CN112249364A - Structural system of a minisatellite, a minisatellite and a method of assembling a minisatellite - Google Patents
Structural system of a minisatellite, a minisatellite and a method of assembling a minisatellite Download PDFInfo
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- CN112249364A CN112249364A CN202011177171.3A CN202011177171A CN112249364A CN 112249364 A CN112249364 A CN 112249364A CN 202011177171 A CN202011177171 A CN 202011177171A CN 112249364 A CN112249364 A CN 112249364A
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- structural system
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- 238000000034 method Methods 0.000 title claims abstract description 13
- 108091092919 Minisatellite Proteins 0.000 title claims description 21
- 108091092878 Microsatellite Proteins 0.000 claims description 5
- 239000003380 propellant Substances 0.000 claims description 4
- 230000003014 reinforcing effect Effects 0.000 claims description 4
- 238000006243 chemical reaction Methods 0.000 claims description 2
- 238000010586 diagram Methods 0.000 description 7
- 230000005540 biological transmission Effects 0.000 description 5
- 230000008878 coupling Effects 0.000 description 5
- 238000010168 coupling process Methods 0.000 description 5
- 238000005859 coupling reaction Methods 0.000 description 5
- 230000010354 integration Effects 0.000 description 3
- 230000003287 optical effect Effects 0.000 description 3
- 238000005452 bending Methods 0.000 description 2
- 238000009434 installation Methods 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 229920000049 Carbon (fiber) Polymers 0.000 description 1
- 239000004917 carbon fiber Substances 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000003384 imaging method Methods 0.000 description 1
- 210000001503 joint Anatomy 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
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Classifications
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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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- 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
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
- B64G1/283—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using reaction wheels
-
- 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
- B64G1/40—Arrangements or adaptations of propulsion systems
- B64G1/402—Propellant tanks; Feeding propellants
-
- 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
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
- B64G1/443—Photovoltaic cell arrays
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- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- Astronomy & Astrophysics (AREA)
- General Physics & Mathematics (AREA)
- Life Sciences & Earth Sciences (AREA)
- Sustainable Development (AREA)
- Radar, Positioning & Navigation (AREA)
- Automatic Assembly (AREA)
Abstract
The embodiment of the invention discloses a structure system of a small satellite, the small satellite and a method for assembling the small satellite, wherein the structure system comprises: a base plate; a plurality of side plates; the bottom plate, the plurality of side plates and the top plate form a prism-like shape when assembled together, the bottom plate and the top plate correspond to the lower bottom surface and the upper bottom surface of the prism-like body respectively, the plurality of side plates correspond to each side surface of the prism-like body respectively, and at least one of the plurality of side plates is provided with a side plate interface used for installing a subsystem single machine of the small satellite on the inner side.
Description
Technical Field
The present invention relates to the field of minisatellites, and more particularly, to a structural system of a minisatellite, and a method of assembling a minisatellite.
Background
A schematic internal structure of a prior art moonlet is shown in fig. 1, which shows in detail the subsystem standalone of the moonlet and the way the payload is installed in the structural system. As shown in fig. 1, the structural system 110P of the moonlet 100P includes an external structural plate 111P, an intra-satellite standalone mounting plate 112P, and a payload mounting frame 113P coupled to the intra-satellite standalone mounting plate 112P. Wherein at least one subsystem unit 130P is mounted on unit mounting plate 112P and payload 120P is mounted on payload mounting frame 113P. However, in this configuration, the overall weight of the moonlet 100P and the weight fraction of the structural system 110P are increased due to the presence of the stand-alone mounting plate 112P; because the payload mounting frame 113P is coupled to the intra-satellite single-machine mounting plate 112P, the force transmission path of the payload is long, mechanical coupling occurs, and the vibration response is large, the satellite platform formed by the single-machine mounting plates 112P also needs to have increased strength to meet the strength requirements of the platform and the load, and the increased strength will increase the mass.
In the Jilin satellite number one launched 10 months in 2015, the structural system of the satellite includes structural outer plates, trusses mounted to the structural outer plates, and central messenger tubes mounted to the trusses. The weight ratio of the structure system is about 21%, and the truss and the central bearing cylinder need to be installed in a time-consuming mode in the assembling process.
The above prior art moonlet architecture system has difficulty meeting the requirements for moonlet payload weight fraction, whole satellite weight, and total assembly integration time in today's commercial satellite field.
Disclosure of Invention
To solve the above technical problems, embodiments of the present invention are directed to a structural system of a small satellite, a small satellite and a method for assembling a small satellite, wherein the structural system of a small satellite has a low weight ratio, the payload of a small satellite has a high weight ratio, a force transmission path is short, mechanical coupling is not generated, and the total assembly and integration time of a small satellite is short.
