WO2021212355A1 - 卫星构型及其分离方法 - Google Patents

卫星构型及其分离方法 Download PDF

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Publication number
WO2021212355A1
WO2021212355A1 PCT/CN2020/086059 CN2020086059W WO2021212355A1 WO 2021212355 A1 WO2021212355 A1 WO 2021212355A1 CN 2020086059 W CN2020086059 W CN 2020086059W WO 2021212355 A1 WO2021212355 A1 WO 2021212355A1
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WIPO (PCT)
Prior art keywords
separation
satellite
sub
platform
plate
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PCT/CN2020/086059
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English (en)
French (fr)
Inventor
安洋
林宝军
蒋桂忠
陈鸿程
田艳
曹冬冬
解放
刘佳伟
Original Assignee
中国科学院微小卫星创新研究院
上海微小卫星工程中心
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Application filed by 中国科学院微小卫星创新研究院, 上海微小卫星工程中心 filed Critical 中国科学院微小卫星创新研究院
Priority to PCT/CN2020/086059 priority Critical patent/WO2021212355A1/zh
Priority to CN202080002017.4A priority patent/CN111954625B/zh
Publication of WO2021212355A1 publication Critical patent/WO2021212355A1/zh

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    • 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
    • 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

Definitions

  • the invention relates to the technical field of spacecraft, in particular to a satellite configuration and a separation method thereof.
  • the multi-satellite launch method In order to increase the utilization rate of the rocket and reduce the launch cost, the multi-satellite launch method has been adopted more and more, the purpose is to use the rocket's carrying capacity as much as possible.
  • Supporting cylindrical satellites have the advantage of large carrying capacity, and are generally used as high-orbit satellites.
  • the multi-satellite launching of the supporting cylindrical satellites mainly adopts the tandem method, that is, two satellites are stacked on top of each other.
  • the current tandem connection release method requires a relatively heavy mechanism, and there is a certain safety risk.
  • the traditional separation control is the process control and drive of the separation of the double star by the rocket. The control process is complicated, the driving demand is greater, and the safety risk is greater.
  • the purpose of the present invention is to provide a satellite configuration and a separation method thereof, so as to solve the problem that the existing dual-satellite separation method is relatively safe.
  • the present invention provides a satellite configuration including a first sub-satellite platform, a second sub-satellite platform, a first separation device, a second separation device, a fairing and a rocket support device, in:
  • the first sub-satellite platform, the first separation device, the second sub-satellite platform, and the second separation device are sequentially connected and fixed in the fairing;
  • the first sub-satellite platform includes a first central bearing tube
  • the second sub-satellite platform includes a second central bearing tube
  • the first central bearing tube and the second central bearing tube pass through the The first separation device is connected;
  • the second central bearing cylinder and the rocket supporting device are connected by the second separating device;
  • the second sub-satellite platform also includes a separation control module and a separation drive module, the separation control module is used to control the startup and operation of the first separation device; the separation drive module is used to provide the first sub-satellite The platform provides separation driving force to separate the first sub-satellite platform from the second sub-satellite platform, and the separation control module is also used to supply power to the separation drive module.
  • the first central force-bearing cylinder includes a first cylindrical body and an upper separation device
  • the second central force-bearing cylinder includes a second cylindrical body and a lower separation device, wherein :
  • the upper separating device is a first truncated cone structure with a narrow top and a wide bottom.
  • the top end of the first truncated cone structure is connected to the bottom end of the first cylindrical body and is fixed to the first cylindrical body.
  • the lower separating device is a second circular truncated cone structure with a wide top and a narrow bottom, and the bottom end of the second circular truncated cone structure is connected to the top of the second cylindrical body and is fixed to the second cylindrical body.
  • the positions of the lug interlocking structure and the lug locking structure are in one-to-one correspondence and then joined to form a plurality of butting portions.
  • each lug interlocking structure has a conical boss
  • the top surface of each lug locking structure has a clamping groove
  • the cone The bosses and the clamping grooves are in one-to-one correspondence and then combined to form a laterally restraining butt joint.
  • the first separation device includes a plurality of pyrotechnic separation elements, the lug interlocking structure and the lug locking structure both have bolt holes, and the pyrotechnic separation element passes A bolt hole longitudinally tightens each of the lug interlocking structure and the corresponding lug locking structure, the explosion of the pyrotechnic separation element causes the lug interlocking structure and the corresponding lug locking structure The longitudinal tightening is released, and the first central force-bearing cylinder and the second central force-bearing cylinder are separated.
  • the pyrotechnic separation element is a combination of a bolt and a separation nut
  • the longitudinal restraint force of the pyrotechnic separation element is 120,000 N
  • the pyrotechnic separation element receives an instruction
  • the separation nut is opened under the action of the pyrotechnic device, and the bolt is drawn out under the action of the spring to realize unlocking.
  • the diameter of the bolt is 12 mm.
  • the first sub-satellite platform and the second sub-satellite platform both include +X board, -X board, +Z board, -Z board, +Y board, and -Y board, where:
  • the shapes of the +X plate, the -X plate, the +Z plate, the -Z plate, the +Y plate, and the -Y plate are all rectangular plates;
  • the +Z plate and the -Z plate are respectively installed at both ends of the first cylindrical body or the second cylindrical body;
  • the +Y plate and the -Y plate are respectively installed on the +Y side and -Y side of the first central bearing tube or the +Y side and -Y side of the second central bearing tube;
  • the +X plate and the -X plate are respectively installed on the +X side and -X side of the first central bearing tube or the +X side and -X side of the second central bearing tube;
  • the rocket supporting device is installed on the -Z surface of the -Z plate of the second sub-satellite platform.
  • both the first sub-satellite platform and the second sub-satellite platform further include a communication antenna, a solar wing, a battery, a payload, and platform equipment, wherein:
  • the communication antennas are respectively installed on the +X surface of the +X board and the -X surface of the -X board;
  • the solar wing is respectively installed on the +Y surface of the +Y plate and the -Y surface of the -Y plate;
  • the batteries are respectively installed on the -X surface of the +X board and the +X surface of the -X board;
  • the platform equipment and load are installed on the -Y surface of the +Y plate and the +Y surface of the -Y plate.
