WO2023179403A1 - 一种用于空间目标的捕捉系统及方法 - Google Patents

一种用于空间目标的捕捉系统及方法 Download PDF

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Publication number
WO2023179403A1
WO2023179403A1 PCT/CN2023/081155 CN2023081155W WO2023179403A1 WO 2023179403 A1 WO2023179403 A1 WO 2023179403A1 CN 2023081155 W CN2023081155 W CN 2023081155W WO 2023179403 A1 WO2023179403 A1 WO 2023179403A1
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WO
WIPO (PCT)
Prior art keywords
capture
transmission
capturing
deceleration
launch
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Application number
PCT/CN2023/081155
Other languages
English (en)
French (fr)
Inventor
郭金生
吴凡
邱实
陈健
魏承
王宏旭
Original Assignee
哈尔滨工业大学
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
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Application filed by 哈尔滨工业大学 filed Critical 哈尔滨工业大学
Publication of WO2023179403A1 publication Critical patent/WO2023179403A1/zh
Priority to US18/390,259 priority Critical patent/US20240116655A1/en

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • B64G1/1078—Maintenance satellites
    • B64G1/1081—Maintenance satellites for debris removal
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41H—ARMOUR; ARMOURED TURRETS; ARMOURED OR ARMED VEHICLES; MEANS OF ATTACK OR DEFENCE, e.g. CAMOUFLAGE, IN GENERAL
    • F41H11/00—Defence installations; Defence devices
    • F41H11/02—Anti-aircraft or anti-guided missile or anti-torpedo defence installations or systems
    • 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

Definitions

  • the present disclosure relates to the field of aerospace technology, and more specifically to a capturing system and method for space targets.
  • the current methods for capturing failed aircraft in orbit generally include: flying nets, robotic arms, darts, and rendezvous and docking methods; among them, using flying nets to capture failed aircraft is less efficient, and the number of flying nets that the spacecraft can carry at the same time is Less; when using a robotic arm to capture a failed aircraft, the mass of the robotic arm is large and the cost is high.
  • High-precision orbit control is required when performing the capture mission, and the reliability of the robotic arm restricts the reliability of the spacecraft it carries; using The rendezvous and docking method is very demanding and difficult to capture a disabled aircraft.
  • the method of using harpoon to capture a failed aircraft is simple and reliable, but it needs to consider lightweight design and prevent problems such as harpoon rebound and repeated capture.
  • Embodiments of the present disclosure are expected to provide a system and method for capturing space targets that can capture failed space targets and drag the failed space targets to an attenuated orbit after successful capture, with low cost and high capture efficiency.
  • embodiments of the present disclosure provide a capture system for space targets, the capture system including: a plurality of capture devices, a transmission device, a launch device, a deceleration and recovery device; wherein,
  • the plurality of capture devices are used to be launched to the target orbit in sequence to capture failed space targets;
  • the conveying device is used to sequentially feed and move the plurality of capturing devices to the first set position in the launching device along a set conveying trajectory;
  • the launching device is used to launch the capturing device in the first set position to the target orbit to capture the space target;
  • the deceleration and recovery device is used to decelerate the capture device after it is launched and flies for a set distance.
  • inventions of the present disclosure provide a method for capturing space targets.
  • the capturing method can be applied to the capturing system described in the first aspect.
  • the capturing method includes:
  • the conveying device sequentially feeds and moves each capturing device to the first set position in the launching device along the set conveying trajectory;
  • the launching device launches the capturing device in the first set position to the target orbit to capture the failed space target;
  • the deceleration and recovery device When the capture device is launched and flies for a set distance, the deceleration and recovery device performs deceleration control on the capture device.
  • Embodiments of the present disclosure provide a system and method for capturing space targets; the capturing device is sequentially fed and moved along a set transmission trajectory through a transmission device to a first set position in the launching device, so that the capturing device It is launched to the target orbit to capture the failed space target; at the same time, when the capture device is launched and flies for a set distance, the deceleration and recovery device can decelerate the capture device to prevent the capture device from rebounding and damaging the spacecraft on board.
  • the capture system provided by the embodiments of the present disclosure can capture space targets of various sizes and types, including space targets of various sizes and types that are difficult to capture with traditional robotic arms and flying nets. It has strong overall adaptability and high capture efficiency. .
  • Figure 1 is a schematic structural diagram of a capturing system for space targets provided by an embodiment of the present disclosure.
  • FIG. 2 is a schematic structural top view of a capturing system for space targets provided by an embodiment of the present disclosure.
  • Figure 3 is a schematic side view of the structure of a capturing system for space targets provided by an embodiment of the present disclosure.
  • Figure 4 is a schematic structural diagram of a capture device provided by an embodiment of the present disclosure.
  • FIG. 5 is a schematic structural diagram of a heating device installed inside a reduction shaft provided by an embodiment of the present disclosure.
  • Figure 6 is a schematic structural diagram of a transmission device provided by an embodiment of the present disclosure.
  • Figure 7 is a partially enlarged schematic diagram of position I in Figure 6.
  • Figure 8 is a schematic structural diagram of a transmitting device provided by an embodiment of the present disclosure.
  • Figure 9 is a schematic diagram of the internal structure of a launching device provided by an embodiment of the present disclosure.
  • Figure 10 is a schematic structural diagram of the second transmission gear provided by an embodiment of the present disclosure.
  • Figure 11 is a schematic structural diagram of a deceleration and recovery device provided by an embodiment of the present disclosure.
  • Figure 12 is a schematic diagram of a partial structure of the deceleration and recovery device provided by an embodiment of the present disclosure.
  • Figure 13 is a schematic flowchart of a method for capturing space targets provided by an embodiment of the present disclosure.
  • the capture system 1 specifically includes: multiple capture devices 11, a transmission device 12, a launch device 13, a deceleration and recovery device. Device 14; wherein,
  • the plurality of capture devices 11 are used to be launched to the target orbit in sequence to capture failed space targets;
  • the conveying device 12 is used to sequentially feed and move the plurality of capturing devices 11 to the first set position in the launching device 13 along the set conveying trajectory;
  • the launching device 13 is used to launch the capturing device 11 at the first set position to the target orbit to capture the space target;
  • the deceleration and recovery device 14 is used to decelerate the capture device 11 after the capture device 11 is launched and flies for a set distance.
  • disabled space targets are not limited to disabled aircraft, and of course also include space debris.
  • conveying trajectory of the conveying device 12 is not limited to the coiled shape shown in FIGS. 1 and 2 , and conveying trajectories of other forms and shapes can also be applied to the capture system 1 in the present disclosure.