The technical scheme of the invention is realized as follows:
in a first aspect, an embodiment of the present invention provides a structural system of a small satellite, where the structural system includes:
a base plate;
a plurality of side plates;
a top plate is arranged on the top plate,
wherein the bottom plate, the side plates and the top plate form a prism-like shape when assembled together, the bottom plate and the top plate respectively correspond to the lower bottom surface and the upper bottom surface of the prism-like shape, the side plates respectively correspond to each side surface of the prism-like shape,
wherein at least one of the plurality of side plates has a side plate interface inside for mounting a subsystem stand-alone of the minisatellite.
In a second aspect, an embodiment of the present invention provides a microsatellite, including:
the structural system according to the first aspect;
at least one subsystem unit, the at least one subsystem unit passes through the curb plate interface install to the curb plate.
In a third aspect, an embodiment of the present invention provides a method for assembling the microsatellite according to the second aspect, the method including:
unfolding the side plates by means of ground equipment until the side plates are in the same horizontal plane with the bottom plate;
mounting the at least one subsystem stand-alone to the side panel in a parallel manner;
and folding the unfolded side plates until the side plates are positioned in a position where the small satellite is assembled.
The embodiment of the invention provides a structure system of a small satellite, the small satellite and a method for assembling the small satellite; subsystem unit and the payload of little satellite direct mount respectively on curb plate and bottom plate, have reduced structural system's weight to a great extent, and the payload direct mount of little satellite has the shortest in power transmission path on the bottom plate, has reduced the mechanics coupling to installation payload and subsystem unit to parallel mode can realize the fast assembly of little satellite.
Drawings
FIG. 1 is a schematic diagram of the internal structure of a prior art moonlet;
FIG. 2 is a schematic diagram of a moonlet according to an embodiment of the present invention;
FIG. 3 is an exploded view of the minisatellite of FIG. 2;
FIG. 4 is a schematic view of the minisatellite of FIG. 2 in a sideplate deployed state;
FIG. 5 is an enlarged schematic view of a portion corresponding to the broken line frame of FIG. 4 to emphasize a hinge connecting the bottom plate and the side plate;
FIG. 6 is a schematic diagram of a backplane of the system architecture of the moonlet of FIG. 2;
FIG. 7 is a schematic view of the moonlet of FIG. 2 with a side panel removed;
fig. 8 is a schematic diagram of a first-order X-direction bending mode of a microsatellite according to an embodiment of the present invention;
FIG. 9 is a schematic diagram of a first-order Z-direction bending mode of a moonlet according to an embodiment of the present invention
FIG. 10 is a stress cloud diagram of a structural plate of a structural system of a moonlet according to an embodiment of the present invention;
fig. 11 is a schematic displacement cloud diagram of a structural plate of a structural system of a moonlet according to an embodiment of the present invention;
FIG. 12 is a sweep plot of a base plate of a structural system for a moonlet having a thickness of 0.5mm according to an embodiment of the present invention;
fig. 13 is a sweep curve when the thickness of the skin of the bottom plate of the structural system of the minisatellite provided by the embodiment of the invention is 0.75 mm;
figure 14 illustrates a method of assembling a moonlet according to an embodiment of the present invention.
Detailed Description
The technical solution in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention.
Referring to fig. 2 and 3, a structural system 110 of a moonlet 100 according to an embodiment of the present invention is shown, where the structural system 110 includes:
a base plate 111;
a plurality of side plates 113;
the top plate 112 is provided with a plurality of holes,
wherein the bottom plate 111, the plurality of side plates 113 and the top plate 112 form a prism-like shape (a hexagonal prism-like shape is taken as an example in the figure) when assembled together, the bottom plate 111 and the top plate 112 correspond to the lower bottom surface and the upper bottom surface of the prism-like shape respectively, the plurality of side plates 113 correspond to each side surface of the prism-like shape respectively,
wherein at least one of the plurality of side panels 113 has a side panel interface 113C for installing the subsystem single machine 130 of the small satellite 100 therein.
Different from the prior art, in the structural system of the small satellite provided by the embodiment of the invention, the single machine of the subsystem of the small satellite can be directly installed on the side plate, the single machine installation plate is removed, and the weight of the structural system is reduced to a greater extent.
With respect to the above-described structural system 110, in a preferred embodiment of the present invention, as shown in fig. 3 and 6, the base plate 111 is used for interfacing with a launch vehicle (not shown in the drawings) of the small satellite 100, and the base plate 111 has a base plate interface 111C on the inside for mounting a payload 120 of the small satellite 100.
