CN117699063A - Spacecraft thermal control design method based on mechanical-thermal integrated structure - Google Patents
Spacecraft thermal control design method based on mechanical-thermal integrated structure Download PDFInfo
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- CN117699063A CN117699063A CN202311440476.2A CN202311440476A CN117699063A CN 117699063 A CN117699063 A CN 117699063A CN 202311440476 A CN202311440476 A CN 202311440476A CN 117699063 A CN117699063 A CN 117699063A
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- 238000000034 method Methods 0.000 title claims abstract description 17
- 238000010521 absorption reaction Methods 0.000 claims abstract description 22
- 239000011248 coating agent Substances 0.000 claims abstract description 20
- 238000000576 coating method Methods 0.000 claims abstract description 20
- 238000004519 manufacturing process Methods 0.000 claims abstract description 8
- QGZKDVFQNNGYKY-UHFFFAOYSA-N Ammonia Chemical compound N QGZKDVFQNNGYKY-UHFFFAOYSA-N 0.000 claims description 18
- 238000010146 3D printing Methods 0.000 claims description 8
- 238000009434 installation Methods 0.000 claims description 7
- 238000005516 engineering process Methods 0.000 claims description 5
- 229920001296 polysiloxane Polymers 0.000 claims description 3
- 239000013354 porous framework Substances 0.000 claims description 3
- 239000002775 capsule Substances 0.000 claims 2
- 238000009413 insulation Methods 0.000 abstract description 8
- 230000009286 beneficial effect Effects 0.000 abstract description 2
- 238000009826 distribution Methods 0.000 abstract description 2
- 230000005855 radiation Effects 0.000 description 7
- 230000017525 heat dissipation Effects 0.000 description 4
- 239000003814 drug Substances 0.000 description 3
- 230000004907 flux Effects 0.000 description 3
- 238000010438 heat treatment Methods 0.000 description 3
- 238000007789 sealing Methods 0.000 description 3
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 description 2
- 229910052782 aluminium Inorganic materials 0.000 description 2
- 238000001816 cooling Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000004519 grease Substances 0.000 description 2
- 230000003647 oxidation Effects 0.000 description 2
- 238000007254 oxidation reaction Methods 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 238000004458 analytical method Methods 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 238000010923 batch production Methods 0.000 description 1
- 238000004364 calculation method Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 230000009977 dual effect Effects 0.000 description 1
- 238000002309 gasification Methods 0.000 description 1
- 229910002804 graphite Inorganic materials 0.000 description 1
- 239000010439 graphite Substances 0.000 description 1
- 238000005286 illumination Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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- Application Of Or Painting With Fluid Materials (AREA)
Abstract
The invention provides a spacecraft thermal control design method based on a mechanical-thermal integrated structure, which comprises the following steps of: step 1: manufacturing a main body structure of the spacecraft; step 2: the main structure is sprayed or attached by adopting a thermal control coating with low solar absorption ratio and high infrared emissivity. The beneficial effects of the invention are as follows: according to the invention, the low solar absorption ratio high infrared emissivity coating is designed at the required position of the spacecraft and is matched with the original surface conductive oxide surface, and the equivalent solar absorption ratio and infrared emissivity are formed by designing the area proportion and distribution of the low solar absorption ratio high infrared emissivity coating and the low solar absorption ratio high infrared emissivity coating, so that the whole-star temperature level control is realized, and the influence of the heat flow outside the orbit on the temperature of the spacecraft is weakened; on the premise of not adding an external heat conduction device, the condition that the temperature gradient of the side with larger external heat flow and the side with smaller external heat flow of the spacecraft is overlarge is effectively avoided by utilizing the high heat conductivity of the mechanical-thermal integrated structure, the use of the traditional multi-layer heat insulation assembly can be canceled, and the temperature index requirement of the internal equipment of the spacecraft is met.
Description
Technical Field
The invention relates to the technical field of aerospace, in particular to a spacecraft thermal control design method based on a mechanical-thermal integrated 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 internal electronics is ensured to meet certain requirements.