  • the capture device 11 is sequentially fed and moved along the set transfer trajectory to the first set position in the launch device 13 through the transfer device 12 , so that the capture device 11 is launched to the target orbit.
  • the deceleration and recovery device 14 can decelerate the capture device 11 to prevent the capture device 11 from rebounding and damaging the spacecraft on board.
  • the capture system 1 provided by the embodiment of the present disclosure can capture space targets of various sizes and types, including space targets of various sizes and types that are difficult to capture with traditional robotic arms and flying nets. It has strong overall adaptability and high capture efficiency. high.
  • each component in the capture device 11 in the embodiment of the present disclosure are made of degradable materials, so the capture device 11 and the failed space target can be fully degraded after being dragged to the attenuation orbit.
  • the capture system 1 can be installed on the surface of other spacecraft (not shown in the figure), such as a satellite, to fly to a set space orbit by carrying the satellite. , to perform the capture task.
  • spacecraft not shown in the figure
  • the capture system 1 can be installed on the surface of other spacecraft (not shown in the figure), such as a satellite, to fly to a set space orbit by carrying the satellite. , to perform the capture task.
  • the hit failed space target can fly to the set attenuation orbit together with the spacecraft carried by the capture system, thereby achieving the failed space target of the target orbit. Drag the task to the set attenuation track.
  • 200 capture devices 11 can be installed in the capture system 1 shown in Figure 1 to perform 200 capture missions, thereby avoiding the ground station from launching the capture system 1 multiple times. Therefore, Can save launch costs.
  • the capture system 1 provided by the embodiment of the present disclosure can carry a 300Kg mass class satellite to perform capture missions in orbit, with low cost and high efficiency.
  • the number of capturing devices 11 can also be changed according to actual conditions, and is not limited to the 200 capturing devices 11 described in the foregoing technical solution.
  • the capture system 1 shown in Figure 1 when the capture system 1 shown in Figure 1 is assembled, as shown in Figures 2 and 3, the maximum radial diameter of the capture system 1 does not exceed 1000 mm, and the maximum height does not exceed 210 mm. Therefore, the capture system 1 in the embodiment of the present disclosure is small in size and light in weight, realizing a lightweight design of the capture system 1 .
  • the deceleration and recovery device 14 is also used to recover the capture device 11 into the launch device 13 when the capture device 11 misses the target. It can be understood that when the capture device 11 fails to hit a space target, the capture device 11 can be re-launched to other target orbits after being recovered to the launch device 13 to continue to search for other failed space targets for capture and dragging.
  • each capture device 11 includes: a head 111, at least one barb unit 112, a positioning unit 113, a The first rope 114, the deceleration shaft 115 and the second rope 116 wrapped around the deceleration shaft 115 inside the capture device 11; wherein,
  • the head 111 is pointed and is used to penetrate and insert into the space target;
  • the barb unit 112 is used to prevent the space target from falling off during towing;
  • the positioning unit 113 is used to position the feed movement position of the capture device 11 in real time to ensure that the capture device 11 can feed and move to the first set position;
  • the second rope 116 is connected to the first rope 114 and is used to decelerate the catching device 11 until the catching device 11 decelerates to zero after the first rope 114 is fully deployed.
  • each component in the capture device 11, for example, the head 111, the barb unit 112 and the positioning unit 113 can be detached from the capture system 1 to be launched to the target orbit.
  • the first rope 114 can gradually detach from the capturing device 11 and gradually unfold.
  • the capture device 11 continues to drive the first rope 114 to fly forward until the first rope 114 is fully deployed; when the first rope 114 is fully deployed, the first rope 114 is fully deployed.
  • One rope 114 will pull the second rope 116 to gradually unfold to control the deceleration of the catching device 11 .
  • the first rope 114 when the first rope 114 is fully deployed and the capture device 11 does not capture the space target, the first rope 114 can also pull the second rope 116 to gradually deploy to control the deceleration of the capture device 11 .
  • the first rope 114 and the second rope 116 can be the same rope, but designed in sections, and the first rope 114 is wound inside the catching device 11 itself, and the second rope 116 is wound around the inside of the catching device 11 itself.
  • On the deceleration shaft 115 such a design can not only help the capture device 11 hit the failed space target, but also effectively control the deceleration of the capture device 11 to prevent the capture device 11 from rebounding and damaging the satellite on board.
  • the mass of the capture device 11 can be designed to be 0.5Kg to achieve a lightweight design of the capture system 1; and the length of the first rope 11 can be designed to be 50 meters to achieve a lightweight design of the capture system 1. Capture disabled space targets within a range of 50 meters. On the other hand, it can be understood that the set distance that the capturing device 11 described in the aforementioned technical solution can fly after being launched is determined by the length of the first rope 11 .
  • a heating device 1151 is provided inside the deceleration shaft 115 .
  • the heating device 1151 is used when the space target is dragged to a set attenuation orbit. Heating is then performed to fuse the second rope 116 .
  • the heating device 1151 may be a power resistor.
  • the transmission device 12 includes: a first frame 121, a plurality of mounting supports 122, and a plurality of deceleration shaft holes. 123.
  • the plurality of mounting brackets 122 are evenly arranged on the upper side of the first frame 121; wherein each of the capturing devices 11 is correspondingly installed on each of the mounting brackets 122;
  • the plurality of deceleration shaft holes 123 are evenly and symmetrically arranged on both sides of the first frame 121; wherein the deceleration shaft 115 is correspondingly installed in the deceleration shaft hole 123 in a rotatable manner;
  • a T-shaped guide rail 1211 is provided at the bottom of the first frame 121. As shown in Figure 7, first transmission racks 128 meshing with the first transmission gear 124 are provided on both sides of the T-shaped guide rail 1211;
  • the first driving mechanism 125 is used to drive the first transmission gear 124 to mesh with the first transmission rack 128 so that a plurality of the capturing devices 11 move sequentially along the set transmission trajectory;
  • the plurality of power supply contacts 126 are evenly arranged on both sides of the first frame 121 to provide electrical energy during the launch and deceleration recovery process of the capture device 11;
  • the plurality of switch contacts 127 are evenly arranged on one side of the first frame 121 .
  • the first driving mechanism 125 drives the first transmission gear 124 to engage with the first transmission rack 128 for transmission, so that the plurality of capturing devices 11 move sequentially until they enter the launching device 13 First setting position. It can be understood that during the feeding movement of the capturing device 11, the first driving mechanism 125 and the first transmission gear 124 do not move, and the first frame 121 and other components provided on the first frame 121, such as the mounting bracket, 122. The deceleration shaft hole 123, the power supply contact 126 and the switch contact 127 will move simultaneously with the capture device 11.
  • the transmission device 12 is assembled together with the mounted satellite through the T-shaped guide rail 1211.