  • the present invention also provides a method for separating the satellite configuration as described above, and the method for separating the satellite configuration includes:
  • the separation control module communicates with the platform equipment of the first sub-satellite platform, and obtains attitude orbit data, orbit information, and time information of the first sub-satellite platform to form double-satellite separation information;
  • the separation control module sends a double-satellite separation instruction to the first separation device and the separation driving module according to the double-satellite separation information;
  • the first separating device performs ignition after receiving the dual-star separation instruction
  • the separation driving module drives the first separation device to perform an axial separation movement according to the double star separation instruction.
  • the separation method of the satellite configuration includes:
  • the star-and-arrow separation control unit of the rocket support device communicates with the platform equipment of the second sub-satellite platform, and obtains attitude orbit data, orbit information, and time information of the second sub-satellite platform to form star-and-arrow separation information ;
  • the star and arrow separation control unit sends a star and arrow separation instruction to the second separation device and the rocket support device according to the star and arrow separation information;
  • the second separating device ignites after receiving the star and arrow separation instruction
  • the rocket supporting device drives the second separating device to perform an axial separation movement according to the star and arrow separation instruction.
  • the start and operation of the first separation device are controlled by the separation control module on the second sub-satellite platform, and the separation drive module on the second sub-satellite platform is the first sub-satellite
  • the platform provides separation driving force to separate the first sub-satellite platform from the second sub-satellite platform, and supplies power to the separation control module and the first separation device, which realizes the separation of the direct control of the off-satellite and the separation of the drive-up, without the need for a rocket support device
  • Figure 1 is a schematic diagram of the existing external tandem multi-satellite layout
  • Figure 2 is a schematic diagram of the existing self-tandem multi-satellite layout
  • Fig. 3 is a schematic diagram of an existing point-type dual-star connection device
  • FIG. 4 is a schematic diagram of a satellite configuration of an embodiment of the present invention.
  • Figure 5 is a schematic diagram of a satellite configuration according to another embodiment of the present invention.
  • Fig. 6 is a schematic diagram of a satellite configuration according to another embodiment of the present invention.
  • FIG. 7 is a schematic diagram of a satellite configuration according to another embodiment of the present invention.
  • the current dual-satellite launching methods of bearing cylindrical satellites generally have two types: parallel and tandem.
  • Traditional tandem multi-satellite launches are divided into external tandem type and self-tandem type.
  • the traditional external tandem multi-satellite layout includes a larger satellite support 104, a connecting structure 103 connecting the double stars, a connecting structure 106 connecting the stars and arrows, and a fairing 105.
  • the satellite support 104 is used to support the upper
  • the satellite 101 and the lower satellite 102 are installed in the satellite bracket 104.
  • the satellite bracket 104 not only needs to be able to envelop the lower satellite 102, but also has sufficient rigidity to support the upper satellite 101. Therefore, the satellite bracket 104 not only needs a larger volume but also a higher strength and rigidity.
  • the weight of the satellite bracket is generally tens of kilograms. It can range from hundreds of kilograms.
  • the disadvantages of the traditional external tandem multi-satellite launch layout are as follows: the use of satellite brackets will increase more weight. For launch, it is hoped that more limited weight will be used for satellites.
  • the envelope size of the satellite placed in the satellite bracket is limited, which is particularly unfavorable for the deployment of equipment such as antennas outside the satellite.
  • the traditional self-tandem multi-satellite launch method includes two satellites (upper satellite flange 201 and lower satellite flange 202) connected by a strap 203, the connection surface is 204, with a typical 1194 Take the strap as an example, the strap weighs about 15 kg. After the double satellites are separated, the strap stays on the off-satellite. A special device is required to fix the satellite strap to prevent it from colliding with the on-board equipment during the separation process. The special device is about 5 kg, and the separation release system totals 20 kg.
  • the disadvantages of the traditional self-tandem multi-satellite launch layout are as follows: the strapping separation method is adopted, the separation device is heavy, and the strap contains the restraint device about 20 kilograms.
  • the limited weight will be used for satellites. After the satellite is separated, the strap is not completely restrained, and the strap is likely to collide with other equipment components on the satellite, which affects the safety of the satellite. If a restraint device is used, additional weight will be added, and a certain amount of movement space will be required in the process of restraining the strap, which squeezes the layout space of the extra-satellite antenna equipment.
  • the traditional double-star connection device is divided into the strap type and the point type.
  • the strap type adopts an outsourcing form to hug the upper and lower flanges of the two satellites together.
  • the strap is tightened by explosive bolts. After receiving the separation instruction, the explosion The bolt is cut under the action of the pyrotechnic device, the strap is loosened, and the satellite is separated.
  • the upper separation device 303 built-in bolts
  • the lower separation device 304 built-in nuts
  • the connecting surface is 305.
  • the separation nut is opened under the action of the pyrotechnic device, the bolt in the separation device 303 is pulled out under the action of the spring, and the satellite is separated.
  • the core idea of the present invention is to provide a satellite configuration and a separation method thereof, so as to reduce the weight of the double-satellite separation device and improve the safety after the double-satellite separation.
  • the present invention provides a satellite configuration and a separation method thereof.
