CN220465799U - Spacecraft uniform temperature heating main body structure - Google Patents
Spacecraft uniform temperature heating main body structure Download PDFInfo
- Publication number
- CN220465799U CN220465799U CN202321868491.2U CN202321868491U CN220465799U CN 220465799 U CN220465799 U CN 220465799U CN 202321868491 U CN202321868491 U CN 202321868491U CN 220465799 U CN220465799 U CN 220465799U
- Authority
- CN
- China
- Prior art keywords
- spacecraft
- body structure
- main body
- heating
- metal cavity
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000010438 heat treatment Methods 0.000 title claims abstract description 56
- 229910052751 metal Inorganic materials 0.000 claims abstract description 51
- 239000002184 metal Substances 0.000 claims abstract description 51
- QGZKDVFQNNGYKY-UHFFFAOYSA-N Ammonia Chemical compound N QGZKDVFQNNGYKY-UHFFFAOYSA-N 0.000 claims abstract description 23
- 239000011888 foil Substances 0.000 claims description 22
- 230000001681 protective effect Effects 0.000 claims description 18
- 230000017525 heat dissipation Effects 0.000 claims description 10
- 238000003466 welding Methods 0.000 claims description 9
- 239000011248 coating agent Substances 0.000 claims description 8
- 238000000576 coating method Methods 0.000 claims description 8
- 229910000914 Mn alloy Inorganic materials 0.000 claims description 7
- UTICYDQJEHVLJZ-UHFFFAOYSA-N copper manganese nickel Chemical compound [Mn].[Ni].[Cu] UTICYDQJEHVLJZ-UHFFFAOYSA-N 0.000 claims description 7
- 229920001721 polyimide Polymers 0.000 claims description 4
- 239000004519 grease Substances 0.000 claims description 3
- 229920001296 polysiloxane Polymers 0.000 claims description 3
- 238000010276 construction Methods 0.000 claims description 2
- 230000009286 beneficial effect Effects 0.000 description 7
- 229910052782 aluminium Inorganic materials 0.000 description 2
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000007704 transition Effects 0.000 description 2
- 238000010146 3D printing Methods 0.000 description 1
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 229910002804 graphite Inorganic materials 0.000 description 1
- 239000010439 graphite Substances 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
- 229910000679 solder Inorganic materials 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Landscapes
- Surface Heating Bodies (AREA)
Abstract
The utility model discloses a spacecraft uniform temperature heating main body structure which comprises a diversion trench, liquid ammonia, a heating film and a metal cavity body in a closed structure, wherein a hollow accommodating cavity is formed by enclosing the inside of the metal cavity body, the liquid ammonia is filled in the hollow accommodating cavity, the diversion trench is further arranged on the inner wall of the metal cavity body, and the heating film is arranged between the inner wall and the outer wall of the metal cavity body. The heating film is electrified to generate heat, so that the liquid ammonia in the hollow accommodating cavity in the metal cavity is heated and vaporized, the vaporized liquid ammonia forms circulation on the inner wall of the metal cavity through the diversion trench, and the heat is timely transferred to the whole spacecraft temperature-equalizing heating main body structure through the fixed circulation path, so that the spacecraft temperature-equalizing heating main body structure has stronger heat conduction performance and heat exchange capacity, the constant temperature and heating of the spacecraft can be realized quickly, and the normal operation of internal devices of the spacecraft at the conventional temperature is ensured.
Description
Technical Field
The utility model relates to the technical field of spacecraft equipment, in particular to a spacecraft uniform temperature heating main body structure.
Background
Spacecraft typically employ multi-layer insulation assemblies to attenuate heat exchange with spatially orbitally external heat flows (including solar direct, earth specular and planetary infrared radiation) and deep air-cooled backgrounds. The heat dissipation and temperature equalization are carried out on the electronic equipment inside the spacecraft by adopting the embedded heat pipe or the externally-attached high heat conduction film, so that the temperature of the electronic equipment inside the spacecraft is ensured to meet certain requirements.
At present, a satellite structure adopts a light aluminum honeycomb structural plate, so that the heat conductivity is low, and the effective heat dissipation cannot be carried out on equipment with high heat consumption; the heat conduction between the aluminum honeycomb structural plates only depends on screws to generate weak heat conduction path strength, and the heat conduction and temperature uniformity capability between the plates is almost negligible. The embedded heat pipe or the high heat conduction graphite film is adopted to enhance the heat conduction and temperature uniformity of the structural plate, the heat conduction performance is enhanced by adopting the attached heat pipe or the high heat conduction film between the plates, a certain satellite space is occupied, and a certain design constraint is brought to the high inheritance of satellites and the automatic assembly. It can be seen that there is a need for improvements and improvements in the art.