  • the transmission device 12 includes a plurality of transmission units, and each transmission unit correspondingly includes a mounting bracket 122, two reduction shaft holes 123 to install the reduction shaft 115, two power supply contacts 126 and a switch. Contact 127; each capture device 11 corresponds to a transfer unit. During the transfer process of the capture device 11, each transfer unit moves simultaneously with the corresponding capture device 11, and cooperates to complete the launch of the capture device 11. Deceleration and recovery tasks.
  • the first driving mechanism 125 may be a stepper motor.
  • the transport trajectory of the capturing device 11 depends on the shape of the first frame 121, such as the coiled shape in FIG. 1 .
  • the launch device 13 includes: a second frame 131, a launch chamber 132, a launch support 133, The second driving mechanism 134, the second transmission mechanism 135, the first elastic mechanism 136, the guide rail 137, the first travel switch 138, the second travel switch 139 and the conductive sheet 140; wherein,
  • the launch chamber 132 is semi-cylindrical and is arranged in the second frame 131;
  • the launch support 133 passes through the launch chamber 132 and is fixedly connected to the launch chamber 132;
  • the second driving mechanism 134 is used to drive the second transmission mechanism 135 for transmission; wherein, when the capturing device 11 has not been fed to the first set position, the second driving mechanism 134 drives the The second transmission mechanism 135 drives so that the launch support 133 and the launch chamber 132 move down to the second set position along the guide rail 137 and compress the first elastic mechanism 136; and, when the When the capturing device 11 has been fed to the first set position, the second driving mechanism 134 drives the second transmission mechanism 135 to continue transmission so that the launching support 135 and the launching chamber 132 are at the same position. Under the action of the first elastic mechanism 135, it moves upward along the guide rail 137 and pushes the capture device 11 to be launched;
  • the first travel switch 138 is used to send an instruction to the second transmission mechanism 135 to stop transmission when the launch support 133 and the launch chamber 132 move down to the second setting position, so that the The launch support 133 is maintained at the second set position;
  • the second travel switch 139 is used to electrically connect with the switch contact 127 to send the capturing device 11 to stop the feeding movement when the catching device 11 has been moved to the first set position. Instructions to keep the capturing device 11 at the first set position;
  • the conductive sheet 140 is disposed on both sides of the second frame 131 and is electrically connected to the power contact 126 to provide electrical energy.
  • the second driving mechanism 134 may be a stepper motor connected to a power supply device (not shown in the figure) on the satellite.
  • the first elastic mechanism 136 can be specifically a spring, which passes through the launch support 133, and the first elastic mechanism 136 can be compressed during the downward movement of the launch support 133 and the launch chamber 132; In addition, after the capturing device 11 is advanced and moved to the first set position, the bottom of the first elastic mechanism 136 can contact the top of the mounting bracket 122 .
  • the second transmission mechanism 135 includes a second transmission gear 1351 and a second transmission rack 1352 meshed with the second transmission gear 1351; and the second transmission gear 1351 is a half Tooth structure.
  • the second transmission gear 1351 is connected to the second driving mechanism 134 ; the second transmission rack 1352 is fixedly provided on one side of the launch support 133 . Therefore, when the capture device 11 reaches the first set position and aims at the space target to be captured, the second drive mechanism 134 drives the second transmission mechanism 135 for transmission.
  • the second transmission gear 1351 Since the second transmission gear 1351 has a half-tooth structure, during the transmission process The second transmission rack 1352 will be released, and then the launch support 133 and the launch chamber 132 will move upward under the elastic force of the first elastic mechanism 136 and push the capture device 11 out of the launch device 13 to complete the launch action.
  • the spring force of the first elastic mechanism 136 is 2000N, which can enable the 0.5kg capture device 11 to obtain a flight speed of 20m/s, thereby enabling the capture device 11 to penetrate and insert into a space target at high speed, such as a failed aircraft. honeycomb sandwich structure.
  • the second transmission mechanism 135 further includes a ratchet 1353 and a pawl 1354; wherein the ratchet 1353 is disposed inside the second transmission gear 1351,
  • the pawl 1354 is connected to the second driving mechanism 134 .
  • the power supply device on the satellite provides electric energy to the second driving mechanism 134, and then the second driving mechanism 134 drives the second transmission gear 1351 and the second transmission tooth.
  • the strip 1352 engages the transmission, thereby moving the launch support 133 downward to the second set position, and compresses the first elastic mechanism 136 through the downwardly moved launch support 133, and simultaneously triggers the first travel switch 138; when the first travel switch 138 After being triggered, the second driving mechanism 134 stops running. At this time, the pawl 1354 resists the outer peripheral wall of the ratchet wheel 1353 so that the second transmission gear 1351 stops transmitting and remains at the current transmission position, thereby causing the launch support 133 to contact the launch chamber. 132 remains fixed at the second set position.
  • the positioning unit 113 in the capture device 11 can abut against the launch chamber 132 , so the elastic force of the first elastic mechanism 136 Under the action, the capture device 11 is launched to the target orbit.
  • the conveying device 12 will continue to feed and move along the set conveying trajectory.
  • the captured device being launched will The corresponding installation parts of 11 such as the installation bracket 122 and other parts such as the reduction shaft 115 are moved out of the launcher 13; on the other hand, the next capture device 11 and its corresponding installation parts will be fed and moved into the launcher 13 to On to the next capture mission.
  • the deceleration and recovery device 14 includes: a solenoid valve 141, a drive shaft 142, and a winding coil around the drive shaft 142. the second elastic mechanism 143, the magnetic damper 144 and the third driving mechanism 145; wherein,
  • the solenoid valve 141 is used to stop working when the catching device 11 moves to the first set position so that the second elastic mechanism 143 pushes the driving shaft 142 and the reduction shaft. 115 connection; and, after the capture device 11 is launched, the drive shaft 142 is separated from the reduction shaft 115 by attracting the drive shaft 142; wherein, the drive shaft 142 passes through a gear coupling 146 is connected to the reduction shaft 115;
  • the magnetic damper 144 is used to generate a damping force when the first rope 114 pulls the second rope 116 and gradually unfolds it to control the deceleration of the capture device 11;
  • the third driving mechanism 145 is connected to the magnetic damper 144 and is used to drive the reduction shaft 115 to rotate in reverse direction to recover the capturing device 11 by pulling the second reduction gear 116 .
  • the capture device 11 moves to the first set position and triggers the second travel switch 139, the capture device 11 stops the feed movement, and the solenoid valve 141 also stops working, so as to pass the second elasticity.