  • the satellite configuration includes a first sub-satellite platform, a second sub-satellite platform, a first separation device, a second separation device, a fairing, and A rocket supporting device, wherein: the first sub-satellite platform, the first separation device, the second sub-satellite platform, and the second separation device are sequentially connected and fixed in the fairing; the first sub-satellite The platform includes a first central force-bearing cylinder, the second sub-satellite platform includes a second central force-bearing cylinder, and the first central force-bearing cylinder and the second central force-bearing cylinder are connected by the first separation device; The second central bearing cylinder and the rocket support device are connected by the second separation device; the second sub-satellite platform also includes a separation control module and a separation drive module, and the separation control module is used to control the The startup and operation of the first separation device; the separation
  • the satellite configuration includes a first sub-satellite platform 10, a second sub-satellite platform 20, a first separation device 30, a second separation device 40, and a fairing. 50 and a rocket supporting device, wherein: the first sub-satellite platform 10, the first separating device 30, the second sub-satellite platform 20, and the second separating device 40 are sequentially connected and fixed in the fairing 50
  • the first sub-satellite platform 10 includes a first central bearing tube 11
  • the second sub-satellite platform 20 includes a second central bearing tube 21, the first central bearing tube 11 and the second center
  • the bearing cylinder 21 is connected through the first separating device 30; the second central bearing cylinder 21 and the rocket support device (not shown in the figure) are connected through the second separating device 40;
  • the second sub-satellite platform 20 also includes a separation control module and a separation drive module.
  • the separation control module is used to control the startup and operation of the first separation device 30; the separation drive module is used for the first sub-satellite platform. 10 provides a separation driving force to separate the first sub-satellite platform 10 from the second sub-satellite platform 20, and the separation control module is also used to supply power to the separation drive module and the first separation device 30 .
  • the first central bearing cylinder 11 includes a first cylindrical body and an upper separating device
  • the second central bearing cylinder 21 includes a second cylindrical body and a lower The separation device, wherein: the upper separation device is a first truncated cone structure 111 with a narrow upper and a width, and the top end of the first truncated cone structure 111 is connected to the bottom end of the first cylindrical body and is connected to the first cylindrical body.
  • the body is fixed, and the outer side of the bottom of the first truncated cone structure 111 has a plurality of lug interlocking structures 31;
  • the lower separating device is a second truncated cone structure 211 with a wide upper and narrow bottom, and the bottom of the second truncated cone structure 211
  • the end is connected to the top end of the second cylindrical body and is fixed to the second cylindrical body.
  • the outer side of the top of the second truncated cone structure 211 has a plurality of corresponding lug locking structures 32; the lug interlocking structure 31 and the second cylindrical body
  • the positions of the lug locking structures 32 correspond to each other and then join to form a plurality of butting portions.
  • each lug interlocking structure 31 has a conical boss 311, and the top surface of each lug locking structure 32 has a card.
  • Connecting grooves 321, the positions of the conical bosses 311 and the clamping grooves 321 are in one-to-one correspondence and then combined to form a laterally constrained butt joint.
  • the first separating device 30 includes a plurality of pyrotechnic separation elements 33, the lug interlocking structure 31 and the lug locking structure 32 both have bolt holes 331, and the pyrotechnic separation element 33.
  • Each lug interlocking structure 31 and its corresponding lug locking structure 32 are tightened longitudinally through the bolt hole 331.
  • the pyrotechnic separation element 33 explodes to cause the lug interlocking structure 31 and its corresponding
  • the lug locking structure 32 releases the longitudinal tightening and separates the first central force-bearing cylinder 11 and the second central force-bearing cylinder 21.
  • the pyrotechnic separation element is a combination of a bolt and a separation nut, and the longitudinal restraint force of the pyrotechnic separation element is 120,000 N.
  • the separation nut is The pyrotechnic device is opened under the action of the trigger, and the bolt is drawn out under the action of the spring to realize unlocking.
  • the diameter of the bolt is 12 mm.
  • the number of pyrotechnic separation elements is 4-12, such as 4, 8 or 12, which can be calculated according to the requirements of the longitudinal restraint force.
  • the first sub-satellite platform 10 and the second sub-satellite platform 20 both include a +X board, a -X board, a +Z board, a -Z board, and a +Y board.
  • -Y plate wherein: the shapes of the +X plate, the -X plate, the +Z plate, the -Z plate, the +Y plate and the -Y plate are all rectangular plates; The +Z plate and the -Z plate are respectively installed on both ends of the first cylindrical body or the second cylindrical body; the +Y plate and the -Y plate are respectively installed on the first cylindrical body The +Y side and -Y side of a central bearing cylinder 11 or the +Y side and -Y side of the second central bearing cylinder 21; the +X plate and the -X plate are respectively installed on the first The +X side and -X side of a central support tube 11 or the +X side and -X side of the second central support tube 21; the rocket support device is installed on the-of the second
  • both the first sub-satellite platform 10 and the second sub-satellite platform 20 further include a communication antenna, a solar wing, a battery, a payload, and platform equipment, wherein: the communication The antennas are installed on the +X side of the +X board and the -X side of the -X board; the solar wing is installed on the +Y side of the +Y board and the -Y side of the -Y board; the batteries are installed on the + The -X surface of the X board and the +X surface of the -X board; the platform equipment and the load are installed on the -Y surface of the +Y board and the +Y surface of the -Y board.
  • the above-mentioned embodiments describe in detail the different configurations of the satellite configuration.
  • the present invention includes but is not limited to the configurations listed in the above-mentioned embodiments, and any changes are made on the basis of the configurations provided in the above-mentioned embodiments.
  • the content all belong to the protection scope of the present invention. Those skilled in the art can draw inferences based on the content of the above-mentioned embodiments.
  • This embodiment also provides a method for separating the satellite configuration as described above.
  • the method for separating the satellite configuration includes: the separation control module communicates with the platform equipment of the first sub-satellite platform 10, and obtains The attitude orbit data, orbit information, and time information of the first sub-satellite platform 10 form binary satellite separation information; the separation control module sends to the first separation device 30 and the separation drive module according to the binary satellite separation information Double-star separation instruction; the first separation device 30 performs ignition after receiving the double-star separation instruction; the separation driving module drives the first separation device 30 to perform an axial separation movement according to the double-star separation instruction.
  • the separation method of the satellite configuration includes: the satellite-arrow separation control unit of the rocket support device communicates with the platform equipment of the second sub-satellite platform 20 , And obtain the attitude orbit data, orbit information, and time information of the second satellite sub-satellite platform 20 to form star-and-arrow separation information; And the rocket support device sends a star and arrow separation instruction; the second separation device 40 receives the star and arrow separation instruction and then ignites; the rocket support device drives the second separation device according to the star and arrow separation instruction 40 for axial separation movement.
