CN110963078A - Frame-type satellite structure suitable for veneer assembly - Google Patents
Frame-type satellite structure suitable for veneer assembly Download PDFInfo
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- CN110963078A CN110963078A CN201911121348.5A CN201911121348A CN110963078A CN 110963078 A CN110963078 A CN 110963078A CN 201911121348 A CN201911121348 A CN 201911121348A CN 110963078 A CN110963078 A CN 110963078A
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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
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Abstract
The invention relates to a frame type satellite structure suitable for single plate general assembly, which comprises a structural plate and a frame assembly, wherein the structural plate comprises a + Z plate, a-Z plate, a + Y plate, a-Y plate, a + X plate, an-X plate, an upper partition plate and a lower partition plate, a core formed by four points of a separation joint of the frame assembly is taken as a coordinate origin, the direction of the coordinate origin pointing to the + Z plate is taken as the + Z direction, the flying direction is taken as the + X direction, and the Y direction is determined according to the right-hand rule; the outer surface of the Z plate is provided with a communication and measurement and control antenna, and the inner surface of the Z plate is provided with a load device; -the Z plate penetrates through the cabin to be provided with a sensor and a measurement and control antenna; platform equipment and load equipment are placed on the inner surfaces of the + Y plate and the-Y plate, and solar wings are fixedly connected on the outer surfaces of the + Y plate and the-Y plate; a satellite storage battery is arranged on the inner surface of the + X plate, and a thruster is arranged on the outer surface of the-X plate. The invention changes the assembly process of the satellite by changing the structural form of the satellite, thereby improving the efficiency of multi-satellite parallel assembly, and the configuration meets various requirements of all other satellite structural designs.
Description
Technical Field
The invention relates to a frame type satellite structure suitable for single plate final assembly, and belongs to the general field of spaceflight.
Background
The commercial aerospace is rapidly developed, and foreign well-known enterprises such as iridium companies, oneWeb companies, SpaceX companies and the like successively launch low-orbit communication constellation projects, plan to complete the development and production of dozens or even thousands of satellites within a few years, and complete constellation networking within a short time (2-5 years).
The construction of low-orbit constellations presents two challenges: one is how to make a large number of satellites in a short time and to ensure the consistency of their states, and the other is how to make the best use of the conditions of transportation to accomplish the launch of satellites growing in orders of magnitude with the lowest cost and the highest efficiency.
The existing satellites with the magnitude of one ton and above are mostly launched by one arrow and one satellite, and a small number of satellites launched by two arrows and two satellites need to be carried and manufactured with special fairings or manufactured with special switching mechanisms between two satellites to enable the satellites to be connected in series up and down in the launching process. On the order of hundreds of stars, at least 50 successful transmissions are required to be sent into orbit. Therefore, it is not feasible to transmit satellites in a conventional manner.
On the other hand, existing communication satellite developments are typically designed, manufactured, assembled and tested according to different requirements of users, wherein the manufacturing, assembling and testing periods are as long as tens of months to tens of months. This is determined by the overall development of the satellite and the design of the satellite platform configuration, since medium and large satellite loads usually require a stable, high-precision frame for support. Such a frame is typically the entire primary load-bearing structure of the satellite, rather than a partial or unitary satellite structural panel.
On the basis, the assembly of the satellite instrument equipment and all levels of tests can be conducted in sequence from inside to outside, from the module to the whole satellite, and the duration is long. If it is desired to break this sequential process, each structural panel would need to have the rigidity, strength and stability of a single final assembly, and if this stable mechanical environment is not available through the satellite itself, a powerful and systematic ground support device would be fitted to achieve the same. Such a terrestrial replacement device is time consuming and expensive to develop for the production of a single or few stars for a communications satellite, and is a good replacement for hundreds or even thousands of low earth constellation satellites.