  • the mechanism 143 promotes the connection between the driving shaft 142 and the reduction shaft 115. At this time, the damping torque is transmitted between the driving shaft 142 and the reduction shaft 115 through the gear coupling 146; it can be understood that when the capture device 11 is launched, the first After the first rope 114 gradually and completely breaks away from the catching device 11, the first rope 114 continues to pull the second rope 116.
  • the deceleration shaft 115 Rotation occurs; on the other hand, the reduction shaft 115 is connected to the drive shaft 142 through the gear coupling 146, so when the magnetic damper 144 generates a damping force, its damping torque can be transmitted to the reduction shaft 115 through the drive shaft 142, so During the specific implementation process, the magnitude of the damping moment can be obtained by measuring the real-time angular velocity of the deceleration shaft 115, so as to decelerate the flight speed of the capture device 11 to zero before the second rope 116 is fully deployed; at the same time, when the capture device 11 hits the space target Finally, the space target can be towed to the attenuation orbit by the spacecraft carried by the capture system 1. It can be understood that when the failed space target is dragged to the attenuation orbit, it can be heated by the heating device 117 on the deceleration shaft 115, and finally the
  • the gear coupling 146 may include a bevel friction gear 1461 and a driven helical gear 1462.
  • the third driving mechanism 145 may be a stepper motor.
  • the connection method between the third driving mechanism and the magnetic damper 144 is not limited to the series connection in FIG. 10 , and can also be designed as an integrated design.
  • FIG. 13 it shows a capturing method for space targets provided by an embodiment of the present disclosure.
  • the capturing method can be applied to the capturing system 1 described in the aforementioned technical solution.
  • the capturing method specifically includes:
  • the transmission device sequentially feeds and moves each capturing device to the first set position in the launching device along the set transmission trajectory;
  • the launching device launches the capturing device at the first set position to the target orbit to capture the failed space target;
  • the deceleration and recovery device performs deceleration control on the capture device.
  • the capturing method further includes:
  • the capture device misses the target, the capture device can be recovered into the launch device.
  • Embodiments of the present disclosure provide a system and method for capturing space targets; the capturing device is sequentially fed and moved along a set transmission trajectory through a transmission device to a first set position in the launching device, so that the capturing device is Launch to the target orbit to capture failed space targets; at the same time, when the capture device is launched and flies for a set distance, the deceleration and recovery device can decelerate the capture device to prevent the capture device from rebounding and damaging the spacecraft on board.