  • the start and operation of the first separation device 30 are controlled by the separation control module on the second sub-satellite platform 20, and the separation drive module on the second sub-satellite platform 20 is the first A sub-satellite platform 10 provides separation driving force to separate the first sub-satellite platform 10 from the second sub-satellite platform 20, and the separation control module supplies power to the first separation device 30, which realizes the separation of the direct control of off-satellite and the separation of driving on-satellite , Without the need for the rocket support device to bypass the lower star to control and drive the upper star, the control is simpler, the driving force can be effectively reduced, and the safety risk is lower.
  • the two satellites of the present invention are interconnected using a serial launch layout.
  • the first satellite sub-satellite platform 10 (hereinafter referred to as the “upper satellite”) is connected to the second sub-satellite platform 20 (hereinafter referred to as the “lower satellite”) through the first separation device 30, and the descending satellite is supported by the rocket via the second separation device 40 Devices (examples include rockets) are connected.
  • a point separation method is adopted between the upper and lower stars. Separate the upper star first, and then separate the lower star. The release of the upper satellite is controlled and powered by the lower satellite.
  • the butting surface of the double star is provided with a conical boss 311 and a clamping groove 321 for inserting and fitting, so as to withstand the shear load between the double stars and avoid the damage of the first separating device 30 under the shear load.
  • the present invention not only avoids the external tandem launching layout of the satellite support with a large weight, but also avoids the heavy strap, avoids bumping into the stand-alone equipment installed outside the satellite when the strap is opened, and increases the safety and reliability.
  • the constraint conditions of the pyrotechnic separation element 33 can be reduced, and the explosive destructive force can be further reduced, so that the upper and lower stars will not be damaged under the separation impact load.
  • the present invention solves two main technical problems: reducing the weight of the dual-satellite separation device, using valuable weight resources for the satellite, saving launch costs, improving the performance of the satellite, and improving the cost-effectiveness of the satellite project. Avoid the movement of the components after the double stars are separated, and improve the safety after the double stars are separated.