Disclosure of Invention
The technical problem solved by the invention is as follows: in order to overcome the defects of the prior art, the frame type satellite structure suitable for the assembly of the single plates is provided, and relates to a novel frame type satellite structure suitable for one-arrow-multi-satellite launching and single plate assembly, the structure is simple, the assembly openness is good, the layout space utilization rate is high, and the structure is suitable for 1000kg and one-arrow-multi-satellite parallel launching.
The technical scheme of the invention is as follows:
a frame-type satellite structure suitable for the assembly of single plates comprises a structural plate and a frame component, wherein the structural plate comprises a + Z plate, a-Z plate, a + Y plate, a-Y plate, a + X plate, an-X plate, an upper partition plate and a lower partition plate,
determining the Y direction according to the right-hand rule by taking the center of a model formed by four points of the frame component separation joint as the origin of coordinates, pointing the direction of the origin of coordinates to a + Z plate as the + Z direction and the flying direction as the + X direction;
the outer surface of the Z plate is provided with a communication and measurement and control antenna, and the inner surface of the Z plate is provided with a load device; -the Z plate penetrates through the cabin to be provided with a sensor and a measurement and control antenna; platform equipment and load equipment are placed on the inner surfaces of the + Y plate and the-Y plate, and solar wings are fixedly connected on the outer surfaces of the + Y plate and the-Y plate; placing a satellite storage battery on the inner surface of the X plate, and placing a thruster on the outer surface of the X plate;
the + Z plate, -Z plate, + Y plate, -Y plate, + X plate, -X plate form the closed trapezoid box;
the upper partition plate, the lower partition plate and the + X plate are parallel, distributed in the closed trapezoidal box body and used for installing a propelling pipeline and control equipment;
the + Z plate, -Z plate, + Y plate, -Y plate, + X plate, -X plate, upper partition plate and lower partition plate are respectively connected with the frame assembly through interfaces on the frame.
Furthermore, the frame-type satellite structure is a rectangular hexahedron structure with a trapezoidal cross section.
Further, the frame assembly comprises a bar system, a buckle, a separation joint, a lifting point joint, a + Z plate joint and a multiple joint, wherein the separation joint, the lifting point joint, the + Z plate joint and the multiple joint are connected through the bar system to form the frame assembly and are connected with the structural plates through the buckle prefabricated on the bar system.
Further, the structural slab is upper and lower aluminium panels + aluminium honeycomb core sub-sandwich panel, pre-buried heat pipe in aluminium honeycomb core.
Furthermore, the thickness of the aluminum honeycomb core is 20 mm-30 mm.
Furthermore, the thickness of the upper and lower aluminum panels is 0.3 mm-0.5 mm.
Furthermore, the total thickness of the structural plate is 20.6mm to 31 mm.
Compared with the prior art, the invention has the beneficial effects that:
(1) after the structure of the traditional satellite is installed, equipment is sequentially installed on each structural plate, at the moment, at most 2-3 small equipment on each satellite can be assembled together, and when equipment with requirements on an assembly pipeline, cables, large-scale equipment and satellite assembly attitude is encountered, all other equipment needs to be suspended, so that the efficiency is low; the frame structure allows all the single boards to be respectively assembled with equipment and then assembled with the star together;
(2) the invention changes the assembly process of the satellite by changing the structural form of the satellite, thereby improving the efficiency of multi-satellite parallel assembly. At the same time, this configuration meets all the requirements of all other satellite structural designs.
Drawings
FIG. 1 is a schematic diagram of a satellite configuration according to the present invention;
fig. 2 is a schematic diagram of the framework of the invention.
Detailed Description
The invention is further illustrated by the following examples.