  • the capture system provided by the embodiments of the present disclosure can capture space targets of various sizes and types, including space targets of various sizes and types that are difficult to capture with traditional robotic arms and flying nets. It has strong overall adaptability and high capture efficiency. .

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Abstract

本公开实施例公开了一种用于空间目标的捕捉系统及方法；所述捕捉系统包括：多个捕捉装置、传送装置、发射装置、减速及回收装置；其中，所述多个捕捉装置，用于依次被发射至目标轨道以捕捉失效的空间目标；所述传送装置，用于沿设定的传送轨迹将所述多个捕捉装置依次进给移动至所述发射装置中的第一设定位置；所述发射装置，用于将处于所述第一设定位置的捕捉装置发射至所述目标轨道以捕捉所述空间目标；所述减速及回收装置，用于当所述捕捉装置被发射并飞行设定距离后对所述捕捉装置进行减速控制。

Description

一种用于空间目标的捕捉系统及方法 相关申请的交叉引用
本申请要求于2022年03月24日提交中国专利局、申请号为202210302809.4，发明名称为“一种用于空间目标的捕捉系统及方法”的中国专利申请的优先权，该中国专利申请的内容在此引入本申请作为参考。
技术领域
本公开涉及航天技术领域，更具体地涉及一种用于空间目标的捕捉系统及方法。
 背景技术
目前，随着在轨空间飞行器数量急剧增加，导致在轨失效飞行器的数量也随之剧增，因此如何捕捉并清除这些失效飞行器已然成为航天领域亟需解决的热点问题之一。目前捕捉在轨失效飞行器的方式大致包括：飞网、机械臂、鱼镖以及交会对接等方式；其中，采用飞网捕捉失效飞行器的方式效率较低，同时搭载的航天器能够携带的飞网数量较少；采用机械臂捕捉失效飞行器时，机械臂的质量较大且成本较高，在执行捕捉任务时需要高精度轨道控制，并且机械臂的可靠性制约了搭载的航天器的可靠性；采用交会对接方式捕捉失效飞行器时要求高，难度也较大；采用鱼镖捕捉失效飞行器的方式简单可靠，但是需要考虑轻量化设计，同时预防鱼镖反弹和重复捕捉等问题。
技术内容
本公开的实施例期望提供一种用于空间目标的捕捉系统及方法；能够捕捉失效的空间目标，并在捕捉成功后可将失效的空间目标拖曳至衰减轨道，成本低且捕捉效率高。
本公开的实施例的技术方案是这样实现的：
第一方面，本公开的实施例提供了一种用于空间目标的捕捉系统，所述捕捉系统包括：多个捕捉装置、传送装置、发射装置、减速及回收装置；其中，
所述多个捕捉装置，用于依次被发射至目标轨道以捕捉失效的空间目标；
所述传送装置，用于沿设定的传送轨迹将所述多个捕捉装置依次进给移动至所述发射装置中的第一设定位置；
所述发射装置，用于将处于所述第一设定位置的捕捉装置发射至所述目标轨道以捕捉所述空间目标；
所述减速及回收装置，用于当所述捕捉装置被发射并飞行设定距离后对所述捕捉装置进行减速控制。
第二方面，本公开的实施例提供了一种用于空间目标的捕捉方法，所述捕捉方法能够应用于第一方面所述的捕捉系统，所述捕捉方法包括：
传送装置沿设定的传送轨迹将每个捕捉装置依次进给移动至发射装置中的第一设定位置中；
所述发射装置将处于所述第一设定位置的所述捕捉装置发射至所述目标轨道以捕捉所述失效的空间目标；
当所述捕捉装置被发射并飞行设定距离后减速及回收装置对所述捕捉装置进行减速控制。
本公开的实施例提供了一种用于空间目标的捕捉系统及方法；通过传送装置将捕捉装置沿设定的传送轨迹依次进给移动至发射装置中的第一设定位置，从而使得捕捉装置被发射至目标轨道以捕捉失效的空间目标；同时，当捕捉装置被发射并飞行设定距离后，减速及回收装置能够对捕捉装置进行减速控制以防止捕捉装置反弹进而损伤搭载的航天器。本公开实施例提供的捕捉系统能够捕捉各种尺寸及类型的空间目标，当然也包括传统机械臂和飞网方式难以捕捉的各种尺寸及类型的空间目标，整体适应性强，且捕捉效率高。
 附图简要说明
为了更清楚地说明本公开实施例的技术方案，下面将对实施例的描述中所需要使用的附图作简单的介绍。下面描述中的附图仅仅是本公开的示例性实施例。
图1为本公开实施例提供的一种用于空间目标的捕捉系统结构示意图。
图2为本公开实施例提供的一种用于空间目标的捕捉系统结构俯视图示意图。
图3为本公开实施例提供的一种用于空间目标的捕捉系统结构侧视图示意图。
图4为本公开实施例提供的捕捉装置结构示意图。
图5为本公开实施例提供的减速轴内部设置有加热装置的结构示意图。
图6为本公开实施例提供的传送装置结构示意图。
图7为图6中I处局部放大示意图。
图8为本公开实施例提供的发射装置结构示意图。
图9为本公开实施例提供的发射装置内部结构示意图。
图10为本公开实施例提供的第二传动齿轮结构示意图。
图11为本公开实施例提供的减速及回收装置结构示意图。
图12为本公开实施例提供的减速及回收装置中局部结构示意图。
图13为本公开实施例提供的一种用于空间目标的捕捉方法流程示意图。
图中：11-捕捉装置；111-头部；112-倒刺单元；113-定位单元；114-第一绳索；115-减速轴；1151-加热装置；116-第二绳索；12-传送装置；121-第一框架；1211-T型导轨；122-安装支座；123-减速轴孔；124-第一传动齿轮；125-第一驱动机构；126-供电触点；127-开关触点；128-第一传动齿条；13-发射装置；131-第二框架；132-发射腔膛；133-发射支座；134-第二驱动机构；135-第二传动机构；1351-第二传动齿轮；1352-第二传动齿条；1353-棘轮；1354-棘爪；136-第一弹性机构；137-导轨；138-第一行程开关；139-第二行程开关；140-导电片；14-减速及回收装置；141-电磁阀；142-驱动轴；143-第二弹性机构；144-磁阻尼器；145-第三驱动机构。
 具体实施方式
为了使得本公开的目的、技术方案和优点更为明显，下面将参照附图详细描述根据本公开的示例实施例。显然，所描述的实施例仅仅是本公开的一部分实施例，而不是本公开的全部实施例，应理解，本公开不受这里描述的示例实施例的限制。
参见图1，其示出了本公开实施例提供的一种用于空间目标的捕捉系统1，所述捕捉系统1具体包括：多个捕捉装置11、传送装置12、发射装置13、减速及回收装置14；其中，
所述多个捕捉装置11，用于依次被发射至目标轨道以捕捉失效的空间目标；
所述传送装置12，用于沿设定的传送轨迹将所述多个捕捉装置11依次进给移动至所述发射装置13中的第一设定位置；