  • a single pyrotechnic separation element 33 is about 1 kilogram, and the mechanical properties of six pyrotechnic separation elements 33 can be equivalent to the mechanical properties of the tape.
  • the total weight is only 6 kg, which is significantly lower than the 20 kg of the tape system. , Increasing the weight of the payload and improving the performance of the satellite.
  • the point-type separation device of the present invention has a small volume and a small space, does not require a restraint device, and provides a good installation space for an extra-satellite stand-alone equipment.
  • the bearing cylinder configuration satellite tandem separation release device does not use the traditional belt type, the bearing cylinder configuration satellite tandem separation release device adopts multi-point separation, the release and separation of the double star is controlled by the lower satellite, and the double star docking surface is through a cone
  • the shaped boss and the groove cooperate to restrict the lateral displacement and prevent the shearing damage of the separating device (conical insertion does not affect the separation).

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Abstract

一种卫星构型及其分离方法，包括第一子卫星平台（10）、第二子卫星平台（20）、第一分离装置（30）、第二分离装置（40）、整流罩（50）及火箭支撑装置，第一子卫星平台（10）、第一分离装置（30）、第二子卫星平台（20）及第二分离装置（40）依次连接并固定于整流罩（50）中；第一子卫星平台（10）包括第一中心承力筒（11），第二子卫星平台（20）包括第二中心承力筒（21），第一中心承力筒（11）与第二中心承力筒（21）通过第一分离装置（30）连接；第二中心承力筒（21）与火箭支撑装置通过第二分离装置（40）进行连接；第二子卫星平台（20）包括分离控制模块及分离驱动模块，分离控制模块用于控制第一分离装置（30）的启动与运行；分离驱动模块用于为第一子卫星平台（10）提供分离驱动力，以使第一子卫星平台（10）与第二子卫星平台（20）分离，分离控制模块还用于为分离驱动模块供电。

Description

卫星构型及其分离方法 技术领域
本发明涉及航天器技术领域，特别涉及一种卫星构型及其分离方法。
背景技术
在目前的航天项目中，发射费占有较大的一块份额，一般仅次于卫星费用，目前发射费用较高。为了提高火箭的利用率和降低发射成本，多星发射方式已经被越来越多的采用，目的就是尽可能的用满火箭的运载能力。承力筒式卫星具有大承载的优点，一般作为高轨道卫星的构型，目前承力筒构型卫星多星发射主要采用串列式方式，即两颗卫星上下层叠。目前的串列式连接释放方式需要的机构重量较大，且存在一定的安全性风险。另外，在双星分离过程中，传统的分离控制是由火箭对双星的分离进行过程控制和驱动，控制过程复杂，驱动需求较大，安全风险较大。
发明内容
本发明的目的在于提供一种卫星构型及其分离方法，以解决现有的双星分离方式安全风险较大的问题。
为解决上述技术问题，本发明提供一种卫星构型，所述卫星构型包括第一子卫星平台、第二子卫星平台、第一分离装置、第二分离装置、整流罩及火箭支撑装置，其中：
所述第一子卫星平台、所述第一分离装置、所述第二子卫星平台及第二分离装置依次连接并固定于所述整流罩中；
所述第一子卫星平台包括第一中心承力筒，所述第二子卫星平台包括第二中心承力筒，所述第一中心承力筒与所述第二中心承力筒通过所述第一分离装置连接；
所述第二中心承力筒与所述火箭支撑装置通过所述第二分离装置进行 连接；
所述第二子卫星平台还包括分离控制模块及分离驱动模块，所述分离控制模块用于控制所述第一分离装置的启动与运行；所述分离驱动模块用于为所述第一子卫星平台提供分离驱动力，以使所述第一子卫星平台与第二子卫星平台分离，分离控制模块还用于为分离驱动模块供电。
可选的，在所述的卫星构型中，所述第一中心承力筒包括第一圆柱本体及上分离装置，所述第二中心承力筒包括第二圆柱本体及下分离装置，其中：
所述上分离装置为一上窄下宽的第一圆台结构，且所述第一圆台结构的顶端连接所述第一圆柱本体的底端并与所述第一圆柱本体固定，所述第一圆台结构的底部外侧具有多个凸耳连锁结构；
所述下分离装置为一上宽下窄的第二圆台结构，且所述第二圆台结构的底端连接所述第二圆柱本体的顶端并与所述第二圆柱本体固定，所述第二圆台结构的顶部外侧具有多个相应的凸耳锁定结构；
所述凸耳连锁结构和所述凸耳锁定结构的位置一一对应后接合，形成多个对接部。
可选的，在所述的卫星构型中，每个所述凸耳连锁结构的底面具有一圆锥凸台，每个所述凸耳锁定结构的顶面具有一卡接凹槽，所述圆锥凸台与所述卡接凹槽的位置一一对应后结合，形成横向约束对接部。
可选的，在所述的卫星构型中，所述第一分离装置包括多个火工分离元件，所述凸耳连锁结构与所述凸耳锁定结构均具有螺栓孔，火工分离元件通过螺栓孔纵向箍紧每个所述凸耳连锁结构及与其所对应的所述凸耳锁定结构，所述火工分离元件爆炸使所述凸耳连锁结构及与其所对应的所述凸耳锁定结构解除纵向箍紧，并分离所述第一中心承力筒和所述第二中心承力筒。
