhile the main spacecraft is maneuvered to realize the separation from the target spacecraft.
The invention discloses an on-orbit docking device for a spacecraft, which is arranged on a main spacecraft, the installation number and the installation positions of the docking devices 1 can be selected according to different task requirements, and based on the on-orbit docking device, the invention also discloses a docking system 2 for the spacecraft, wherein the docking system 2 comprises a spacecraft main body and at least one docking device 1, the spacecraft main body is constructed into a cube structure, the docking system 2 comprises at least one docking device 1, and refer to the attached drawings 12-13, wherein the docking system 2 comprises four docking devices 1. The docking device 1 is rotatably connected with the main body of the spacecraft, the docking device 1 can rotate in space by any angle within 0 ° to 180 °, specifically, the rotatable connection of the docking device 1 with the main body of the spacecraft means that: the docking device 1 is connected with the spacecraft main body through a cylindrical rotary joint 3, each docking device 1 can rotate in a fixed plane within a range of 180 degrees, and the docking system 2 can control the rotation angle of each docking device 1 respectively.
Further, the docking system 2 is configured to: when the docking devices 1 do not rotate, each docking device 1 is accommodated in the spacecraft main body in a mutually parallel configuration, specifically, referring to fig. 12 to 13, the four docking devices 1 are respectively arranged on four mutually parallel edges of the cube when not rotating, and through the configuration, the structural stability of the spacecraft in the launching and flying processes can be improved, the installation envelope of the spacecraft can be reduced, and meanwhile, the influence of a magnetic field at the tail end of the docking device 1 on the spacecraft main body can be reduced when docking is established.
It should be noted that: the technical schemes described in the embodiments of the present invention can be combined arbitrarily without conflict.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention, and all the changes or substitutions should be covered within the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the appended claims.
Claims (10)
1. In-orbit docking device for a spacecraft to be mounted on a main spacecraft for docking a target spacecraft, characterized in that the docking device is a rod-like mechanism, divided into a control section, a buffer section and a capture section, wherein,
the capturing section comprises a limiting module and a capturing module capable of inclining towards any direction, the capturing module is configured to acquire attitude information and position information of the target spacecraft and is in butt joint with or separated from the target spacecraft through electromagnetic force, and the limiting module is configured to limit the capturing section;
the control section comprises a control unit which is configured to control the main spacecraft to approach the target spacecraft according to the attitude information and the position information and send an electromagnetic adjusting instruction to the capture module to control the capture module to adsorb or repel the target spacecraft;
the buffer section is configured to absorb an impact when the master spacecraft is docked with the target spacecraft.
2. The docking device as claimed in claim 1, wherein the capturing module comprises:
a first plate on which a tracking unit configured to attract or repel the target spacecraft with different strengths of electromagnetic force and at least one electromagnetic capturing unit configured to acquire the position information and the attitude information in optical communication are disposed;
the rod-shaped piece is fixedly connected to the bottom of the first plate-shaped piece, and a spherical bulge is arranged at the middle section of the rod-shaped piece;
a second plate provided with a spherical through slot through which the rod passes such that the spherical protrusion fits in the spherical through slot, the capture module being configured to: the spherical protrusion can rotate in the spherical through groove relative to the second plate-shaped piece, so that the first plate-shaped piece can incline by 0 degrees to a first angle in any direction, and the first angle refers to the maximum angle value of the inclination of the first plate-shaped piece.
3. The docking device as claimed in claim 2, wherein the bottom end of the rod is connected with a rigid ball by a spring, the position-limiting module comprises a spherical recess, the second plate is mounted on the top of the position-limiting module to cover the spherical recess, the rod extends into the spherical recess to make the rigid ball always abut against the spherical surface of the spherical recess, and the capturing section is configured to: when the rod-shaped piece is not inclined, the spring is in an original state; when the rod-shaped piece tilts, the spring is in a compressed state; when the rod is inclined to the first angle, the rigid ball is positioned at the edge of the spherical groove, and the rigid ball is in contact with the bottom end of the rod.