Disclosure of Invention
In view of the shortcomings of the prior art, the utility model aims to provide a spacecraft uniform temperature heating main body structure which is used for solving the problem that the heat conduction capacity of a spacecraft in the prior art is poor.
In order to achieve the above purpose, the utility model adopts the following technical scheme: the utility model provides a spacecraft samming heating major structure, includes guiding gutter, liquid ammonia, heating film and is the metal cavity of enclosed construction, a cavity holding chamber is enclosed to the inside of metal cavity, cavity holding intracavity is filled with liquid ammonia, still be provided with on the inner wall of metal cavity the guiding gutter, be equipped with between the inner wall of metal cavity and the outer wall the heating film.
In an embodiment of the utility model, the heating film further comprises welding spots and power lines, wherein the heating film comprises an upper layer of protective film, a metal foil and a lower layer of protective film which are arranged in a stacked mode, the welding spots are arranged at two ends of the metal foil, and each welding spot is connected with one power line.
The beneficial effects of the above embodiment are that: the upper protective film and the lower protective film can form good protection for the metal foil between the upper protective film and the lower protective film, and the service life of the metal foil is effectively prolonged.
In an embodiment of the present utility model, the upper protective film and the lower protective film are both polyimide films, and the metal foil is a copper nickel manganese alloy foil.
The beneficial effects of the above embodiment are that: the metal foil made of the copper-nickel-manganese alloy has certain conductivity and resistance, and can immediately generate heat in an electrified state to heat the spacecraft.
In an embodiment of the utility model, the thickness of the heating film is 0.09mm to 0.14mm.
The beneficial effects of the above embodiment are that: the heating film with the thickness of 0.09mm to 0.14mm has smaller volume, can be flexibly applied to spacecrafts, and has stronger applicability.
In an embodiment of the utility model, the device further comprises blind holes arranged in the diversion trench, and the blind holes are uniformly distributed along the diversion trench.
In an embodiment of the utility model, a diameter of the blind hole is smaller than a width of the diversion trench.
The beneficial effects of the above embodiment are that: the diameter of the blind hole is smaller than the width of the diversion trench, liquid ammonia flows in the diversion trench after being heated and vaporized, and the surface tension of the liquid ammonia accelerates the flow under the action of the capillary pump force of the blind hole when flowing through the edge of the blind hole, so that the liquid ammonia forms circulation in the hollow accommodating cavity.
In an embodiment of the present utility model, the number of the diversion trenches is a plurality, and the diversion trenches are staggered.
The beneficial effects of the above embodiment are that: the staggered diversion trenches provide various flow channels for circulation and are helpful for forming a closed loop circulation path.
In an embodiment of the utility model, the metal cavity is an integrally formed structure.
In an embodiment of the utility model, the heat dissipation coating is coated on the outer wall of the metal cavity.
The beneficial effects of the above embodiment are that: the heat exchange capacity of the spacecraft temperature-equalizing heating main body structure can be enhanced by the heat dissipation coating on the outer wall of the metal cavity.
In an embodiment of the utility model, the heat dissipation coating is a heat-conductive silicone grease.
As described above, the spacecraft homogeneous heating main body structure has the following beneficial effects: the heating film is electrified to generate heat, so that liquid ammonia in a hollow accommodating cavity in the metal cavity is heated and vaporized, the vaporized liquid ammonia forms circulation on the inner wall of the metal cavity through the diversion trench, and the heat is timely transferred to the whole spacecraft temperature-equalizing heating main body structure through a fixed circulation path, so that the spacecraft temperature-equalizing heating main body structure has stronger heat conduction performance and heat exchange capacity, constant temperature and heating of the spacecraft can be realized rapidly, and normal operation of internal devices of the spacecraft at the conventional temperature is ensured.
Drawings
In order to more clearly illustrate the embodiments of the utility model or the technical solutions in the prior art, the drawings that are required in the embodiments or the description of the prior art will be briefly described, it being obvious that the drawings in the following description are only some embodiments of the utility model, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic structural view of a spacecraft homogeneous heating main body structure provided by the utility model;
fig. 2 is a schematic structural diagram of a heating film according to the present utility model.
Description of main reference numerals: 1-heating film, 2-metal cavity, 3-cavity holding chamber, 4-upper protective film, 5-metal foil, 6-lower protective film, 7-solder joint, 8-power cord, 9-blind hole.
Detailed Description
The utility model provides a spacecraft homogeneous heating main body structure, which is used for making the purposes, technical schemes and effects of the utility model clearer and more definite, and the utility model is further described in detail below by referring to the accompanying drawings and the embodiments.