所述发射装置13，用于将处于所述第一设定位置的捕捉装置11发射至所述目标轨道以捕捉所述空间目标；
所述减速及回收装置14，用于当所述捕捉装置11被发射并飞行设定距离后对所述捕捉装置11进行减速控制。
需要说明的是，本公开实施例中，对于失效的空间目标并不局限于失效的飞行器，当然也包括空间碎片。
此外，传送装置12的传送轨迹也并不限定于图1和图2中所示的盘绕状，其他形式及形状的传送轨迹也能够适用于本公开中的捕捉系统1。
对于图1所示的捕捉系统1，通过传送装置12将捕捉装置11沿设定的传送轨迹依次进给移动至发射装置13中的第一设定位置，从而使得捕捉装置11被发射至目标轨道以捕捉失效的空间目标；同时，当捕捉装置11被发射并飞行设定距离后，减速及回收装置14能够对捕捉装置11进行减速控制以防止捕捉装置11反弹进而损伤搭载的航天器。本公开实施例提供的捕捉系统1能够捕捉各种尺寸及类型的空间目标，当然也包括传统机械臂和飞网方式难以捕捉的各种尺寸及类型的空间目标，整体适应性强，且捕捉效率高。
需要说明的是，本公开实施例中的捕捉装置11中各部件的材质均采用可降解的材料，因此捕捉装置11及失效的空间目标被拖曳至衰减轨道后均能够进行充分降解。
对于图1所示的捕捉系统1，可以理解的是，捕捉系统1可以安装于其他航天器（图中未示出）表面，例如卫星，以通过搭载卫星的方式飞行至设定的空间轨道上，从而进行捕捉任务。
可以理解地，当捕捉装置11命中失效的空间目标后，被命中的失效空间目标能够随着捕捉系统搭载的航天器一起飞行至设定的衰减轨道，从而实现了将目标轨道的失效的空间目标拖曳至设定的衰减轨道的任务。
此外，需要说明的是，在具体实施过程中图1所示的捕捉系统1中能够安装200枚捕捉装置11，以执行200次的捕捉任务，从而避免了地面站多次发射捕捉系统1，因此能够节约发射成本。同时，本公开实施例提供的捕捉系统1能够搭载300Kg质量级的卫星在轨实施捕捉任务，成本低且效率高。当然，可以理解的是，在具体实施过程中也可以根据实际情况改变捕捉装置11的数量，并不局限于前述技术方案所述的200枚捕捉装置11。
又一方面，具体来说，当图1所示的捕捉系统1组装完成后，如图2和图3所示，捕捉系统1的最大径向直径不超过1000毫米，最大高度不超过210毫米。因此，本公开实施例中的捕捉系统1尺寸小，质量轻，实现了捕捉系统1的轻量化设计。
对于图1所示的捕捉系统1，在一些可能的实施方式中，所述减速及回收装置14还用于当所述捕捉装置11脱靶时，回收所述捕捉装置11至所述发射装置13中。可以理解地，当捕捉装置11没有命中空间目标时，捕捉装置11被回收至发射装置13后能够重新被发射至其他目标轨道以继续寻找其他失效的空间目标进行捕捉和拖拽。
对于图1所示的捕捉系统1，在一些可能的实施方式中，如图4所示，每个所述捕捉装置11包括：头部111、至少一个倒刺单元112、定位单元113、缠绕于所述捕捉装置11内部的第一绳索114、减速轴115以及缠绕于所述减速轴115上的第二绳索116；其中，
所述头部111呈尖状，用于击穿并插入所述空间目标内部；
所述倒刺单元112，用于防止所述空间目标在拖曳过程中发生脱落；
所述定位单元113，用于实时定位所述捕捉装置11的进给移动位置，以保证所述捕捉装置11能够进给移动至所述第一设定位置；
所述第二绳索116与所述第一绳索114连接，用于当所述第一绳索114完全展开后，对所述捕捉装置11进行减速控制直至所述捕捉装置11减速至零。
需要说明的是，在具体实施过程中，捕捉装置11中的各部件，举例来说，头部111、倒刺单元112以及定位单元113均可脱离捕捉系统1以发射至目标轨道。此外，随着捕捉装置11的发射，第一绳索114能够逐渐脱离捕捉装置11并逐渐展开。当第一绳索114在完全展开之前捕捉装置11已命中失效的空间目标时，捕捉装置11继续带动第一绳索114向前飞行直至第一绳索114完全展开；当第一绳索114完全展开后，第一绳索114会拉动第二绳索116逐渐展开以对捕捉装置11进行减速控制。当然，当第一绳索114完全展开且捕捉装置11没有捕捉到空间目标时，第一绳索114也能够拉动第二绳索116逐渐展开以对捕捉装置11进行减速控制。可以理解的是，在捕捉装置11中第一绳索114和第二绳索116可以为同一条绳索，只是呈分段设计，且第一绳索114缠绕于捕捉装置11自身的内部，第二绳索116缠绕于减速轴115上，这样的设计既能有助于捕捉装置11命中失效的空间目标，又能够有效地对捕捉装置11进行减速控制，以防止捕捉装置11反弹进而损伤搭载的卫星。
需要说明的是，在具体实施过程中，捕捉装置11的质量能够设计为0.5Kg，以实现捕捉系统1的轻量化设计；且第一绳索11的长度能够设计为50米，以在捕捉系统1的50米范围内捕捉失效的空间目标。另一方面，可以理解的是，在前述技术方案中所述的捕捉装置11被发射后能够飞行设定的距离则由第一绳索11的长度决定。
对于上述的实施方式，在一些示例中，如图5所示，所述减速轴115的内部设置有加热装置1151，所述加热装置1151用于当所述空间目标被拖曳至设定的衰减轨道后进行加热以熔断所述第二绳索116。可以理解地，加热装置1151具体可以为功率电阻。
对于图1所示的空间目标捕捉系统1，在一些可能的实施方式中，如图6所示，所述传送装置12包括：第一框架121、多个安装支座122、多个减速轴孔123、第一传动齿轮124、与所述第一传动齿轮124连接的第一驱动机构125、多个供电触点126以及多个开关触点127；其中，
所述多个安装支座122均匀地设置于所述第一框架121的上侧；其中，每个所述捕捉装置11对应地安装于每个所述安装支座122上；
所述多个减速轴孔123均匀且对称地设置于所述第一框架121的两侧；其中，所述减速轴115以可旋转的方式对应地安装于所述减速轴孔123内；
所述第一框架121的底部设置有T型导轨1211，如图7所示，所述T型导轨1211的两侧面分别设置有与所述第一传动齿轮124啮合的第一传动齿条128；
所述第一驱动机构125，用于驱动所述第一传动齿轮124与所述第一传动齿条128啮合传动以使得多个所述捕捉装置11依次沿设定的传送轨迹进给移动；
所述多个供电触点126均匀地设置于所述第一框架121的两侧，用于在所述捕捉装置11的发射及减速回收过程中提供电能；
所述多个开关触点127均匀地设置于所述第一框架121的一侧。
需要说明的是，在具体实施过程中，第一驱动机构125驱动第一传动齿轮124与第一传动齿条128啮合传动，从而使得多个捕捉装置11依次进给移动直至进入发射装置13中的第一设定位置。可以理解地，在捕捉装置11的进给移动过程中，第一驱动机构125与第一传动齿轮124不发生移动，第一框架121及设置于第一框架121上的其他部件，例如安装支座122、减速轴孔123、供电触点126和开关触点127会与捕捉装置11同时进给移动。此外，需要说明的是，传送装置12通过T型导轨1211和搭载的卫星进行配合组装。
可以理解地，在传送装置12中包括多个传送单元，每个传送单元则对应地包括一个安装支座122、两个减速轴孔123以安装减速轴115、两个供电触点126以及一个开关触点127；每个捕捉装置11均对应于一个传送单元，在捕捉装置11的传送过程中，每个传送单元随着对应地捕捉装置11同时进给移动，并配合完成捕捉装置11的发射、减速及回收任务。
