CN112193441B - Multi-satellite locking ejection mechanism in space - Google Patents
Multi-satellite locking ejection mechanism in space Download PDFInfo
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- CN112193441B CN112193441B CN202011046787.7A CN202011046787A CN112193441B CN 112193441 B CN112193441 B CN 112193441B CN 202011046787 A CN202011046787 A CN 202011046787A CN 112193441 B CN112193441 B CN 112193441B
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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
- B64G1/645—Separators
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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
- B64G1/641—Interstage or payload connectors
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Abstract
一种太空中多卫星锁紧弹射机构，属于航天技术领域。本发明解决了现有的传统卫星锁紧弹射机构不适用于一箭多星的星链式发射方式的问题。其中所述卫星组件包括N个由上到下依次叠设的卫星机构，每个卫星机构的外部均固装有星间弹射组件，所述星间弹射组件包括若干卫星承力杆及若干星间弹射弹簧，若干卫星承力杆均竖直设置且沿卫星机构周向布置，每上下相临的两个卫星承力杆之间配合连接以实现对相临两个卫星机构之间的周向固定，若干星间弹射弹簧对应设置在每上下相临的两个卫星承力杆之间。到达空间预定轨道后通过指令驱动解锁，在星箭弹射弹簧的作用下，释放卫星组件，卫星机构彼此之间在星间弹射弹簧的作用下相互弹射，释放卫星机构。
A multi-satellite locking and ejection mechanism in space belongs to the field of aerospace technology. The invention solves the problem that the existing traditional satellite locking and ejection mechanism is not suitable for the star-chain launching mode of multiple satellites with one arrow. The satellite assembly includes N satellite mechanisms stacked in sequence from top to bottom, and an inter-satellite ejection assembly is fixed on the outside of each satellite mechanism. Ejection spring, several satellite bearing rods are vertically arranged and arranged along the circumference of the satellite mechanism, and the two satellite bearing rods adjacent to each other up and down are matched and connected to realize the circumferential fixation between the two adjacent satellite mechanisms , and a plurality of inter-satellite ejection springs are correspondingly arranged between two satellite bearing rods adjacent to each other up and down. After reaching the predetermined orbit in space, it is unlocked by command drive. Under the action of the star-arrow ejection spring, the satellite assembly is released, and the satellite mechanisms are ejected from each other under the action of the inter-satellite ejection spring to release the satellite mechanism.
Description
Technical Field
The invention relates to a multi-satellite locking ejection mechanism in space, belonging to the technical field of spaceflight.
Background
In recent years, with the deep development of aerospace technology, a novel cubic satellite has smaller volume and mass, and simultaneously, satellite-borne energy becomes more precious, an explosion bolt and a fusing mechanism are mostly adopted in a traditional satellite locking and releasing mechanism, huge impact force generated in the releasing process of the explosion bolt can generate huge impact damage to a microsatellite, and meanwhile, initiating explosive devices put forward higher requirements on storage environment; the fuse mechanism needs larger power when working, and the energy of the satellite is excessively consumed. Particularly, in recent years, a star-chain type launching mode of one rocket and multi-star is emerging in international space, and the traditional satellite locking ejection mechanism is not suitable any more.
Disclosure of Invention
The invention aims to solve the problem that the conventional satellite locking ejection mechanism is not suitable for a one-arrow-multi-satellite star chain type launching mode, and further provides a multi-satellite locking ejection mechanism in space.