可选的，在所述的卫星构型中，所述火工分离元件为螺栓与分离螺母的组合，所述火工分离元件的纵向约束力为120000牛，所述火工分离元件收到指令后，所述分离螺母在火工品作动下张开，所述螺栓在弹簧作动下抽出，实现解锁，所述螺栓的直径为12mm。
可选的，在所述的卫星构型中，所述第一子卫星平台及所述第二子卫星平台均包括+X板、-X板、+Z板、-Z板、+Y板及-Y板，其中：
所述+X板、所述-X板、所述+Z板、所述-Z板、所述+Y板及所述-Y板的形状均为矩形板；
所述+Z板和所述-Z板分别安装在所述第一圆柱本体的两端或所述第二圆柱本体的两端；
所述+Y板和所述-Y板分别安装在所述第一中心承力筒的+Y侧和-Y侧或所述第二中心承力筒的+Y侧和-Y侧；
所述+X板和所述-X板分别安装在所述第一中心承力筒的+X侧和-X侧或所述第二中心承力筒的+X侧和-X侧；
所述火箭支撑装置安装在所述第二子卫星平台的-Z板的-Z面。
可选的，在所述的卫星构型中，所述第一子卫星平台及所述第二子卫星平台均还包括通信天线、太阳翼、蓄电池、载荷及平台设备，其中：
所述通信天线分别安装在+X板的+X面及-X板的-X面；
所述太阳翼分别安装在+Y板的+Y面及-Y板的-Y面；
所述蓄电池分别安装在+X板的-X面及-X板的+X面；
所述平台设备及载荷安装在+Y板的-Y面及-Y板的+Y面。
本发明还提供一种如上所述的卫星构型的分离方法，所述卫星构型的分离方法包括：
所述分离控制模块与所述第一子卫星平台的平台设备进行通信，并获取所述第一子卫星平台的姿轨数据、轨道信息及时间信息，形成双星分离信息；
所述分离控制模块根据所述双星分离信息向所述第一分离装置及所述分离驱动模块发送双星分离指令；
所述第一分离装置收到所述双星分离指令后进行点火；
所述分离驱动模块根据所述双星分离指令驱动所述第一分离装置作轴向分离运动。
可选的，在所述的卫星构型的分离方法中，所述卫星构型的分离方法包括：
所述火箭支撑装置的星箭分离控制单元与所述第二子卫星平台的平台设备进行通信，并获取所述第二子卫星平台的姿轨数据、轨道信息及时间信息，形成星箭分离信息；
所述星箭分离控制单元根据所述星箭分离信息向所述第二分离装置及所述火箭支撑装置发送星箭分离指令；
所述第二分离装置收到所述星箭分离指令后进行点火；
所述火箭支撑装置根据所述星箭分离指令驱动所述第二分离装置作轴向分离运动。
在本发明提供的卫星构型及其分离方法中，通过第二子卫星平台上的分离控制模块控制第一分离装置的启动与运行，第二子卫星平台上的分离驱动模块为第一子卫星平台提供分离驱动力，以使第一子卫星平台与第二子卫星平台分离，且为分离控制模块及第一分离装置供电，实现了下星直接控制和驱动上星的分离，无需火箭支撑装置绕过下星对上星进行控制和驱动，控制更加简单，驱动力可有效降低，安全风险较低。
附图说明
图1是现有的外串列式多星布局示意图；
图2是现有的自串列式多星布局示意图；
图3是现有的点式双星连接装置示意图；
图4是本发明一实施例的卫星构型示意图；
图5是本发明另一实施例的卫星构型示意图；
图6是本发明另一实施例的卫星构型示意图；
图7是本发明另一实施例的卫星构型示意图；
图中所示：10-第一子卫星平台；11-第一中心承力筒；111-第一圆台结构；20-第二子卫星平台；21-第二中心承力筒；211-第二圆台结构；30-第一分离装置；31-凸耳连锁结构；311-圆锥凸台；32-凸耳锁定结构；321-卡接凹槽；33-火工分离元件；331-螺栓孔；40-第二分离装置；50-整流罩。
具体实施方式
以下结合附图和具体实施例对本发明提出的卫星构型及其分离方法作进一步详细说明。根据下面说明和权利要求书，本发明的优点和特征将更清楚。需说明的是，附图均采用非常简化的形式且均使用非精准的比例，仅用以方便、明晰地辅助说明本发明实施例的目的。
目前的承力筒式卫星双星发射方式一般有并列式和串列式两种。传统的串列式多星发射分为外串联式及自串联式。
如图1所示，传统的外串列式多星布局包括一个较大的卫星支架104，连接双星的连接结构103、连接星箭的连接结构106及整流罩105，卫星支架104用于支撑上星101，下星102安装在卫星支架104内。卫星支架104既要能包络下星102，也要有足够的刚度强度支撑上星101，因此卫星支架104不仅需要较大的体积还需要较高的强度刚度，卫星支架重量一般为几十公斤到数百公斤不等。传统的外串列式多星发射布局缺点如下：采用卫星支架，会增加较多的重量，对于发射来说，希望将有限的重量多用于卫星。放在卫星支架中的卫星，其包络尺寸受到限制，尤其不利于布局星外的天线等设备。卫星分离时，先分离上星，再分离筒状卫星支架，最后分离下星，分离筒状卫星支架的过程中很容易与下星发生磕碰，且多次分离降低了可靠性和安全性。
如图2所示，传统的自串联式多星发射方式包括，两颗卫星(上星法兰201与下星法兰202)之间通过包带203连接，连接面为204，以典型的1194包带为例，该包带重量约为15公斤。双星分离后，包带留在下星，需要专用装置固定卫星包带，防止其在分离过程中碰撞星上设备，专用装置约5公斤，分离释放系统合计20公斤。传统的自串列式多星发射布局缺点如下：采用包带式分离方式，分离装置重量大，包带含约束装置约20公斤，对于发射来说，希望将有限的重量多用于卫星。卫星分离后，包带处于未完全约束状态，包带容易与星上其他设备部件发生碰撞，影响卫星的安全性。如果采用约束装置，将额外增加重量，同时对包带约束过程中需要一定的运动空间，挤压了星外天线设备的布局空间。
另外，传统的双星连接装置分为包带式和点式，包带式采用外包形式将两颗卫星的上下法兰抱箍在一起，包带通过爆炸螺栓收紧，收到分离指 令后，爆炸螺栓在火工品作用下切断，包带松开，卫星分离。如图3所示，点式采用上分离装置303(内置螺栓)及下分离装置304(内置螺母)连接的方式将两颗卫星的上法兰301及下法兰302连接在一起，连接面为305，收到分离指令后，分离螺母在火工品作用下张开，分离装置303内的螺栓在弹簧作用下抽出，卫星分离。
本发明的核心思想在于提供一种卫星构型及其分离方法，实现降低双星分离装置的重量和提升双星分离后的安全性。
为实现上述思想，本发明提供了一种卫星构型及其分离方法，所述卫星构型包括第一子卫星平台、第二子卫星平台、第一分离装置、第二分离装置、整流罩及火箭支撑装置，其中：所述第一子卫星平台、所述第一分离装置、所述第二子卫星平台及第二分离装置依次连接并固定于所述整流罩中；所述第一子卫星平台包括第一中心承力筒，所述第二子卫星平台包括第二中心承力筒，所述第一中心承力筒与所述第二中心承力筒通过所述第一分离装置连接；所述第二中心承力筒与所述火箭支撑装置通过所述第二分离装置进行连接；所述第二子卫星平台还包括分离控制模块及分离驱动模块，所述分离控制模块用于控制所述第一分离装置的启动与运行；所述分离驱动模块用于为所述第一子卫星平台提供分离驱动力，以使所述第一子卫星平台与所述第二子卫星平台分离，所述分离控制模块还用于为所述分离驱动模块及所述第一分离装置供电。