另一方面，可以理解的是，第一驱动机构125具体可以为步进电机。
又一方面，需要说明的是，捕捉装置11的传送轨迹取决于第一框架121的形状，例如图1中的盘绕状。
对于图1所示的捕捉系统1，在一些可能的实施方式中，如图8和图9所示，所述发射装置13包括：第二框架131、发射腔膛132、发射支座133、第二驱动机构134、第二传动机构135、第一弹性机构136、导轨137、第一行程开关138、第二行程开关139以及导电片140；其中，
所述发射腔膛132呈半圆柱状，设置于所述第二框架131内；
所述发射支座133穿过所述发射腔膛132且与所述发射腔膛132固定连接；
所述第二驱动机构134，用于驱动所述第二传动机构135进行传动；其中，当所述捕捉装置11未进给至第一设定位置时，所述第二驱动机构134驱动所述第二传动机构135进行传动以使得所述发射支座133与所述发射腔膛132沿所述导轨137下移至第二设定位置且压缩所述第一弹性机构136；以及，当所述捕捉装置11已进给至所述第一设定位置时，所述第二驱动机构134驱动所述第二传动机构135继续传动以使得所述发射支座135与所述发射腔膛132在所述第一弹性机构135的作用下沿所述导轨137上移并推动所述捕捉装置11被发射；
所述第一行程开关138，用于当所述发射支座133与所述发射腔膛132下移至第二设定位置时，发送所述第二传动机构135停止传动的指令，以使得所述发射支座133保持于所述第二设定位置处；
所述第二行程开关139，用于当所述捕捉装置11已进给移动至所述第一设定位置时与所述开关触点127电连接以发送所述捕捉装置11停止进给移动的指令，以使得所述捕捉装置11保持于所述第一设定位置处；
所述导电片140设置于所述第二框架131的两侧，与所述供电触点126电连接以提供电能。
此外，在具体实施过程中，第二驱动机构134具体可以为步进电机，与搭载的卫星上的供电装置（图中未示出）连接。
同时，在具体实施过程中，第一弹性机构136具体可以为弹簧，其穿过发射支座133，在发射支座133与发射腔膛132下移的过程中第一弹性机构136能够被压缩；此外，在捕捉装置11进给移动至第一设定位置后，第一弹性机构136的底部能够与安装支座122的顶部抵接。
对于上述实施方式，在一些示例中，所述第二传动机构135包括第二传动齿轮1351，以及与第二传动齿轮1351啮合的第二传动齿条1352；且所述第二传动齿轮1351为半齿结构。可以理解地，在本公开的具体实施过程中，如图8所示，第二传动齿轮1351与第二驱动机构134连接；第二传动齿条1352固定地设置于发射支座133的一侧。因此，当捕捉装置11达到第一设定位置且瞄准待捕捉的空间目标后，第二驱动机构134驱动第二传动机构135传动，由于第二传动齿轮1351为半齿结构，因此在传动的过程中第二传动齿条1352会被释放，进而发射支座133与发射腔膛132在第一弹性机构136的弹性力作用下上移并将捕捉装置11推射出发射装置13，完成发射动作。需要说明的是，第一弹性机构136的弹簧力为2000N，可使得0.5kg的捕捉装置11获得20m/s的飞行速度，进而能够使得捕捉装置11高速击穿并插入空间目标体内，比如失效飞行器的蜂窝夹层结构。
对于上述实施方式，在一些示例中，如图10所示，所述第二传动机构135还包括棘轮1353与棘爪1354；其中，所述棘轮1353设置于所述第二传动齿轮1351的内部，所述棘爪1354与第二驱动机构134连接。具体来说，在捕捉装置11未进给至第一设定位置时，卫星上的供电装置为第二驱动机构134提供电能，进而第二驱动机构134驱动第二传动齿轮1351与第二传动齿条1352啮合传动，从而将发射支座133下移至第二设定位置，并通过下移的发射支座133压缩第一弹性机构136，同时触发第一行程开关138；当第一行程开关138被触发后，第二驱动机构134停止运转，此时棘爪1354抵住棘轮1353的外周壁以使得第二传动齿轮1351停止传动并保持于当前传动位置，进而使得发射支座133与发射腔膛132保持于第二设定位置固定不动。
可以理解的是，当捕捉装置11进给移动至第一设定位置且准备发射时，捕捉装置11中的定位单元113能够与发射腔膛132相抵接，因而在第一弹性机构136的弹性力作用下，捕捉装置11被发射至目标轨道。
又一方面，需要说明的是，在进给移动至发射装置13中的当前捕捉装置11被发射后，传送装置12会继续沿设定的传送轨迹进给移动，一方面使得被发射的捕捉装置11的对应安装部件如安装支座122及其他部件如减速轴115被移动出发射装置13；另一方面，下一个捕捉装置11及其对应的安装部件会被进给移动至发射装置13中以进行下一次的捕捉任务。
对于图1所示的捕捉系统1，在一些可能的实施方式中，如图11和12所示，所述减速及回收装置14包括：电磁阀141、驱动轴142、缠绕于所述驱动轴142上的第二弹性机构143、磁阻尼器144以及第三驱动机构145；其中，
所述电磁阀141，用于当所述捕捉装置11进给移动至所述第一设定位置时，通过停止工作以使得所述第二弹性机构143推动所述驱动轴142与所述减速轴115连接；以及，当所述捕捉装置11被发射后，通过吸合所述驱动轴142以使得所述驱动轴142与所述减速轴115分离；其中，所述驱动轴142通过齿轮联轴器146与所述减速轴115连接；
所述磁阻尼器144，用于在所述第一绳索114拉动所述第二绳索116逐渐展开时产生阻尼力以对所述捕捉装置11进行减速控制；
所述第三驱动机构145与所述磁阻尼器144连接，用于驱动所述减速轴115反向转动以通过拉动第二减速116回收所述捕捉装置11。
需要说明的是，当捕捉装置11进给移动至第一设定位置并触发第二行程开关139时，此时捕捉装置11停止进给移动，电磁阀141同样也停止工作，以通过第二弹性机构143推动驱动轴142与减速轴115连接，此时通过齿轮联轴器146在驱动轴142与减速轴115之间进行阻尼力矩的传输；可以理解的是，当捕捉装置11被发射后，第一绳索114逐渐并且完全脱离捕捉装置11后，第一绳索114继续拉动第二绳索116，由于第二绳索116缠绕在减速轴115上，因此在第二绳索116被拉动的过程中，减速轴115发生转动；而另一方面，减速轴115通过齿轮联轴器146与驱动轴142连接，因此当磁阻尼器144产生阻尼力时，其阻尼力矩能够通过驱动轴142传输至减速轴115，因此在具体实施过程中通过测量减速轴115的实时角速度能够获得阻尼力矩的大小，以实现在第二绳索116完全展开之前将捕捉装置11的飞行速度减速为零；同时，当捕捉装置11命中空间目标后，可以通过捕捉系统1搭载的航天器将空间目标拖曳至衰减轨道中。可以理解地，当失效的空间目标拖拽至衰减轨道后，可以通过减速轴115上的加热装置117加热，最终将第二绳索116熔断。
对于上述实施方式，在一些示例中，所述齿轮联轴器146可以包括锥形摩擦齿轮1461和从动斜齿轮1462。
需要说明的是，在具体实施过程中，第三驱动机构145可以为步进电机。并且第三驱动机构与磁阻尼器144的连接方式并不局限于图10中的串联连接，也可以设计为一体化设计。
参见图13，其示出了本公开实施例提供的一种用于空间目标的捕捉方法，所述捕捉方法能够应用于前述技术方案所述的捕捉系统1中，所述捕捉方法具体包括：