The technical scheme adopted by the invention for solving the technical problems is as follows:
a multi-satellite locking ejection mechanism in space comprises a carrying platform, a locking assembly, a satellite-rocket ejection assembly and a plurality of inter-satellite ejection assemblies,
wherein the satellite component comprises N satellite mechanisms which are sequentially stacked from top to bottom, an inter-satellite ejection component is fixedly arranged outside each satellite mechanism,
the inter-satellite ejection assembly comprises a plurality of satellite bearing rods and a plurality of inter-satellite ejection springs, the satellite bearing rods are vertically arranged and are circumferentially arranged along the satellite mechanisms, every two adjacent satellite bearing rods are connected in a matched mode to achieve circumferential fixing between every two adjacent satellite mechanisms, and the inter-satellite ejection springs are correspondingly arranged between every two adjacent satellite bearing rods;
the satellite and rocket ejection assembly comprises a plurality of satellite and rocket ejection springs and a plurality of satellite and rocket ejection mandrels vertically and fixedly arranged on the carrying platform, the satellite and rocket ejection mandrels are vertically and oppositely arranged with a plurality of satellite bearing rods fixedly arranged on the Nth satellite mechanism, the satellite bearing rods fixedly arranged on the Nth satellite mechanism are matched and connected with the satellite and rocket ejection mandrels under the locking state of the locking assembly to circumferentially fix the Nth satellite mechanism, and the satellite and rocket ejection springs are correspondingly arranged between the satellite bearing rods fixedly arranged on the Nth satellite mechanism and the satellite and rocket ejection mandrels.
Furthermore, the locking assembly comprises a linkage steel wire rope and a plurality of pressing rods, the pressing rods are circumferentially arranged on the outer side of the satellite assembly, the lower end of each pressing rod is rotatably connected to the carrying platform, the top ends of the pressing rods are connected through the linkage steel wire rope, the linkage steel wire rope realizes restraint and unlocking through a steel wire rope unlocking device fixedly arranged on the pressing rods, a limiting block is fixedly arranged on each satellite load bearing rod, a plurality of limiting grooves are formed in each pressing rod from top to bottom, the limiting blocks are arranged in one-to-one correspondence with the limiting grooves, the limiting blocks are correspondingly matched and connected with the limiting grooves in the locking assembly locking state, and the limiting blocks are separated from the limiting grooves in the locking assembly unlocking state.
Furthermore, the limiting block is of a wedge-shaped block structure.
Furthermore, a plurality of unlocking supports are fixedly arranged on the upper surface of the carrying platform, a pressing block is fixedly arranged on the lower portion of each pressing rod, and the pressing blocks and the unlocking supports are vertically arranged in a corresponding mode and fixedly connected through unlocking release devices in the locking state of the locking assembly.
Furthermore, the upper part of the satellite bearing rod is of a columnar structure, the lower part of the satellite bearing rod is of a barrel-shaped structure, and the lower part of each satellite bearing rod is sleeved with the upper part of the satellite bearing rod adjacent to the lower part of the satellite bearing rod.
Furthermore, an inter-satellite ejection mandrel is vertically and fixedly installed at the top end of the satellite bearing rod, an inter-satellite ejection spring is sleeved on the inter-satellite ejection mandrel, and when two adjacent satellite bearing rods are connected in a matched mode, the inter-satellite ejection mandrel is inserted into the lower portion of the satellite bearing rod above the inter-satellite ejection mandrel.
Furthermore, an inter-satellite ejection plug cover is coaxially and fixedly arranged at the lower part of the inter-satellite ejection core shaft, and the lower end of an inter-satellite ejection spring is fixedly arranged on the inter-satellite ejection plug cover.
Furthermore, a satellite and rocket ejection plug cover is coaxially and fixedly arranged at the lower part of the satellite and rocket ejection core shaft, and the lower end of a satellite and rocket ejection spring is fixedly arranged on the satellite and rocket ejection plug cover.
Furthermore, the number of the satellite bearing rods is four, and the satellite bearing rods are symmetrically arranged at four corners of the satellite mechanism in pairs.
Furthermore, a plurality of shear cones are fixedly arranged at the bottom end of each satellite mechanism, a plurality of grooves are formed in the upper surface of the carrying platform and the upper surface of each satellite mechanism, and the shear cones on the satellite mechanisms are correspondingly matched and connected with the grooves below the shear cones in the locking state of the locking assembly.
Compared with the prior art, the invention has the following effects:
the satellite mechanism is installed in a cabin section of a rocket before launching, the satellite assembly is locked before taking off and in the launching process, the satellite assembly is unlocked after reaching a space preset orbit through command driving, the satellite assembly is released under the action of a satellite and rocket launching spring, and then the satellite mechanisms are mutually launched under the action of an inter-satellite launching spring, so that the effect of releasing the satellite mechanisms is achieved.