The heat insulation assembly is used for insulating the external heat flow of the orbit where the spacecraft is located and the deep space cold background, the workload of the heat control implementation stage is large and the automatic production cannot be realized in the side satellite assembly stage with small total external heat flow value and small variation fluctuation, and the production efficiency is low, so that the existing design production mode cannot meet the requirements of quick development of commercial aerospace.
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 paths, 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.
Description of prior art problems and defects: the conventional heat design of the spacecraft only focuses on practicality, and the post-heating heat control multi-layer heat insulation assembly and the heat pipe are utilized as main means for heat insulation and heat conduction, so that the batch automatic production requirement of the commercial satellite and the high inheritance integrated design concept are not fully considered. The conventional heat design method brings about the increase of the heat control implementation period in the final assembly stage of the spacecraft, the soft multi-layer heat insulation assembly is more dependent on manual assembly, and meanwhile, the hooking risk is easily generated on the outer layers of the spacecraft, so that the batch emission is not facilitated; the non-integrated design leads to an increase of coordination interfaces, so that the engineering complexity is enhanced. At this time, a new heat control design method is needed, so that normal temperature control requirements of the spacecraft are met, and complexity of assembly operation of the spacecraft is reduced.
Disclosure of Invention
The invention provides a spacecraft thermal control design method based on a mechanical-thermal integrated structure, which comprises the following steps of:
step 1: manufacturing a main body structure of the spacecraft;
step 2: the main structure is sprayed or attached by adopting a thermal control coating with low solar absorption ratio and high infrared emissivity.
In the step 1, a main structure of the spacecraft is manufactured, a porous skeleton structure is arranged in the main structure, the main structure is vacuumized after being filled with liquid ammonia, the liquid ammonia is heated and gasified to absorb heat at the installation position of equipment, and ammonia gas is changed into saturated liquid ammonia when moving to a region with lower temperature, so that a large amount of latent heat is released; when the liquid ammonia with certain surface tension contacts the porous framework, certain capillary pump force can be generated, and then the liquid ammonia can flow back to the installation position of the equipment under the action of the capillary pump force, so that circulation is formed, and the temperature of the equipment is maintained within a certain temperature range.
As a further improvement of the present invention, in the step 1, the main structure of the spacecraft is integrally formed by a 3D printing technology, and the porous skeleton structure is printed in the main structure.
As a further improvement of the invention, the thermal conductivity of the host structure is greater than 5000W/m/K.
As a further improvement of the invention, the spacecraft also comprises a cabin sealing structure, wherein the cabin sealing structure is contacted with the main body structure and is coated with heat-conducting silicone grease to strengthen the contact heat exchange capability.
As a further development of the invention, in said step 2, the structural surface thermal parameters of the sunny side of the spacecraft are controlled by designing the area ratio of the thermal control coating.
As a further improvement of the present invention, the solar absorption/infrared emissivity of the thermally controlled coating is less than 1.
As a further improvement of the invention, in said step 2, the desired equivalent solar absorption ratio and infrared emissivity is achieved by controlling the area of the thermally controlled coating.
The beneficial effects of the invention are as follows: according to the invention, the low solar absorption ratio high infrared emissivity coating is designed at the required position of the spacecraft and is matched with the original surface conductive oxide surface, and the equivalent solar absorption ratio and infrared emissivity are formed by designing the area proportion and distribution of the low solar absorption ratio high infrared emissivity coating and the low solar absorption ratio high infrared emissivity coating, so that the whole-star temperature level control is realized, and the influence of the heat flow outside the orbit on the temperature of the spacecraft is weakened; on the premise of not adding an external heat conduction device, the condition that the temperature gradient of the side with larger external heat flow and the side with smaller external heat flow of the spacecraft is overlarge is effectively avoided by utilizing the high heat conductivity of the mechanical-thermal integrated structure, the use of the traditional multi-layer heat insulation assembly can be canceled, and the temperature index requirement of the internal equipment of the spacecraft is met.
Drawings
FIG. 1 is a schematic diagram of a thermal integrated structure based on 3D printing;
FIG. 2 is a schematic diagram of the design of the overall temperature control of the spacecraft.