In the description of the present utility model, it should be understood that the azimuth or positional relationship indicated by the terms "up, down, left, right", etc. are based on the azimuth or positional relationship shown in the drawings, and are merely for convenience in describing the present utility model and for simplifying the description, and are not to be construed as limiting the present utility model; furthermore, the terms "mounted," "connected," and the like, are to be construed broadly and, as appropriate, the specific meaning of the terms in the present utility model will be understood by those of ordinary skill in the art.
Referring to fig. 1 and 2, the utility model provides a spacecraft homogeneous heating main body structure, which comprises a diversion trench (not shown), liquid ammonia, a heating film 1 and a metal cavity 2 with a closed structure, wherein a hollow accommodating cavity 3 is enclosed in the metal cavity 2, the liquid ammonia is filled in the hollow accommodating cavity 3, the diversion trench is further arranged on the inner wall of the metal cavity 2, and the heating film 1 is arranged between the inner wall and the outer wall of the metal cavity 2.
In this embodiment, the heating film 1 includes an upper protective film 4, a metal foil 5, a lower protective film 6, a welding spot 7, and a power line 8, where the welding spots 7 are disposed at two ends of the metal foil 5, and each welding spot 7 is connected with one power line 8; it can be appreciated that the upper protective film 4 and the lower protective film 6 can form good protection for the metal foil 5 therebetween, and effectively prolong the service life of the metal foil 5. Optionally, the upper protective film 4 and the lower protective film 6 are polyimide films, and the metal foil 5 is made of copper-nickel-manganese alloy, i.e. the metal foil 5 is a copper-nickel-manganese alloy foil. In this embodiment, the polyimide film can tightly wrap the metal foil 5 made of the copper-nickel-manganese alloy to protect the metal foil, and the metal foil 5 made of the copper-nickel-manganese alloy has certain conductivity and resistance, and can immediately generate heat in an electrified state to heat the spacecraft.
Optionally, the thickness of the heating film 1 is 0.09mm to 0.14mm, and the heating film 1 with the thickness of 0.09mm to 0.14mm has smaller volume, can be flexibly applied to a spacecraft, and has stronger applicability. In this embodiment, the thickness of the heating film 1 is 0.1mm.
The device further comprises blind holes 9 arranged in the diversion trenches, wherein the blind holes 9 are uniformly distributed along the diversion trenches; the diameter of the blind hole 9 is smaller than the width of the diversion trench; the diameter of the blind hole 9 is smaller than the width of the diversion trench, liquid ammonia flows in the diversion trench after being heated and vaporized, and the surface tension of the liquid ammonia accelerates to flow under the action of capillary pump force of the blind hole 9 when flowing through the edge of the blind hole 9, so that the liquid ammonia forms circulation in the hollow accommodating cavity 3. Specifically, the number of the diversion trenches is multiple, the diversion trenches are staggered, it can be understood that the two adjacent diversion trenches are in smooth transition, the smooth transition diversion trenches facilitate the flow of liquid ammonia, and the circulation of liquid ammonia in the hollow accommodating cavity 3 is quickened; and the staggered diversion trenches provide various flow channels for liquid ammonia circulation, which is helpful for forming a closed loop circulation path.
Optionally, the metal cavity 2 is integrated into one piece structure, and integrated into one piece's structure can the rapid prototyping, can reduce the degree of difficulty and the cost of production, improves spacecraft samming heating main body structure's production efficiency. Preferably, the metal cavity 2 is integrally formed by 3D printing.
In this embodiment, the heat dissipation device further includes a heat dissipation coating coated on the outer wall of the metal cavity 2, where the heat dissipation coating on the outer wall of the metal cavity 2 can enhance the heat exchange capability of the spacecraft homogeneous heating main structure; optionally, the heat dissipation coating is a heat conductive silicone grease.
In summary, according to the spacecraft temperature-equalizing heating main body structure disclosed by the utility model, heat is generated after the heating film 1 is electrified, so that the liquid ammonia in the hollow accommodating cavity 3 in the metal cavity 2 is heated and vaporized, the vaporized liquid ammonia forms a circulation on the inner wall of the metal cavity 2 through the diversion trench, and the heat is timely transferred to the whole spacecraft temperature-equalizing heating main body structure through the fixed circulation path, so that the spacecraft temperature-equalizing heating main body structure has stronger heat conduction performance and heat exchange capacity, the constant temperature and heating of the spacecraft can be realized quickly, and the normal operation of internal devices of the spacecraft at the conventional temperature is ensured. Therefore, the utility model effectively overcomes various defects in the prior art and has high industrial utilization value.
It will be understood that equivalents and modifications will occur to those skilled in the art based on the present utility model and its spirit, and all such modifications and substitutions are intended to be included within the scope of the present utility model.