The device has the advantages of simple structure, low energy consumption, low impact damage, convenience in installation and storage and high reliability.
Drawings
Fig. 1 is a schematic perspective view of the present application in a locked state;
FIG. 2 is an enlarged schematic view at P of FIG. 1;
FIG. 3 is an enlarged schematic view at I of FIG. 1;
FIG. 4 is a schematic view of the fit connection between the hold-down rod and the satellite force-bearing rod;
FIG. 5 is a schematic top view of the present application in a locked condition;
FIG. 6 is a schematic perspective view of the carrying platform;
fig. 7 is a schematic perspective view of the present application in an unlocked state;
FIG. 8 is a schematic perspective structure diagram of a satellite and rocket ejection assembly;
fig. 9 is a schematic perspective view of the inter-satellite ejection assembly.
Detailed Description
The first embodiment is as follows: the embodiment is described with reference to fig. 1 to 9, and a multi-satellite locking ejection mechanism in space comprises a carrying platform 1, a locking assembly, a satellite-rocket ejection assembly and a plurality of inter-satellite ejection assemblies,
wherein the satellite component comprises N satellite mechanisms 2 which are sequentially stacked from top to bottom, an inter-satellite ejection component is fixedly arranged outside each satellite mechanism 2,
the inter-satellite ejection assembly comprises a plurality of satellite bearing rods 3 and a plurality of inter-satellite ejection springs 4, the satellite bearing rods 3 are vertically arranged and are circumferentially arranged along the satellite mechanisms 2, every two adjacent satellite bearing rods 3 are connected in a matched mode to achieve circumferential fixing between every two adjacent satellite mechanisms 2, and the inter-satellite ejection springs 4 are correspondingly arranged between every two adjacent satellite bearing rods 3;
the satellite and rocket ejection assembly comprises a plurality of satellite and rocket ejection springs 5 and a plurality of satellite and rocket ejection mandrels 6 vertically and fixedly arranged on the carrying platform 1, the plurality of satellite and rocket ejection mandrels 6 and a plurality of satellite force-bearing rods 3 fixedly arranged on an Nth satellite mechanism 2 are arranged in an up-and-down opposite mode, under the locking state of the locking assembly, the plurality of satellite force-bearing rods 3 fixedly arranged on the Nth satellite mechanism 2 are matched and connected with the plurality of satellite and rocket ejection mandrels 6 to circumferentially fix the Nth satellite mechanism 2, and the plurality of satellite and rocket ejection springs 5 are correspondingly arranged between the plurality of satellite force-bearing rods 3 fixedly arranged on the Nth satellite mechanism 2 and the plurality of satellite and rocket ejection mandrels 6.
Two adjacent satellite bearing rods 3 from top to bottom are mutually matched and connected to circumferentially fix the two adjacent satellite mechanisms 2, but axial displacement between the two adjacent satellite mechanisms 2 is not limited, namely, the up-and-down displacement between the satellite assembly and each adjacent two satellite mechanisms 2 is limited. The satellite mechanisms 2 are configured and assembled through the matching of the satellite bearing rods 3. The satellite bearing rod 3 solves the problems of positioning and matching when two adjacent satellite mechanisms 2 are assembled, and simultaneously resists most of transverse shearing force borne by the satellite mechanisms 2 through the satellite bearing rod 3, thereby protecting the satellite mechanisms 2.
Under the state that a plurality of satellite mechanisms 2 are overlapped, an inter-satellite ejection spring 4 is compressed between two adjacent satellite bearing rods 3. When the locking assembly is in an unlocked state, the inter-satellite ejection spring 4 provides elastic potential energy for ejection between the satellite mechanisms 2.
The number of the satellite rocket ejection mandrels 6 is the same as the number of the satellite bearing rods 3 in each group of the inter-satellite ejection assemblies.