A frame-type satellite structure suitable for single-plate assembly is shown in figure 1 and comprises a structural plate and a frame assembly 9, wherein the structural plate comprises a + Z plate 1, a-Z plate 2, a + Y plate 3, a-Y plate 4, a + X plate 5, an-X plate 6, an upper partition plate 7 and a lower partition plate 8,
the center of the model formed by four points of the frame component 9 separating the joint 93 is taken as the origin of coordinates, the direction of the origin of coordinates pointing to the + Z plate 1 is taken as the + Z direction, the flying direction is taken as the + X direction, and the Y direction is determined according to the right-hand rule;
the outer surface of the + Z plate 1 is provided with a communication and measurement and control antenna, and the inner surface is provided with load equipment; the Z plate 2 penetrates through the cabin to be provided with a sensor and a measurement and control antenna; platform equipment and load equipment are placed on the inner surfaces of the + Y plate 3 and the-Y plate 4, and solar wings are fixedly connected on the outer surfaces of the + Y plate 3 and the-Y plate 4; a satellite storage battery is arranged on the inner surface of the X plate 5, and a thruster is arranged on the outer surface of the X plate 6;
a closed trapezoidal box body is formed by the + Z plate 1, the-Z plate 2, the + Y plate 3, the-Y plate 4, the + X plate 5 and the-X plate 6;
the upper partition plate 7, the lower partition plate 8 and the + X plate are parallel 5, distributed in the closed trapezoidal box body and used for installing a propulsion pipeline and control equipment;
the + Z plate 1, -Z plate 2, + Y plate 3, -Y plate 4, + X plate 5, -X plate 6, upper baffle 7 and lower baffle 8 are respectively connected with the frame assembly 9 through interfaces on the frame.
The frame-type satellite structure is a rectangular hexahedron structure with a trapezoidal cross section.
As shown in fig. 2, the frame assembly 9 comprises a bar system 91, a buckle 92, a separation joint 93, a lifting point joint 94, a + Z plate joint 95 and a multiple joint 96, wherein the separation joint 93, the lifting point joint 94, the + Z plate joint 95 and the multiple joint 96 are connected through the bar system 91 to form the frame assembly 9 and are connected with the structural plate through the buckle 92 prefabricated on the bar system 91.
The structural plate is an upper aluminum panel, a lower aluminum panel and an aluminum honeycomb core sandwich plate, the heat pipe is embedded in the aluminum honeycomb core, the thickness of the aluminum honeycomb core is 20 mm-30 mm, the thickness of the upper aluminum panel and the lower aluminum panel is 0.3 mm-0.5 mm, and the total thickness of the structural plate is 20.6 mm-31 mm.
The structure assembling process comprises the following steps:
(1) constructing a frame component: connecting the separation joint 93, the hoisting point joint 94, the + Z plate joint 95 and the composite joint 96 through a rod system 91 to form a frame assembly 9, and connecting the frame assembly with a structural plate through a buckle 92 prefabricated on the rod system 91;
(2) mounting instrumentation on the respective structural plates: the outer surface of the + Z plate 1 is provided with a communication and measurement and control antenna, and the inner surface is provided with load equipment; the Z plate 2 penetrates through the cabin to be provided with a sensor and a measurement and control antenna; platform equipment and load equipment are placed on the inner surfaces of the + Y plate 3 and the-Y plate 4, and solar wings are fixedly connected on the outer surfaces of the + Y plate 3 and the-Y plate 4; a satellite storage battery is arranged on the inner surface of the X plate 5, and a thruster is arranged on the outer surface of the X plate 6;
(3) mounting the structural panel to the frame assembly: the-Z plate, + Y plate, -Y plate, + X plate, -X plate, + Z plate are installed in order.
Under the condition of one-arrow-multi-satellite launching, the satellite is ensured to have 1) enough direct incidence area of the solar wing substrate under various working conditions and external heat flow conditions; 2) enough heat dissipation surface; 3) enough antenna installation space can only select the current body-shaped section configuration;
the frame structure is selected to address assembly efficiency in multi-star simultaneous final assembly. After the structure of the traditional satellite is installed, equipment is sequentially installed on each structural plate, at the moment, at most 2-3 small equipment on each satellite can be assembled together, and when equipment with requirements on an assembly pipeline, cables, large-scale equipment and satellite assembly attitude is encountered, all other equipment needs to be suspended, so that the efficiency is low; the frame structure allows all the single boards to be respectively assembled with equipment and then assembled with the star.