<实施例一>
本实施例提供一种卫星构型，如图4所示，所述卫星构型包括第一子卫星平台10、第二子卫星平台20、第一分离装置30、第二分离装置40、整流罩50及火箭支撑装置，其中：所述第一子卫星平台10、所述第一分离装置30、所述第二子卫星平台20及第二分离装置40依次连接并固定于所述整流罩50中；所述第一子卫星平台10包括第一中心承力筒11，所述第二子卫星平台20包括第二中心承力筒21，所述第一中心承力筒11与所述第二中心承力筒21通过所述第一分离装置30连接；所述第二中心承力筒21与所述火箭支撑装置(图中未示出)通过所述第二分离装置40进行连接；所述第二子卫星平台20还包括分离控制模块及分离驱动模块，所述 分离控制模块用于控制所述第一分离装置30的启动与运行；所述分离驱动模块用于为所述第一子卫星平台10提供分离驱动力，以使所述第一子卫星平台10与所述第二子卫星平台20分离，所述分离控制模块还用于为所述分离驱动模块及所述第一分离装置30供电。
如图5所示，在所述的卫星构型中，所述第一中心承力筒11包括第一圆柱本体及上分离装置，所述第二中心承力筒21包括第二圆柱本体及下分离装置，其中：所述上分离装置为一上窄下宽的第一圆台结构111，且所述第一圆台结构111的顶端连接所述第一圆柱本体的底端并与所述第一圆柱本体固定，所述第一圆台结构111的底部外侧具有多个凸耳连锁结构31；所述下分离装置为一上宽下窄的第二圆台结构211，且所述第二圆台结构211的底端连接所述第二圆柱本体的顶端并与所述第二圆柱本体固定，所述第二圆台结构211的顶部外侧具有多个相应的凸耳锁定结构32；所述凸耳连锁结构31和所述凸耳锁定结构32的位置一一对应后接合，形成多个对接部。
如图5～7所示，在所述的卫星构型中，每个所述凸耳连锁结构31的底面具有一圆锥凸台311，每个所述凸耳锁定结构32的顶面具有一卡接凹槽321，所述圆锥凸台311与所述卡接凹槽321的位置一一对应后结合，形成横向约束对接部。在所述的卫星构型中，所述第一分离装置30包括多个火工分离元件33，所述凸耳连锁结构31与所述凸耳锁定结构32均具有螺栓孔331，火工分离元件33通过螺栓孔331纵向箍紧每个所述凸耳连锁结构31及与其所对应的所述凸耳锁定结构32，所述火工分离元件33爆炸使所述凸耳连锁结构31及与其所对应的所述凸耳锁定结构32解除纵向箍紧，并分离所述第一中心承力筒11和所述第二中心承力筒21。
具体的，在所述的卫星构型中，所述火工分离元件为螺栓与分离螺母的组合，火工分离元件的纵向约束力为120000牛，火工分离元件收到指令后，分离螺母在火工品作动下张开，螺栓在弹簧作动下抽出，实现解锁，所述螺栓的直径为12mm。火工分离元件的数量为4～12个，例如4个，8个或12个，可根据纵向约束力的需求进行计算得出。
进一步的，在所述的卫星构型中，所述第一子卫星平台10及所述第二 子卫星平台20均包括+X板、-X板、+Z板、-Z板、+Y板及-Y板，其中：所述+X板、所述-X板、所述+Z板、所述-Z板、所述+Y板及所述-Y板的形状均为矩形板；所述+Z板和所述-Z板分别安装在所述第一圆柱本体的两端或所述第二圆柱本体的两端；所述+Y板和所述-Y板分别安装在所述第一中心承力筒11的+Y侧和-Y侧或所述第二中心承力筒21的+Y侧和-Y侧；所述+X板和所述-X板分别安装在所述第一中心承力筒11的+X侧和-X侧或所述第二中心承力筒21的+X侧和-X侧；所述火箭支撑装置安装在所述第二子卫星平台20的-Z板的-Z面。
进一步的，在所述的卫星构型中，所述第一子卫星平台10及所述第二子卫星平台20均还包括通信天线、太阳翼、蓄电池、载荷及平台设备，其中：所述通信天线分别安装在+X板的+X面及-X板的-X面；所述太阳翼分别安装在+Y板的+Y面及-Y板的-Y面；所述蓄电池分别安装在+X板的-X面及-X板的+X面；所述平台设备及载荷安装在+Y板的-Y面及-Y板的+Y面。
综上，上述实施例对卫星构型的不同构型进行了详细说明，当然，本发明包括但不局限于上述实施中所列举的构型，任何在上述实施例提供的构型基础上进行变换的内容，均属于本发明所保护的范围。本领域技术人员可以根据上述实施例的内容举一反三。
<实施例二>
本实施例还提供一种如上所述的卫星构型的分离方法，所述卫星构型的分离方法包括：所述分离控制模块与所述第一子卫星平台10的平台设备进行通信，并获取所述第一子卫星平台10的姿轨数据、轨道信息及时间信息，形成双星分离信息；所述分离控制模块根据所述双星分离信息向所述第一分离装置30及所述分离驱动模块发送双星分离指令；所述第一分离装置30收到所述双星分离指令后进行点火；所述分离驱动模块根据所述双星分离指令驱动所述第一分离装置30作轴向分离运动。
进一步的，在所述的卫星构型的分离方法中，所述卫星构型的分离方法包括：所述火箭支撑装置的星箭分离控制单元与所述第二子卫星平台20的平台设备进行通信，并获取所述第二子卫星平台20的姿轨数据、轨道信 息及时间信息，形成星箭分离信息；所述星箭分离控制单元根据所述星箭分离信息向所述第二分离装置40及所述火箭支撑装置发送星箭分离指令；所述第二分离装置40收到所述星箭分离指令后进行点火；所述火箭支撑装置根据所述星箭分离指令驱动所述第二分离装置40作轴向分离运动。
在本发明提供的卫星构型及其分离方法中，通过第二子卫星平台20上的分离控制模块控制第一分离装置30的启动与运行，第二子卫星平台20上的分离驱动模块为第一子卫星平台10提供分离驱动力，以使第一子卫星平台10与第二子卫星平台20分离，分离控制模块为第一分离装置30供电，实现了下星直接控制和驱动上星的分离，无需火箭支撑装置绕过下星对上星进行控制和驱动，控制更加简单，驱动力可有效降低，安全风险较低。
本发明的两颗卫星互联，采用串列式发射布局。第一子卫星平台10(以下简称为“上星”)通过第一分离装置30与第二子卫星平台20(以下简称为“下星”)连接，下星通过第二分离装置40与火箭支撑装置(示例包括火箭)连接。上星与下星之间采用点式分离方式。先分离上星，再分离下星。上星的分离释放由下星控制和供电。双星对接面有圆锥凸台311和卡接凹槽321插入配合，以承受双星之间的剪切载荷，避免第一分离装置30在剪切载荷下的损坏。本发明不仅仅避免了增加重量较大的卫星支架的外串列式发射布局，也避免了沉重的包带，避免包带打开时磕碰到星外安装的单机设备，安全可靠性增加。另外，由于第一分离装置30不担心剪切载荷的应力，可以降低火工分离元件33的约束条件，并进一步减小爆炸破坏力，不至于使上下星在分离冲击载荷下受到破坏。