Different from the prior art, in the structure system of the small satellite provided by the embodiment of the invention, the bottom plate, the top plate and the side plates of the structure system form a closed structure, a structure similar to a force bearing cylinder is integrally formed, a special part for installing an effective load is removed, and the weight of the structure system is further reduced; in addition, under the condition that the bottom plate is used for being in butt joint with a carrier rocket of the small satellite and the effective load of the small satellite is directly installed on the bottom plate, the effect of the shortest force transmission path can be achieved, the effective load can be installed in a mode independent of the side plate and the subsystem single machine installed on the side plate, and therefore mechanical coupling is reduced, and therefore the influence of micro-vibration of the small satellite on the effective load (particularly high-resolution optical load with high requirements on micro-vibration) can be reduced.
In one practical application of the invention, the total weight of the small satellite is 502kg, the weight of the structural system is 69.3kg, and the weight of the payload is 201 kg.
With regard to the above-described structural system 110, in a preferred embodiment of the present invention, as shown in fig. 4 and 5, wherein fig. 4 is a schematic view of the minisatellite in fig. 2 in a side panel unfolded state, and fig. 5 is an enlarged schematic view of a portion corresponding to a broken line frame DF in fig. 4, at least one of the plurality of side panels 113 is assembled to the bottom panel 111 through a hinge H (see fig. 5) in a manner capable of being unfolded and folded. Because the side plate 113 can be expanded relative to the bottom plate 111, the modularized structural layout is convenient to form, and the parallel assembly of the plurality of subsystem single machines 130 is convenient, so that the assembly integration efficiency is greatly improved, and the independent test can be always carried out on each subsystem single machine before the satellite closes the cabin.
With respect to the above-described structural system 110, in a preferred embodiment of the present invention, the bottom plate 111, the top plate 112 and the plurality of side plates 113 have a honeycomb sandwich structure.
In the above structure system 110, as shown in fig. 7, in the preferred embodiment of the present invention, a reinforcing beam B is embedded in the honeycomb sandwich structure of the bottom plate 111. The bottom plate 111 is a key component of the whole star structure and carries both the payload 120 and the mechanical loads of the side plates 113 and the top plate 112. By embedding the reinforcing beam B in the honeycomb sandwich structure of the bottom plate 111, it is ensured that the strength and rigidity of the bottom plate 111 meet the use requirements, and the structural system 110 can meet the strength requirements when the mass of the payload 120 is increased. The skin of the honeycomb sandwich structure of the bottom plate 111 may be made of M55J carbon fiber material, for example, and the thickness of the skin may be arbitrarily selected from 0.3mm to 1mm according to task requirements, and the whole fundamental frequency of the star corresponding to the thickness is changed from 22Hz to 35Hz, thereby enabling parametric design of the structural system 110. In more detail, see fig. 8 and 9, where fig. 8 shows a sweep curve for a thickness of the floor skin of 0.5mm and fig. 9 shows a sweep curve for a thickness of the floor skin of 0.75 mm.
For the above structure system 110, in a preferred embodiment of the present invention, in order to increase the stability of the whole star structure, a reinforcing frame may be embedded inside the side plate 113 to share certain axial and torsional loads. As shown in fig. 8 to 11, the small satellite provided by the embodiment of the present invention is found to satisfy the use requirement after performing both the modal and the static analysis.
With regard to the above structural system 110, in a preferred embodiment of the present invention, as shown in fig. 3, the bottom plate interface 111C includes at least three clip projections P extending from the inner side of the bottom plate 111 toward the top plate 1120.
Embodiments of the present invention also provide a minisatellite 100, and referring to fig. 2 to 7, the minisatellite 100 includes (see fig. 3 in particular):
the structural system 110 described above;
at least one subsystem unit 130, said at least one subsystem unit 130 being mounted to said side panel 113 through said side panel interface 113C.
With respect to the above-described moonlet 100, in a preferred embodiment of the present invention, as shown in fig. 3, the moonlet 100 further includes a payload 120, and the payload 120 is mounted to the backplane 111 through the backplane interface 111C.
In the field of small satellites, the requirement for satellite micro-vibration is high. In the small satellite provided by the embodiment of the invention, the effective load is directly arranged on the bottom plate of the satellite, so that the force transmission path and the mechanical coupling are reduced, and the small satellite is particularly suitable for high-resolution optical loads with higher requirements on micro-vibration of the small satellite.