In order to reduce the weight of the structure of the satellite, the satellite adopts a composite plate with an aluminum honeycomb at the center, compared with a common solid plate, the structure is low in rigidity, the structure can deform after equipment is installed, when one structural plate is longer than 3 meters, the two plates after the equipment is installed are difficult to be assembled normally, and for the composite plate, a structural connecting hole is formed by embedding a threaded hole in a plate core in advance in a preset embedding manner, so that the problem that the composite plate cannot be assembled can not be remedied by means of correction or error repair and the like is solved.
The invention changes the assembly process of the satellite by changing the structural form of the satellite, thereby improving the efficiency of multi-satellite parallel assembly. At the same time, this configuration meets all the requirements of all other satellite structural designs.
The above description is only for the best mode of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.
Claims (7)
1. A frame type satellite structure suitable for single plate assembly is characterized by comprising a structural plate and a frame component (9), wherein the structural plate comprises a + Z plate (1), -Z plate (2), -Y plate (3), -Y plate (4), -X plate (5), -X plate (6), upper partition plate (7) and lower partition plate (8),
the center of a model formed by four points of the frame assembly (9) separating joint (93) is taken as a coordinate origin, the direction of the coordinate origin pointing to the + Z plate (1) is taken as the + Z direction, the flying direction is taken as the + X direction, and the Y direction is determined according to the right-hand rule;
the outer surface of the + Z plate (1) is provided with a communication and measurement and control antenna, and the inner surface is provided with load equipment; a sensor and a measurement and control antenna are arranged on the Z plate (2) in a cabin penetrating manner; platform equipment and load equipment are placed on the inner surfaces of the + Y plate (3) and the-Y plate (4), and solar wings are fixedly connected on the outer surfaces of the + Y plate and the-Y plate; a satellite storage battery is arranged on the inner surface of the + X plate (5), and a thruster is arranged on the outer surface of the-X plate (6);
the enclosed trapezoidal box body is formed by + Z plates (1), -Z plates (2), -Y plates (3), -Y plates (4), -X plates (5), -X plates (6);
the upper partition plate (7), the lower partition plate (8) and the + X plate are parallel (5), are distributed in the closed trapezoidal box body and are used for installing a propulsion pipeline and control equipment;
the + Z plate (1), -Z plate (2), + Y plate (3), -Y plate (4), + X plate (5), -X plate (6), upper baffle (7), lower baffle 8 are connected with the frame assembly (9) through the interface on the frame respectively.
2. A framed satellite structure adapted for single board assembly, according to claim 1, wherein the framed satellite structure is a rectangular parallelepiped structure with a trapezoidal cross-section.
3. A framed satellite structure adapted for assembly of veneers according to claim 1, wherein the frame assembly (9) comprises a bar system (91), a clasp (92), a separation joint (93), a lifting point joint (94), + Z plate joint (95) and a multiple joint (96), wherein the separation joint (93), the lifting point joint (94), + Z plate joint (95) and the multiple joint (96) are connected by the bar system (91) to form the frame assembly (9) and are connected to the structural plates by the clasp (92) prefabricated on the bar system (91).
4. A frame satellite structure adapted for use with a veneer assembly as recited in claim 1, wherein the structural panels are upper and lower aluminum panels + sandwich panels of aluminum honeycomb core, and heat pipes are embedded in the aluminum honeycomb core.
5. A framed satellite structure adapted for veneer assembly, according to claim 4, wherein the aluminum honeycomb core has a thickness of 20mm to 30 mm.
6. A framed satellite structure adapted for use with a veneer assembly as claimed in claim 4, wherein the upper and lower aluminum panels have a thickness of 0.3mm to 0.5 mm.