Detailed Description
The invention discloses a spacecraft thermal control design method based on a mechanical-thermal integrated structure, which comprises the following steps of:
step 1: manufacturing a main body structure of the spacecraft;
step 2: the main structure is sprayed or attached by adopting the thermal control coating with low solar absorption ratio and high infrared emissivity, so that the surface thermal parameters (solar absorption ratio and infrared emissivity) of the main structure after conducting oxidation treatment are changed, and further the comprehensive temperature influence of the out-of-orbit heat flow and the deep space extreme cold background on the star is controlled, and the overall temperature of the spacecraft is at a good temperature level.
In the step 1, the main structure of the spacecraft is integrally formed by a 3D printing technology, a porous skeleton structure is printed in the main structure, a certain amount of liquid ammonia is filled in the main structure, then the main structure is vacuumized, the heat of the installation position of the equipment is absorbed by utilizing the heated gasification of the liquid ammonia, and the ammonia gas is transformed into saturated liquid ammonia when moving to a region with lower temperature, so that a large amount of latent heat is released; when the liquid ammonia with certain surface tension contacts the porous framework, certain capillary pump force can be generated, and then the liquid ammonia can flow back to the installation position of the equipment under the action of the capillary pump force, so that circulation is formed, and the temperature of the equipment is maintained within a certain temperature range.
In step 1, the main structure of the spacecraft is integrally printed by a 3D printing technology, the thermal conductivity of the structure is larger than 5000W/m/K, the cabin sealing structure and the main structure are in contact with each other in a larger area, and heat conduction silicone grease is coated to strengthen contact heat exchange capacity, so that the overall temperature equalization design of the satellite is realized, and the design principle is shown in figure 1. In step 2, the structural surface thermal parameters of the sunny side of the spacecraft are controlled by designing the area proportion of the thermal control coating (the solar absorption ratio/infrared emissivity is smaller than 1), so that the heat exchange capacity of the whole external heat flow absorbed by the spacecraft and the heat exchange capacity of the whole deep space extremely cold background are controlled, the whole temperature level of the spacecraft is at a good temperature level, and the design principle is shown in figure 2.
The realization principle is as follows: the whole samming of spacecraft is realized mainly through the heat conduction capability of structure itself, and this part is the mechatronic structure based on 3D printing technique, and its heat conduction heat dissipation capability:
wherein: q is the heat conduction and heat dissipation capacity of the integrated structure, and the unit is W;
k is the heat conductivity of the integrated structure, and the unit is W/square meter/DEG C;
l is the heat conduction path distance, unit m;
s is the sectional area of a heat conduction path of an integrated structure, and the unit square meter;
delta T is the temperature difference between the heat source position and the cold end of the integrated structure, and is in units of ℃.
The temperature level of the whole spacecraft is balanced by controlling the self heating value, the total absorption amount of heat flow outside the orbit and the radiant heat of the deep space cold background, and the energy balance formula is as follows:
Q outer part +Q Inner part ＝Q Powder medicine
Wherein Q is Outer part Out-of-orbit heat flow absorbed by spacecraft, Q Inner part For the heating value of the internal equipment of the spacecraft, Q Powder medicine The unit of the radiating heat of the spacecraft to the deep air cooling background is W.
Q Outer part Including the ability to absorb direct solar radiation, the energy of the planet's albedo to the sun's light, and the energy of the planet's own infrared radiation to the spacecraft.
Q Outer part ＝q Straight line ×α s ×S Straight line +q Reverse-rotation ×α s ×S Reverse-rotation +q Red colour ×ε×S Red colour
Wherein: q Straight line For heat flux density of solar radiation to spacecraft, q Reverse-rotation Heat flux density, q, of planet-pair solar radiation energy reflected to spacecraft Red colour The heat flux density of the infrared energy emitted by the planet to the spacecraft is W/square meter;
α s the solar absorption ratio is the thermal parameter of the surface of the spacecraft subjected to solar direct illumination and planetary albedo;
epsilon is the infrared emissivity of the outer surface of the spacecraft;
s is the area of each face.
Q Powder medicine For the radiant heat of each surface of the spacecraft facing the deep air cooling background, if the surface infrared emissivity of each surface is different, the calculation and the summation are needed to be carried out on each surface respectively.