With respect to the above-described minisatellite 100, in a preferred embodiment of the present invention, as shown in FIG. 3, the minisatellite further comprises at least one solar cell array 140.
With respect to the above-described moonlet 100, in a preferred embodiment of the present invention, the moonlet further includes a reaction flywheel mounted on the side panel 113. This arrangement allows no connection between the side plates 113 and the payload 120 other than the base plate 111, so that micro-vibrations have minimal impact on the imaging in the case of a high resolution optical payload.
With respect to the above-described moonlet 100, in a preferred embodiment of the present invention, and with particular reference to fig. 2 and 7, the moonlet 100 further includes a propellant tank 150, the propellant tank 150 being disposed externally of the structural system 110. Placing the propellant tanks 150 outside the structural system 110, for example in the support cabin of a launch vehicle, allows the height of the side panels 113 of the structural system 110 of the moonlet 100 to be reduced, reducing the height of the moonlet 100 while saving the weight of the structural system 110.
Embodiments of the present invention also provide a method for assembling the above-described moonlet 100, referring to fig. 14, the method including:
s1401: unfolding the side panels 113 by means of a floor device 200 (see fig. 4) until they are in the same horizontal plane as the bottom panel 111;
s1402: mounting the at least one subsystem unit 130 to the side panel 113 in a parallel manner;
s1403: the unfolded side panels 130 are folded up to a position where they should be in a state where the small satellite 100 is assembled.
Installing the payload 120 and the at least one subsystem standalone 130 in a parallel manner enables fast assembly of the moonlet 100.
It should be noted that: the technical schemes described in the embodiments of the present invention can be combined arbitrarily without conflict.
The above description is only for the specific embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily conceive of the changes or substitutions within the technical scope of the present invention, and all the changes or substitutions should be covered within the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the appended claims.
Claims (10)
1. A structural system for a minisatellite, said structural system comprising:
a base plate;
a plurality of side plates;
a top plate is arranged on the top plate,
wherein the bottom plate, the side plates and the top plate form a prism-like shape when assembled together, the bottom plate and the top plate respectively correspond to the lower bottom surface and the upper bottom surface of the prism-like shape, the side plates respectively correspond to each side surface of the prism-like shape,
wherein at least one of the plurality of side plates has a side plate interface inside for mounting a subsystem stand-alone of the minisatellite.
2. The structural system of claim 1, wherein the backplane is for interfacing with a launch vehicle of the moonlet and the backplane has a backplane interface on an interior side for mounting a payload of the moonlet.
3. The structural system of claim 1, wherein at least one of the plurality of side panels is mounted to the bottom panel in an expandable and collapsible manner by a hinge.
4. The structural system of claim 1, wherein the bottom panel, the top panel, and the plurality of side panels have a honeycomb sandwich structure.
5. Structural system according to claim 4, characterised in that the honeycomb sandwich structure of the bottom plate has embedded therein a reinforcing beam.
6. A minisatellite, comprising:
structural system according to any one of claims 1 to 5;
at least one subsystem unit, the at least one subsystem unit passes through the curb plate interface install to the curb plate.
7. The moonlet of claim 6, further comprising a payload mounted to the backplane through the backplane interface.
8. The microsatellite according to claim 6 further comprising a reaction flywheel mounted on said side panels.
9. The microsatellite according to claim 6 further comprising a propellant tank disposed outside said structural system.
10. A method of assembling a minisatellite according to any one of claims 6 to 9 comprising:
unfolding the side plates by means of ground equipment until the side plates are in the same horizontal plane with the bottom plate;
mounting the at least one subsystem stand-alone to the side panel in a parallel manner;
and folding the unfolded side plates until the side plates are positioned in a position where the small satellite is assembled.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011177171.3A CN112249364A (en) | 2020-10-28 | 2020-10-28 | Structural system of a minisatellite, a minisatellite and a method of assembling a minisatellite |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011177171.3A CN112249364A (en) | 2020-10-28 | 2020-10-28 | Structural system of a minisatellite, a minisatellite and a method of assembling a minisatellite |
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| CN112249364A true CN112249364A (en) | 2021-01-22 |
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| CN202011177171.3A Pending CN112249364A (en) | 2020-10-28 | 2020-10-28 | Structural system of a minisatellite, a minisatellite and a method of assembling a minisatellite |
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Citations (9)
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|---|---|---|---|---|
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| CN108438253A (en) * | 2018-03-06 | 2018-08-24 | 航天东方红卫星有限公司 | A kind of microsatellite structure suitable for optics load |
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-
2020
- 2020-10-28 CN CN202011177171.3A patent/CN112249364A/en active Pending
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| Title |
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