7. A framed satellite structure adapted for assembly of veneers, according to claim 4, wherein the total thickness of the structural panels amounts to 20.6mm to 31 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911121348.5A CN110963078A (en) | 2019-11-15 | 2019-11-15 | Frame-type satellite structure suitable for veneer assembly |
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| CN201911121348.5A CN110963078A (en) | 2019-11-15 | 2019-11-15 | Frame-type satellite structure suitable for veneer assembly |
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| CN110963078A true CN110963078A (en) | 2020-04-07 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113371228A (en) * | 2021-06-15 | 2021-09-10 | 北京空间飞行器总体设计部 | Truss type satellite structure suitable for point type satellite-rocket separation mode |
| CN113524128A (en) * | 2021-06-29 | 2021-10-22 | 中国空间技术研究院 | Low-orbit internet satellite load module precision maintaining tool |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0238840A2 (en) * | 1986-03-22 | 1987-09-30 | ERNO Raumfahrttechnik Gesellschaft mit beschränkter Haftung | Payload transport and operating system for space vehicles |
| JPH11208596A (en) * | 1998-01-23 | 1999-08-03 | Mitsubishi Electric Corp | Satellite structure |
| US6131857A (en) * | 1998-10-30 | 2000-10-17 | Hebert; Barry Francis | Miniature spacecraft |
| KR101008142B1 (en) * | 2008-08-07 | 2011-01-14 | 한국항공우주연구원 | Advanced Spacecraft Integration Dolly |
| CN103863577A (en) * | 2014-02-25 | 2014-06-18 | 上海微小卫星工程中心 | Frame panel type satellite configuration and modular satellite |
| CN103963360A (en) * | 2014-05-06 | 2014-08-06 | 上海卫星工程研究所 | Honeycomb sandwich plate and frame bottom surface integrated structure |
| CN106516163A (en) * | 2016-11-17 | 2017-03-22 | 上海卫星工程研究所 | Large-size and high stable truss structure based on high thermal conductive carbon fiber composite materials |
| CN109850186A (en) * | 2019-02-28 | 2019-06-07 | 中国空间技术研究院 | A kind of flat-rack satellite structure and assembly method for several satellite in a rocket parallel connection transmitting |
-
2019
- 2019-11-15 CN CN201911121348.5A patent/CN110963078A/en active Pending
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0238840A2 (en) * | 1986-03-22 | 1987-09-30 | ERNO Raumfahrttechnik Gesellschaft mit beschränkter Haftung | Payload transport and operating system for space vehicles |
| JPH11208596A (en) * | 1998-01-23 | 1999-08-03 | Mitsubishi Electric Corp | Satellite structure |
| US6131857A (en) * | 1998-10-30 | 2000-10-17 | Hebert; Barry Francis | Miniature spacecraft |
| KR101008142B1 (en) * | 2008-08-07 | 2011-01-14 | 한국항공우주연구원 | Advanced Spacecraft Integration Dolly |
| CN103863577A (en) * | 2014-02-25 | 2014-06-18 | 上海微小卫星工程中心 | Frame panel type satellite configuration and modular satellite |
| CN103963360A (en) * | 2014-05-06 | 2014-08-06 | 上海卫星工程研究所 | Honeycomb sandwich plate and frame bottom surface integrated structure |
| CN106516163A (en) * | 2016-11-17 | 2017-03-22 | 上海卫星工程研究所 | Large-size and high stable truss structure based on high thermal conductive carbon fiber composite materials |
| CN109850186A (en) * | 2019-02-28 | 2019-06-07 | 中国空间技术研究院 | A kind of flat-rack satellite structure and assembly method for several satellite in a rocket parallel connection transmitting |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113371228A (en) * | 2021-06-15 | 2021-09-10 | 北京空间飞行器总体设计部 | Truss type satellite structure suitable for point type satellite-rocket separation mode |
| CN113524128A (en) * | 2021-06-29 | 2021-10-22 | 中国空间技术研究院 | Low-orbit internet satellite load module precision maintaining tool |
| CN113524128B (en) * | 2021-06-29 | 2022-08-12 | 中国空间技术研究院 | Low-orbit internet satellite load module precision maintaining tool |
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Application publication date: 20200407 |