S1301、传送装置沿设定的传送轨迹将每个捕捉装置依次进给移动至发射装置中的第一设定位置中；
S1302、所述发射装置将处于所述第一设定位置的所述捕捉装置发射至所述目标轨道以捕捉所述失效的空间目标；
S1303、当所述捕捉装置被发射并飞行设定距离后减速及回收装置对所述捕捉装置进行减速控制。
可以理解地，由于图13所示的捕捉方法能够应用于前述技术方案所述的捕捉系统1中，因此对于捕捉方法的具体细节可参见前述技术方案中对于捕捉系统1中各个组件的详细描述，在此不再赘述。
示例性地，对于图13所示的捕捉方法，在一些可能的实施方式中，所述捕捉方法还包括：
当所述捕捉装置脱靶时，所述捕捉装置能够被回收至所述发射装置中。
在上面详细描述的本公开的示例实施例仅仅是说明性的，而不是限制性的。本领域技术人员应该理解，在不脱离本公开的原理和精神的情况下，可对这些实施例或其特征进行各种修改和组合，这样的修改应落入本公开的范围内。
 工业实用性
本公开实施例提供了一种用于空间目标的捕捉系统及方法；通过传送装置将捕捉装置沿设定的传送轨迹依次进给移动至发射装置中的第一设定位置，从而使得捕捉装置被发射至目标轨道以捕捉失效的空间目标；同时，当捕捉装置被发射并飞行设定距离后，减速及回收装置能够对捕捉装置进行减速控制以防止捕捉装置反弹进而损伤搭载的航天器。本公开实施例提供的捕捉系统能够捕捉各种尺寸及类型的空间目标，当然也包括传统机械臂和飞网方式难以捕捉的各种尺寸及类型的空间目标，整体适应性强，且捕捉效率高。

Claims (10)

  1. 一种用于空间目标的捕捉系统，其特征在于，所述捕捉系统包括：多个捕捉装置、传送装置、发射装置、减速及回收装置；其中，
    所述多个捕捉装置，用于依次被发射至目标轨道以捕捉失效的空间目标；
    所述传送装置，用于沿设定的传送轨迹将所述多个捕捉装置依次进给移动至所述发射装置中的第一设定位置；
    所述发射装置，用于将处于所述第一设定位置的捕捉装置发射至所述目标轨道以捕捉所述空间目标；
    所述减速及回收装置，用于当所述捕捉装置被发射并飞行设定距离后对所述捕捉装置进行减速控制。
  2. 根据权利要求1所述的捕捉系统，其特征在于，所述减速及回收装置还用于当所述捕捉装置脱靶时，回收所述捕捉装置至所述发射装置中。
  3. 根据权利要求1所述的捕捉系统，其特征在于，每个所述捕捉装置包括：头部、至少一个倒刺单元、定位单元、缠绕于所述捕捉装置内部的第一绳索、减速轴以及缠绕于所述减速轴上的第二绳索；其中，
    所述头部呈尖状，用于击穿并插入所述空间目标内部；
    所述倒刺单元，用于防止所述空间目标在拖曳过程中发生脱落；
    所述定位单元，用于实时定位所述捕捉装置的进给移动位置，以保证所述捕捉装置能够进给移动至所述第一设定位置；
    所述第二绳索与所述第一绳索连接，用于当所述第一绳索完全展开后，对所述捕捉装置进行减速控制直至所述捕捉装置减速至零。
  4. 根据权利要求3所述的捕捉系统，其特征在于，所述减速轴的内部设置有加热装置，所述加热装置用于当所述空间目标被拖曳至设定的衰减轨道后进行加热以熔断所述第二绳索。
  5. 根据权利要求1所述的捕捉系统，其特征在于，所述传送装置包括：第一框架、多个安装支座、多个减速轴孔、第一传动齿轮、与所述第一传动齿轮连接的第一驱动机构、多个供电触点以及多个开关触点；其中，
    所述多个安装支座均匀地设置于所述第一框架的上侧；其中，每个所述捕捉装置对应地安装于每个所述安装支座上；
    所述多个减速轴孔均匀且对称地设置于所述第一框架的两侧；其中，所述减速轴以可旋转的方式对应地安装于所述减速轴孔内；
    所述第一框架的底部设置有T型导轨，所述T型导轨的两侧面分别设置有与所述第一传动齿轮啮合的第一传动齿条；
    所述第一驱动机构，用于驱动所述第一传动齿轮与所述第一传动齿条啮合传动以使得多个所述捕捉装置依次沿设定的传送轨迹进给移动；
    所述多个供电触点均匀地设置于所述第一框架的两侧，用于在所述捕捉装置的发射及减速回收过程中提供电能；
    所述多个开关触点均匀地设置于所述第一框架的一侧。
  6. 根据权利要求5所述的捕捉系统，其特征在于，所述发射装置包括：第二框架、发射腔膛、发射支座、第二驱动机构、第二传动机构、第一弹性机构、导轨、第一行程开关、第二行程开关以及导电片；其中，
    所述发射腔膛呈半圆柱状，设置于所述第二框架内；
    所述发射支座穿过所述发射腔膛且与所述发射腔膛固定连接；
    所述第二驱动机构，用于驱动所述第二传动机构进行传动；其中，当所述捕捉装置未进给至第一设定位置时，所述第二驱动机构驱动所述第二传动机构进行传动以使得所述发射支座与所述发射腔膛沿所述导轨下移至第二设定位置且压缩所述第一弹性机构；以及，当所述捕捉装置已进给至所述第一设定位置时，所述第二驱动机构驱动所述第二传动机构继续传动以使得所述发射支座与所述发射腔膛在所述第一弹性机构的作用下沿所述导轨上移并推动所述捕捉装置被发射；
    所述第一行程开关，用于当所述发射支座与所述发射腔膛下移至第二设定位置时，发送所述第二传动机构停止传动的指令，以使得所述发射支座保持于所述第二设定位置处；
    所述第二行程开关，用于当所述捕捉装置已进给移动至所述第一设定位置时与所述开关触点电连接以发送所述捕捉装置停止进给移动的指令，以使得所述捕捉装置保持于所述第一设定位置处；
    所述导电片设置于所述第二框架的两侧，与所述供电触点电连接以提供电能。
  7. 根据权利要求6所述的捕捉系统，其特征在于，所述第二传动机构包括第二传动齿轮，以及与第二传动齿轮啮合的第二传动齿条；且所述第二传动齿轮为半齿结构。
  8. 根据权利要求6所述的捕捉系统，其特征在于，所述第二传动机构还包括棘轮与棘爪；其中，所述棘轮设置于所述第二传动齿轮的内部，所述棘爪与第二驱动机构连接。
  9. 根据权利要求3所述的捕捉系统，其特征在于，所述减速及回收装置包括：电磁阀、驱动轴、缠绕于所述驱动轴上的第二弹性机构、磁阻尼器以及第三驱动机构；其中，
    所述电磁阀，用于当所述捕捉装置进给移动至所述第一设定位置时，通过停止工作以使得所述第二弹性机构推动所述驱动轴与所述减速轴连接；以及，当所述捕捉装置被发射后，通过吸合所述驱动轴以使得所述驱动轴与所述减速轴分离；其中，所述驱动轴通过齿轮联轴器与所述减速轴连接；
    所述磁阻尼器，用于在所述第一绳索拉动所述第二绳索逐渐展开时产生阻尼力以对所述捕捉装置进行减速控制；
    所述第三驱动机构与所述磁阻尼器连接，用于驱动所述减速轴反向转动以通过拉动第二减速回收所述捕捉装置。
  10. 一种用于空间目标的捕捉方法，其特征在于，所述捕捉方法能够应用于权利要求1至9任一项所述的捕捉系统，所述捕捉方法包括：
    传送装置沿设定的传送轨迹将每个捕捉装置依次进给移动至发射装置中的第一设定位置中；
    所述发射装置将处于所述第一设定位置的所述捕捉装置发射至所述目标轨道以捕捉所述失效的空间目标；
    当所述捕捉装置被发射并飞行设定距离后减速及回收装置对所述捕捉装置进行减速控制。
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