The satellite and rocket ejection spring 5 provides elastic potential energy for the ejection of the satellite components.
The satellite mechanism 2 is installed in a cabin section of a rocket before launching, the satellite assembly is locked before taking off and in the launching process, the satellite assembly is unlocked after reaching a space preset orbit through command driving, the satellite assembly is released under the action of a satellite and rocket launching spring 5, and then the satellite mechanisms 2 are mutually launched under the action of an inter-satellite launching spring 4, so that the effect of releasing the satellite mechanisms 2 is achieved.
The locking assembly comprises a linkage steel wire rope 7 and a plurality of pressing rods 8, the pressing rods 8 are arranged on the outer side of the satellite assembly along the circumferential direction, the lower end of each pressing rod 8 is rotatably connected to the carrying platform 1, the top ends of the pressing rods 8 are connected through the linkage steel wire rope 7, the linkage steel wire rope 7 is restrained and unlocked through a steel wire rope unlocking device 9 fixedly mounted on the pressing rods 8, a limiting block 10 is fixedly mounted on each satellite force bearing rod 3, a plurality of limiting grooves 11 are formed in each pressing rod 8 from top to bottom, the limiting blocks 10 are arranged in one-to-one correspondence with the limiting grooves 11, the limiting blocks 10 are correspondingly matched and connected with the limiting grooves 11 in a locking state, and the limiting blocks 10 are separated from the limiting grooves 11 in an unlocking state. Through the cooperation between stopper 10 and the spacing groove 11 is connected, realize the restriction to the up-and-down direction displacement between satellite assembly and every two adjacent satellite mechanism 2. The top end of the pressing rod 8 is provided with a through hole, the linkage steel wire ropes 7 can be one and penetrate through the through hole, the linkage steel wire ropes are locking and releasing devices, the steel wire ropes are locked under the action of the locking and releasing devices, and the steel wire ropes are released from constraint after being released. Two ends of the linkage steel wire rope 7 are fixedly connected to the steel wire rope unlocking device 9 respectively, and the top end of the pressing rod 8 is restrained, namely the pressing rod 8 is restrained along the transverse swing. When the steel wire rope unlocking device 9 is unlocked, the steel wire rope with limited length can be released, so that the pressing rod 8 can only swing for a limited angle along the transverse direction, and the swing angle of the pressing rod 8 is controlled. The hold-down bar 8 is guaranteed to rotate only at a small angle around the rotation connection point of its bottom with the carrying platform 1. Through the mode of processing a plurality of spacing grooves 11 on one compression rod 8 for one compression rod 8 is simultaneously with the cooperation of a plurality of satellite mechanism 2 and is implemented and compress tightly the locking, simplifies the device structure.
The limiting block 10 is a wedge-shaped block structure.
A plurality of unlocking supports 12 are fixedly arranged on the upper surface of the carrying platform 1, a pressing block 13 is fixedly arranged on the lower portion of each pressing rod 8, and the pressing blocks 13 and the unlocking supports 12 are vertically arranged correspondingly and fixedly connected through unlocking release devices 14 in a locking state of the locking assembly. A set of hinged support 15 is further arranged between the two support monomers of each set of unlocking support 12, and the bottoms of the plurality of pressing rods 8 are correspondingly rotatably connected to the carrying platform 1 through the plurality of sets of hinged supports 15. The lower part of each compression bar 8 can be provided with a compression block 13 on one side only, and can also be provided with compression blocks 13 on both sides, correspondingly, the number and the positions of the unlocking supports 12 on the carrying platform 1 are arranged in one-to-one correspondence with the compression blocks 13. Locking and unlocking are realized through the matched connection of the unlocking support 12 and the pressing block 13, and the installation position is provided for the pressing rod 8 and the unlocking release device through the unlocking support 12 and the hinged support 15, so that the space on a satellite is saved.
The unlocking and releasing device can realize automatic release through a received instruction, also can be of a compression bolt structure, and can be fixed by compressing the common grail rod.