Claims (10)
1. The utility model provides a spacecraft samming heating major structure, its characterized in that includes guiding gutter, liquid ammonia, heating film and is the metal cavity of enclosed construction, a cavity holding chamber is enclosed to the inside of metal cavity, cavity holding intracavity is filled with liquid ammonia, be provided with on the inner wall of metal cavity the guiding gutter, be equipped with between the inner wall of metal cavity and the outer wall the heating film.
2. The spacecraft homogeneous heating main body structure according to claim 1, further comprising welding spots and power wires, wherein the heating film comprises an upper layer protection film, a metal foil and a lower layer protection film which are arranged in a laminated mode, the welding spots are arranged at two ends of the metal foil, and each welding spot is connected with one power wire.
3. The spacecraft homogeneous heating main body structure according to claim 2, wherein the upper protective film and the lower protective film are polyimide films, and the metal foil is a copper-nickel-manganese alloy foil.
4. The spacecraft homogeneous heating body structure of claim 1, wherein the thickness of the heating film is from 0.09mm to 0.14mm.
5. The spacecraft homogeneous heating main body structure according to claim 1, further comprising blind holes arranged in the diversion trench, wherein the blind holes are uniformly distributed along the diversion trench.
6. The spacecraft homogeneous heating main body structure according to claim 5, wherein the diameter of the blind hole is smaller than the width of the diversion trench.
7. The spacecraft homogeneous heating main body structure according to claim 5, wherein the number of the diversion trenches is a plurality, and a plurality of diversion trenches are arranged in a staggered manner.
8. The spacecraft homogeneous heating main body structure of claim 1, wherein the metal cavity is an integrally formed structure.
9. The spacecraft homogeneous heating body structure of claim 1, further comprising a heat dissipating coating applied to an outer wall of the metal cavity.
10. The spacecraft homogeneous heating body structure of claim 9, wherein the heat dissipation coating is a thermally conductive silicone grease.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202321868491.2U CN220465799U (en) | 2023-07-14 | 2023-07-14 | Spacecraft uniform temperature heating main body structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202321868491.2U CN220465799U (en) | 2023-07-14 | 2023-07-14 | Spacecraft uniform temperature heating main body structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN220465799U true CN220465799U (en) | 2024-02-09 |
Family
ID=89797846
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202321868491.2U Active CN220465799U (en) | 2023-07-14 | 2023-07-14 | Spacecraft uniform temperature heating main body structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN220465799U (en) |
-
2023
- 2023-07-14 CN CN202321868491.2U patent/CN220465799U/en active Active
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN204230382U (en) | A kind of heating panel and electrokinetic cell system | |
| CN107046144A (en) | Supply unit and electric power system | |
| US11611119B2 (en) | Battery module, battery device, and battery system having thermal management design | |
| CN108336048B (en) | Heat superconducting fin radiator with phase change heat storage function | |
| CN107275559B (en) | Battery pack device | |
| CN205491593U (en) | A heat radiation structure and cooling parts for electronic equipment | |
| JP2018088305A (en) | Cooling system | |
| CN203040093U (en) | Radiating device applied to high-power heat source | |
| CN105870373A (en) | Radiating bracket for placing cylindrical batteries in cylindrical battery pack | |
| CN206388796U (en) | Supply unit and electric power system | |
| CN110380157A (en) | A kind of battery thermal management system coupled based on liquid cooling and phase-change thermal storage | |
| CN210430029U (en) | Plate-type heating and cooling heat conduction device and temperature-controllable lithium battery pack adopting same | |
| CN206040903U (en) | Electric automobile cylindricality group battery water cooling jacket | |
| CN220465799U (en) | Spacecraft uniform temperature heating main body structure | |
| CN105870538A (en) | Electric automobile battery pack heat management structure | |
| CN216597749U (en) | Power battery thermal management system | |
| CN209434589U (en) | A kind of high-power heating device heat transmission cooling plate bodies, cooling structure and the superpower laser containing the cooling structure | |
| CN116780038A (en) | Lithium battery protection board and lithium battery device with same | |
| CN216960588U (en) | Heat dissipation device and electronic equipment | |
| CN110350270A (en) | The cooling heat-transfer device of plate type heating and the controllable temperature lithium battery group using the device | |
| CN216145679U (en) | Thermal safety management system and battery | |
| CN205847818U (en) | The working medium cooling system by contact of high power device | |
| CN114082854A (en) | Manufacturing method of battery shell, battery shell and battery | |
| CN213279515U (en) | Cooling device, inverter, and cooling system | |
| CN211606913U (en) | Easy radiating circuit board |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| GR01 | Patent grant | ||
| GR01 | Patent grant |