4. The docking device as recited in claim 2, wherein the tracking unit comprises a signal light emitter and a signal light receiver, the tracking unit configured to: the signal light emitter emits signal light to the target spacecraft, the signal light is processed by the target spacecraft and then captured and analyzed by the signal light receiver, so that the position information and the attitude information are obtained, and the tracking unit transmits the position information and the attitude information to the control unit.
5. Docking device according to claim 2, characterized in that the at least one electromagnetic capture unit is arranged circumferentially uniformly on the first plate-like member, the electromagnetic capture unit comprising coaxially arranged electromagnets and permanent magnets, the permanent magnets being configured to attract the target spacecraft with a first attracting force, the electromagnets being configured to switch, depending on the electromagnetic adjustment command, between a first magnetic pole state in which the electromagnets attract the target spacecraft with a second attracting force and a second magnetic pole state in which the electromagnets repel the target spacecraft with a first repelling force.
6. The docking device as recited in claim 5, wherein the electromagnet comprises a coil and a core, and the direction and intensity of current passing through the coil can be adjusted.
7. The docking device of claim 1, wherein the buffer section comprises a bellows configured to enclose a flexible cable for connecting the capture section and the control section.
8. A docking system for a spacecraft, the docking device comprising any one of claims 1 to 7, wherein the docking system comprises a spacecraft body and at least one of the docking devices, the spacecraft body being configured as a cube structure, the docking system being configured to: the docking device is rotatably connected with the spacecraft main body and can rotate within the space by any angle within the range of 0-180 degrees.
9. The docking system of claim 8, wherein the at least one docking device is configured to: each of the at least one docking device is parallel to each other when the docking device is not rotated.
10. The docking system of claim 8, wherein each of the at least one docking device is independently controllable.
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| CN202211631205.0A CN115610713A (en) | 2022-12-19 | 2022-12-19 | In-orbit docking device and docking system for spacecraft |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20120076629A1 (en) * | 2011-12-06 | 2012-03-29 | Altius Space Machines | Sticky Boom Non-Cooperative Capture Device |
| CN105000200A (en) * | 2015-07-24 | 2015-10-28 | 北京空间飞行器总体设计部 | Flexible misjudgment preventing butt joint rod with signal feedback |
| CN108100311A (en) * | 2017-12-21 | 2018-06-01 | 星际漫步（北京）航天科技有限公司 | Microsatellite separator and its method for releasing |
| US20190367192A1 (en) * | 2017-02-15 | 2019-12-05 | Astroscale Japan Inc. | Capturing system, aerospace vehicle, and plate-like body |
| CN111994306A (en) * | 2020-07-23 | 2020-11-27 | 北京空间飞行器总体设计部 | High-precision electromagnetic docking mechanism with large-angle tolerance |
| US20220332443A1 (en) * | 2021-04-19 | 2022-10-20 | Roopnarine | Servicing systems for on-orbit spacecrafts |
-
2022
- 2022-12-19 CN CN202211631205.0A patent/CN115610713A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20120076629A1 (en) * | 2011-12-06 | 2012-03-29 | Altius Space Machines | Sticky Boom Non-Cooperative Capture Device |
| CN105000200A (en) * | 2015-07-24 | 2015-10-28 | 北京空间飞行器总体设计部 | Flexible misjudgment preventing butt joint rod with signal feedback |
| US20190367192A1 (en) * | 2017-02-15 | 2019-12-05 | Astroscale Japan Inc. | Capturing system, aerospace vehicle, and plate-like body |
| CN108100311A (en) * | 2017-12-21 | 2018-06-01 | 星际漫步（北京）航天科技有限公司 | Microsatellite separator and its method for releasing |
| CN111994306A (en) * | 2020-07-23 | 2020-11-27 | 北京空间飞行器总体设计部 | High-precision electromagnetic docking mechanism with large-angle tolerance |
| US20220332443A1 (en) * | 2021-04-19 | 2022-10-20 | Roopnarine | Servicing systems for on-orbit spacecrafts |
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