During unlocking, after the unlocking device of the compression rod 8 releases the axial constraint on the compression rod 8, under the driving action of the satellite ejection assembly, the satellite force bearing rod 3 generates a transverse driving force on the compression rod 8 through the inclined plane matched with the compression rod 8, so that the compression rod 8 rotates for a certain angle along the hinged support 15 on the carrying platform 1, the compression rod 8 is enabled to release the axial constraint on the satellite assembly, and under the driving action of the satellite ejection assembly, the ejection of the satellite assembly and the carrying platform 1 is realized.
The upper part of the satellite bearing rod 3 is a columnar structure, the lower part of the satellite bearing rod 3 is a barrel-shaped structure, and the lower part of each satellite bearing rod 3 is sleeved with the upper part of the satellite bearing rod 3 adjacent to the lower part of the satellite bearing rod. The upper part of the satellite bearing rod 3 is preferably a stepped cylindrical structure, and the inner wall of the lower barrel-shaped structure is in a stepped shape matched with the upper part.
An inter-satellite ejection mandrel 16 is vertically and fixedly installed at the top end of each satellite bearing rod 3, an inter-satellite ejection spring 4 is sleeved on the inter-satellite ejection mandrel 16, and when two adjacent satellite bearing rods 3 are connected in a matched mode, the inter-satellite ejection mandrel 16 is inserted into the lower portion of the satellite bearing rod 3 above the inter-satellite ejection mandrel. The design makes the ejection performance more reliable.
The lower part of the inter-satellite ejection mandrel 16 is coaxially and fixedly provided with an inter-satellite ejection plug cover 17, and the lower end of the inter-satellite ejection spring 4 is fixedly arranged on the inter-satellite ejection plug cover 17. The inter-satellite ejection spring 4 is axially limited by the inter-satellite ejection blocking cover 17.
The lower part of the satellite and rocket ejection mandrel 6 is coaxially and fixedly provided with a satellite and rocket ejection plug cover 18, and the lower end of the satellite and rocket ejection spring 5 is fixedly arranged on the satellite and rocket ejection plug cover 18. The satellite and rocket ejection spring 5 is axially limited by the satellite and rocket ejection blocking cover 18.
The number of the satellite bearing rods 3 is four, and the satellite bearing rods are symmetrically arranged at four corners of the satellite mechanism 2 in pairs.
The bottom end of each satellite mechanism 2 is fixedly provided with a plurality of shear cones 19, the upper surface of the carrying platform 1 and the upper surface of each satellite mechanism 2 are respectively provided with a plurality of grooves 20, and the shear cones 19 on each satellite mechanism 2 are correspondingly connected with the grooves 20 below the same in a matching manner in the locking state of the locking assembly. The inner surface of the groove 20 is a conical bore surface. The shear cone 19 is arranged to play a role in installation positioning and bearing transverse shear force. The satellite mechanism 2 is protected. By adopting the double shear-resistant structure of the shear-resistant cone and the satellite bearing rod 3, 270 percent of transverse shear force of the satellite mechanism is overcome by the satellite bearing rod 3, and the possibility of damaging the satellite mechanism 2 by the transverse shear force in the launching process is effectively reduced. The satellite mechanism and the satellite mechanism are installed and positioned through the shear cone, and the transverse shearing force is resisted through the matching of the satellite bearing rod structures between the satellite mechanisms, so that the transverse load borne by the satellite mechanism is greatly reduced, and the satellite mechanism is protected.
The working principle is as follows:
the satellite mechanism 2 is assembled together through the inter-satellite ejection assembly, then the satellite assembly and the carrying platform 1 are assembled together through the satellite ejection assembly, the lower portion of the pressing rod 8 is rotatably mounted on the carrying platform 1 through the hinged support 15, after the limiting groove 11 in the pressing rod 8 is matched with the limiting block 10 in the satellite bearing rod 3, the pressing rod 8 is longitudinally locked through the unlocking device of the pressing rod 8, and the satellite assembly is longitudinally pressed. And tensioning the upper linkage steel wire rope 7, limiting the transverse swing of the pressing rod 8, and locking the linkage steel wire rope 7 through a steel wire rope unlocking device 9.