Wherein: epsilon is the infrared emissivity of each surface of the spacecraft;
sigma is the radiation constant of a blackbody (stefin-boltzmann constant);
s is the radiation surface area of each surface of the spacecraft;
t is the absolute temperature of the spacecraft surface, unit K.
Through analysis, the invention adopts a mechanical-thermal integrated structure based on a 3D printing technology, and can realize the temperature equalization design of the whole spacecraft by utilizing the high heat conduction and bearing dual functions of the mechanical-thermal integrated structure, and the temperature control of the whole spacecraft can be realized by reasonably designing the equivalent solar absorption ratio and infrared emissivity of the surface of the spacecraft.
In summary, the invention constructs a satellite main structure frame by utilizing a high heat conduction and high bearing integrated structure based on 3D printing, and then combines a low infrared emissivity and high infrared emissivity thermal control coating with an industrial conductive oxidation coating, and realizes the expected equivalent solar absorption ratio and infrared emissivity by controlling the area of the thermal control coating, thereby leading the internal equipment of the spacecraft to be in a better temperature range, avoiding the complex work of the installation of the multilayer heat insulation component of the spacecraft, and being better suitable for short-period and batch production of the spacecraft.
The foregoing is a further detailed description of the invention in connection with the preferred embodiments, and it is not intended that the invention be limited to the specific embodiments described. It will be apparent to those skilled in the art that several simple deductions or substitutions may be made without departing from the spirit of the invention, and these should be considered to be within the scope of the invention.
Claims (8)
1. A spacecraft thermal control design method based on a mechanical-thermal integrated structure is characterized by comprising the following steps:
step 1: manufacturing a main body structure of the spacecraft;
step 2: the main structure is sprayed or attached by adopting a thermal control coating with low solar absorption ratio and high infrared emissivity.
2. The method for designing the thermal control of the spacecraft according to claim 1, wherein in the step 1, a main body structure of the spacecraft is manufactured, a porous skeleton structure is arranged in the main body structure, the main body structure is filled with liquid ammonia and then vacuumized, the liquid ammonia is heated and gasified to absorb heat at the installation position of the equipment, and ammonia gas moves to a region with lower temperature to be transformed into saturated liquid ammonia, so that a large amount of latent heat is released; when the liquid ammonia with certain surface tension contacts the porous framework, certain capillary pump force can be generated, and then the liquid ammonia can flow back to the installation position of the equipment under the action of the capillary pump force, so that circulation is formed, and the temperature of the equipment is maintained within a certain temperature range.
3. The method according to claim 2, wherein in the step 1, the main structure of the spacecraft is integrally formed by a 3D printing technology, and the porous skeleton structure is printed in the main structure.
4. The spacecraft thermal control design method of claim 1, wherein the thermal conductivity of the host structure is greater than 5000W/m/K.
5. The method of claim 1, further comprising a capsule structure, wherein the capsule structure is in contact with the body structure and is coated with a heat conductive silicone to enhance contact heat transfer capability.
6. The method according to claim 1, wherein in the step 2, the structural surface thermal parameters of the sunny side of the spacecraft are controlled by designing the area ratio of the thermal control coating.
7. The method of thermal control design of a spacecraft of claim 6, wherein the solar absorption/infrared emissivity of the thermal control coating is less than 1.
8. The spacecraft thermal control design method of claim 1, wherein in said step 2, the desired equivalent solar absorption ratio and infrared emissivity are achieved by controlling the thermal control coating area.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311440476.2A CN117699063A (en) | 2023-10-31 | 2023-10-31 | Spacecraft thermal control design method based on mechanical-thermal integrated structure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311440476.2A CN117699063A (en) | 2023-10-31 | 2023-10-31 | Spacecraft thermal control design method based on mechanical-thermal integrated structure |
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| CN117699063A true CN117699063A (en) | 2024-03-15 |
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| CN202311440476.2A Pending CN117699063A (en) | 2023-10-31 | 2023-10-31 | Spacecraft thermal control design method based on mechanical-thermal integrated structure |
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- 2023-10-31 CN CN202311440476.2A patent/CN117699063A/en active Pending
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