因此，本发明解决的技术问题主要有两个：减小了双星分离装置的重量，将宝贵的重量资源用于卫星，可以节约发射费用，提升卫星的性能，提升卫星项目的性价比。避免双星分离后的部件运动，提高双星分离后的安全性。
具体的，单个火工分离元件33约1公斤，六个火工分离元件33的力学性能可以等效于包带的力学性能，总重量仅为6公斤，相对于包带系统的20公斤大幅降低，增加了有效载荷的重量，提升了卫星的性能。
双星分离后，所有的零部件运动发生在第一分离装置内部，不会对星外设备产生碰撞危险。本发明的点式分离装置体积小，空间小，不需要约束装置，且给星外单机设备提供了良好的安装空间。
承力筒构型卫星串列分离释放装置不采用传统的包带式，承力筒构型卫星串列分离释放装置采用多点式分离，双星的释放分离由下星控制，双星对接面通过圆锥形凸台和凹槽相配合，约束横向位移，防止分离装置剪切损坏(锥形插入，不影响分离)。
本说明书中各个实施例采用递进的方式描述，每个实施例重点说明的都是与其他实施例的不同之处，各个实施例之间相同相似部分互相参见即可。对于实施例公开的系统而言，由于与实施例公开的方法相对应，所以描述的比较简单，相关之处参见方法部分说明即可。
上述描述仅是对本发明较佳实施例的描述，并非对本发明范围的任何限定，本发明领域的普通技术人员根据上述揭示内容做的任何变更、修饰，均属于权利要求书的保护范围。

Claims (9)

  1. 一种卫星构型，其特征在于，所述卫星构型包括第一子卫星平台、第二子卫星平台、第一分离装置、第二分离装置、整流罩及火箭支撑装置，其中：
    所述第一子卫星平台、所述第一分离装置、所述第二子卫星平台及第二分离装置依次连接并固定于所述整流罩中；
    所述第一子卫星平台包括第一中心承力筒，所述第二子卫星平台包括第二中心承力筒，所述第一中心承力筒与所述第二中心承力筒通过所述第一分离装置连接；
    所述第二中心承力筒与所述火箭支撑装置通过所述第二分离装置进行连接；
    所述第二子卫星平台还包括分离控制模块及分离驱动模块，所述分离控制模块用于控制所述第一分离装置的启动与运行；所述分离驱动模块用于为所述第一子卫星平台提供分离驱动力，以使所述第一子卫星平台与第二子卫星平台分离，分离控制模块还用于为分离驱动模块供电。
  2. 如权利要求1所述的卫星构型，其特征在于，所述第一中心承力筒包括第一圆柱本体及上分离装置，所述第二中心承力筒包括第二圆柱本体及下分离装置，其中：
    所述上分离装置为一上窄下宽的第一圆台结构，且所述第一圆台结构的顶端连接所述第一圆柱本体的底端并与所述第一圆柱本体固定，所述第一圆台结构的底部外侧具有多个凸耳连锁结构；
    所述下分离装置为一上宽下窄的第二圆台结构，且所述第一圆台结构的底端连接所述第二圆柱本体的顶端并与所述第二圆柱本体固定，所述第二圆台结构的顶部外侧具有多个相应的凸耳锁定结构；
    所述凸耳连锁结构和所述凸耳锁定结构的位置一一对应后接合，形成多个对接部。
  3. 如权利要求2所述的卫星构型，其特征在于，每个所述凸耳连锁结构的底面具有一圆锥凸台，每个所述凸耳锁定结构的顶面具有一卡接凹槽，所述圆锥凸台与所述卡接凹槽的位置一一对应后结合，形成横向约束对接 部。
  4. 如权利要求2所述的卫星构型，其特征在于，所述第一分离装置包括多个火工分离元件，所述凸耳连锁结构与所述凸耳锁定结构均具有螺栓孔，火工分离元件通过螺栓孔纵向箍紧每个所述凸耳连锁结构及与其所对应的所述凸耳锁定结构，所述火工分离元件爆炸使所述凸耳连锁结构及与其所对应的所述凸耳锁定结构解除纵向箍紧，并分离所述第一中心承力筒和所述第二中心承力筒。
  5. 如权利要求4所述的卫星构型，其特征在于，所述火工分离元件为螺栓与分离螺母的组合，所述火工分离元件的纵向约束力为120000牛，所述火工分离元件收到指令后，所述分离螺母在火工品作动下张开，所述螺栓在弹簧作动下抽出，实现解锁，所述螺栓的直径为12mm。
  6. 如权利要求2所述的卫星构型，其特征在于，所述第一子卫星平台及所述第二子卫星平台均包括+X板、-X板、+Z板、-Z板、+Y板及-Y板，其中：
    所述+X板、所述-X板、所述+Z板、所述-Z板、所述+Y板及所述-Y板的形状均为矩形板；
    所述+Z板和所述-Z板分别安装在所述第一圆柱本体的两端或所述第二圆柱本体的两端；
    所述+Y板和所述-Y板分别安装在所述第一中心承力筒的+Y侧和-Y侧或所述第二中心承力筒的+Y侧和-Y侧；
    所述+X板和所述-X板分别安装在所述第一中心承力筒的+X侧和-X侧或所述第二中心承力筒的+X侧和-X侧；
    所述火箭支撑装置安装在所述第二子卫星平台的-Z板的-Z面。
  7. 如权利要求6所述的卫星构型，其特征在于，所述第一子卫星平台及所述第二子卫星平台均还包括通信天线、太阳翼、蓄电池、载荷及平台设备，其中：
    所述通信天线分别安装在+X板的+X面及-X板的-X面；
    所述太阳翼分别安装在+Y板的+Y面及-Y板的-Y面；
    所述蓄电池分别安装在+X板的-X面及-X板的+X面；
    所述平台设备及载荷安装在+Y板的-Y面及-Y板的+Y面。
  8. 一种如权利要求7所述的卫星构型的分离方法，其特征在于，所述卫星构型的分离方法包括：
    所述分离控制模块与所述第一子卫星平台的平台设备进行通信，并获取所述第一子卫星平台的姿轨数据、轨道信息及时间信息，形成双星分离信息；
    所述分离控制模块根据所述双星分离信息向所述第一分离装置及所述分离驱动模块发送双星分离指令；
    所述第一分离装置收到所述双星分离指令后进行点火；
    所述分离驱动模块根据所述双星分离指令驱动所述第一分离装置作轴向分离运动。
  9. 如权利要求8所述的卫星构型的分离方法，其特征在于，所述卫星构型的分离方法包括：
    所述火箭支撑装置的星箭分离控制单元与所述第二子卫星平台的平台设备进行通信，并获取所述第二子卫星平台的姿轨数据、轨道信息及时间信息，形成星箭分离信息；
    所述星箭分离控制单元根据所述星箭分离信息向所述第二分离装置及所述火箭支撑装置发送星箭分离指令；
    所述第二分离装置收到所述星箭分离指令后进行点火；
    所述火箭支撑装置根据所述星箭分离指令驱动所述第二分离装置作轴向分离运动。
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