After the space antenna reaches the preset orbit, the ejection mechanism is locked to release the satellite assembly after receiving an ejection command. Firstly, the compressing rod 8 unlocking device acts after receiving an instruction, longitudinal restraint on the compressing rod 8 is released, then the steel wire rope unlocking device 9 acts to release a steel wire rope with a certain length, the length of the linkage steel wire rope 7 is extended, the compressing rod 8 is allowed to swing around a hinged support 15 on the carrying platform 1 at a certain angle, under the ejection action of the satellite and rocket ejection assembly, the compressing rod 8 swings outwards at a certain angle around the hinged support 15 on the carrying platform 1, a limiting groove 11 on the compressing rod 8 is separated from contact with a limiting block 10 on the satellite bearing rod 3, restraint on the satellite assembly is further released, and ejection of the satellite assembly is achieved. And then the ejection between the satellite mechanism 2 and the satellite mechanism 2 is realized under the action of the inter-satellite ejection spring 4. The whole process is rapid in action and reliable in function, the power consumption is low through instruction control, the energy of the satellite is saved, the action impact is small, and the satellite is protected through the bearing rod structure.
In this application, implement the restraint to satellite mechanism through stopper and spacing groove cooperation, adopt compressing tightly pole unlocking ware, linkage wire rope and wire rope unlocking ware to implement locking and release compressing tightly the pole. The adjacent two satellite mechanisms are installed and positioned through the shear cones and the grooves, and the transverse shearing force is resisted through the matching of the satellite bearing rods between the satellite mechanisms, so that the transverse load borne by the satellite mechanisms is greatly reduced, and the satellite mechanisms are protected.
Claims (9)
1. A multi-satellite locking ejection mechanism in space is characterized in that: it comprises a carrying platform (1), a locking component, a satellite and rocket ejection component and a plurality of inter-satellite ejection components,
wherein the satellite component comprises N satellite mechanisms (2) which are sequentially stacked from top to bottom, an inter-satellite ejection component is fixedly arranged outside each satellite mechanism (2),
the inter-satellite ejection assembly comprises a plurality of satellite bearing rods (3) and a plurality of inter-satellite ejection springs (4), the satellite bearing rods (3) are vertically arranged and are circumferentially arranged along the satellite mechanisms (2), every two adjacent satellite bearing rods (3) are connected in a matched mode to achieve circumferential fixing between every two adjacent satellite mechanisms (2), and the inter-satellite ejection springs (4) are correspondingly arranged between every two adjacent satellite bearing rods (3);
the satellite and rocket ejection assembly comprises a plurality of satellite and rocket ejection springs (5) and a plurality of satellite and rocket ejection mandrels (6) vertically and fixedly arranged on a carrying platform (1), the plurality of satellite and rocket ejection mandrels (6) and a plurality of satellite bearing rods (3) fixedly arranged on an Nth satellite mechanism (2) are arranged in an up-and-down opposite mode, under the locking state of the locking assembly, the plurality of satellite bearing rods (3) fixedly arranged on the Nth satellite mechanism (2) are matched and connected with the plurality of satellite and rocket ejection mandrels (6) to realize circumferential fixing of the Nth satellite mechanism (2), and the plurality of satellite and rocket ejection springs (5) are correspondingly arranged between the plurality of satellite bearing rods (3) fixedly arranged on the Nth satellite mechanism (2) and the plurality of satellite and rocket ejection mandrels (6); the locking component comprises a linkage steel wire rope (7) and a plurality of pressing rods (8), the plurality of pressing rods (8) are arranged on the outer side of the satellite component along the circumferential direction, the lower end of each pressing rod (8) is rotatably connected to the carrying platform (1), the top ends of a plurality of pressing rods (8) are connected through a linkage steel wire rope (7), the linkage steel wire rope (7) realizes restraint and unlocking through a steel wire rope unlocking device (9) fixedly arranged on a pressing rod (8), each satellite force bearing rod (3) is fixedly provided with a limiting block (10), each pressing rod (8) is provided with a plurality of limiting grooves (11) from top to bottom, the limiting blocks (10) and the limiting grooves (11) are arranged in a one-to-one correspondence manner, and the locking assembly is locked, the plurality of limiting blocks (10) are correspondingly matched and connected with the plurality of limiting grooves (11), and when the locking assembly is in an unlocking state, the limiting blocks (10) are separated from the limiting grooves (11).
2. The in-space multi-satellite locking ejection mechanism of claim 1, wherein: the limiting block (10) is of a wedge-shaped block structure.
3. The in-space multi-satellite locking ejection mechanism of claim 1, wherein: a plurality of unlocking supports (12) are fixedly arranged on the upper surface of the carrying platform (1), a pressing block (13) is fixedly arranged on the lower portion of each pressing rod (8), and the pressing blocks (13) and the unlocking supports (12) are vertically arranged in a corresponding mode and fixedly connected through unlocking release devices (14) in a locking state of the locking assembly.
4. A multi-satellite locking ejection mechanism in space as claimed in claim 1, 2 or 3, wherein: the upper part of the satellite bearing rod (3) is of a columnar structure, the lower part of the satellite bearing rod is of a barrel-shaped structure, and the lower part of each satellite bearing rod (3) is sleeved on the upper part of the satellite bearing rod (3) adjacent to the lower part of the satellite bearing rod.
5. The in-space multi-satellite locking ejection mechanism of claim 4, wherein: an inter-satellite ejection mandrel (16) is vertically and fixedly installed at the top end of each satellite bearing rod (3), an inter-satellite ejection spring (4) is sleeved on the inter-satellite ejection mandrel (16), and when two adjacent satellite bearing rods (3) are connected in a matched mode, the inter-satellite ejection mandrel (16) is inserted into the lower portion of the satellite bearing rod (3) above the inter-satellite ejection mandrel.
6. The in-space multi-satellite locking ejection mechanism of claim 5, wherein: an inter-satellite ejection plug cover (17) is coaxially and fixedly arranged at the lower part of the inter-satellite ejection mandrel (16), and the lower end of an inter-satellite ejection spring (4) is fixedly arranged on the inter-satellite ejection plug cover (17).
7. The in-space multi-satellite locking ejection mechanism of claim 1, 2, 3, 5 or 6, wherein: the lower part of the satellite and rocket ejection mandrel (6) is coaxially and fixedly provided with a satellite and rocket ejection plug cover (18), and the lower end of the satellite and rocket ejection spring (5) is fixedly arranged on the satellite and rocket ejection plug cover (18).
8. The in-space multi-satellite locking ejection mechanism of claim 1, wherein: the number of the satellite bearing rods (3) is four, and the satellite bearing rods are symmetrically arranged at four corners of the satellite mechanism (2) in pairs.
9. The in-space multi-satellite locking ejection mechanism of claim 1, wherein: the bottom of each satellite mechanism (2) is fixedly provided with a plurality of shear cones (19), the upper surface of the carrying platform (1) and the upper surface of each satellite mechanism (2) are respectively provided with a plurality of grooves (20), and the shear cones (19) on each satellite mechanism (2) are correspondingly matched and connected with the grooves (20) below the shear cones under the locking state of the locking assemblies.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202011046787.7A CN112193441B (en) | 2020-09-29 | 2020-09-29 | Multi-satellite locking ejection mechanism in space |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011046787.7A CN112193441B (en) | 2020-09-29 | 2020-09-29 | Multi-satellite locking ejection mechanism in space |
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| CN112193441A CN112193441A (en) | 2021-01-08 |
| CN112193441B true CN112193441B (en) | 2021-11-30 |
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| CN115196052B (en) * | 2022-07-13 | 2024-12-17 | 航天行云科技有限公司 | Fixing device of stacked satellite and control method thereof |
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| CN115783314B (en) * | 2022-11-21 | 2024-04-26 | 北京中科宇航技术有限公司 | Multi-flat-plate stacked satellite unlocking system |
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