Thermal control design method for cabin opening type spacecraft

Abstract

The invention provides a thermal control design method for a cabin opening type spacecraft. The thermal control design method comprises the steps that 1, a main body structure of the spacecraft is integrally formed through an aluminum alloy machining process; and 2, spraying or mounting each surface of the main body structure by adopting a thermal control coating with low solar absorptivity and high infrared emissivity. The invention has the following beneficial effects: by designing the relatively strong multi-layer flexible cabin plate on the side surface with relatively large external heat flow of the spacecraft, the sun direct illumination heat flow isolation of the maximum heat flow in the external heat flow of the orbit is realized, so that the influence of the external heat flow on the temperature of the spacecraft is greatly weakened; on the premise that an external heat conduction device does not need to be additionally arranged, the situation that the temperature gradient of the side with the large external heat flow and the side with the small external heat flow of the spacecraft is too large is effectively avoided through the high heat conduction performance of the integrated structure; a thermal control film is mounted on a traditional industrial conductive anodic oxidation coating or thermal control paint is sprayed on the traditional industrial conductive anodic oxidation coating, so that the solar absorptivity and the infrared emissivity of the five outer side walls of the whole satellite main structure are adjusted, and the temperature of spacecraft equipment meets the corresponding temperature index requirement.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and in particular to a thermal control design method for an open-cabin spacecraft. Background Art

[0002] Spacecraft generally use lightweight aluminum honeycomb materials as the main structure of the satellite, which has low thermal conductivity. Each panel needs to be designed for radiation heat exchange and thermal conductivity separately. Almost every panel is an independent thermal control design object. By controlling the size of the heat dissipation surface design of each panel, multi-layer insulation component coating and enhanced temperature uniformity, combined with the extra-orbital heat flux reaching each panel (including direct sunlight, earth reflection and planetary infrared radiation) and the heat consumption of the instruments and equipment on each panel, the instruments and equipment on each panel can operate within a certain temperature range.

[0003] Multi-layer insulation components are used to insulate the spacecraft from the heat flux outside its orbit and the cold background of deep space. In particular, the insulation design of extravehicular instruments and equipment is complex. The workload of the thermal control implementation phase during the satellite assembly phase is large and automated production cannot be achieved, resulting in low production efficiency. As a result, the existing design and production model is unable to adapt to the needs of rapid development of commercial aerospace.

[0004] At present, satellite structures use lightweight aluminum honeycomb structural panels, which have low thermal conductivity and cannot form a whole satellite heat bus. They cannot effectively dissipate heat for equipment with high heat consumption. The heat conduction between aluminum honeycomb structural panels only relies on screws to produce a weak heat conduction path, and the heat conduction and temperature equalization capacity between panels is almost negligible. Using pre-buried heat pipes or high thermal conductivity graphite films to enhance the heat conduction and temperature equalization capacity of the structural panels themselves, and using mounted heat pipes or high thermal conductivity films between panels to enhance thermal conductivity, requires a certain amount of satellite space, which brings certain design constraints to the satellite's high inheritance and automated assembly.

[0005] Description of existing technical problems and defects: The thermal design of conventional spacecraft only focuses on practicality, using later added thermal control multi-layer insulation components and heat pipes as the main means of insulation and heat conduction, without fully considering the needs of mass automated production of commercial satellites and the concept of high-inheritance integrated design. This conventional thermal design method has increased the thermal control implementation cycle during the spacecraft assembly phase. The soft multi-layer insulation components rely more on manual assembly. At the same time, the multi-layer external spacecraft is prone to hooking risks, which is not conducive to batch launch; the non-integrated integrated design leads to an increase in coordination interfaces, which increases the complexity of the project. At this time, a new thermal control design method is needed to meet the normal temperature control needs of the spacecraft and reduce the complexity of the spacecraft assembly operation. Summary of the invention

[0006] The present invention provides a thermal control design method for an open-cabin spacecraft, comprising the following steps:

[0007] Step 1: The main structure of the spacecraft is integrally formed by aluminum alloy machining process;

[0008] Step 2: Spray or mount thermal control coatings with low solar absorption ratio and high infrared emissivity on all surfaces of the main structure.

[0009] As a further improvement of the present invention, in step 1, the main structure of the spacecraft is designed as a five-sided frame structure, which is integrally formed by aluminum alloy machining process.

[0010] As a further improvement of the present invention, a multi-layer thermal insulation assembly is used to construct a multi-layer flexible cabin panel.

[0011] As a further improvement of the present invention, the multi-layer flexible cabin panel is composed of a layer of double-sided aluminum-coated polyester film and a layer of polyester mesh stacked to form a unit, the double-sided aluminum-coated polyester film and the polyester mesh are alternately stacked multiple times, and then a layer of double-sided aluminum-coated polyester film is stacked on the top, and finally a layer of single-sided aluminum-coated polyimide film is covered on the side of the multi-layer insulation assembly facing the outside of the satellite cabin as a mask.

[0012] As a further improvement of the present invention, the multi-layer flexible cabin panel is composed of a unit consisting of a layer of 6μm thick double-sided aluminum-coated polyester film and a layer of polyester mesh. The double-sided aluminum-coated polyester film and the polyester mesh are alternately laminated 15 times, and then a layer of double-sided aluminum-coated polyester film is laminated on the top, and finally a layer of single-sided aluminum-coated polyimide film is covered on the side of the multi-layer insulation assembly facing the outside of the satellite cabin as a mask.

[0013] As a further improvement of the present invention, the thermal conductivity of the main structure is greater than 121 W / m / K.

[0014] As a further improvement of the present invention, in step 2, the equivalent thermal parameters of the spacecraft structure surface are controlled by designing the mounting ratio and area of ​​the F46 silver-plated secondary surface mirror on the industrial conductive oxidation surface.

[0015] As a further improvement of the present invention, the thermal control coating surface is industrially anodized and the thermal control coating is used to calculate the equivalent infrared emissivity based on the area ratio.

[0016] The beneficial effects of the present invention are as follows: the present invention achieves isolation of the direct solar heat flow of the maximum heat flow in the orbital heat flow by designing a stronger multi-layer flexible cabin panel on the side of the spacecraft with larger external heat flow, so that the influence of the external heat flow on the temperature of the spacecraft is greatly reduced; without the need to add an external heat conduction device, the high thermal conductivity of the integrated structure itself is utilized to effectively avoid the excessive temperature gradient between the side with larger external heat flow and the side with smaller external heat flow, and the use of heat pipes and external graphite film can be eliminated; by mounting a thermal control film or spraying a thermal control paint on the traditional industrial conductive anodized coating, the solar absorption ratio and infrared emissivity of the five outer walls of the main structure of the entire satellite are adjusted, so that the temperature of the spacecraft equipment meets the corresponding temperature index requirements; reducing one side of the metal structure cabin panel, and constructing a multi-layer flexible cabin panel using a multi-layer thermal insulation component, which effectively reduces the weight of the spacecraft; avoiding the external cabin panel covering the multi-layer thermal insulation component, so that each structural panel is more convenient for automated assembly, improves production efficiency, and shortens the development cycle. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 It is the thermal control design state diagram of the spacecraft;

[0018] Figure 2 It is the design schematic diagram of temperature control. DETAILED DESCRIPTION

[0019] The present invention discloses a thermal control design method for an open-cabin spacecraft, comprising the following steps:

[0020] Step 1: The main structure of the spacecraft is integrally formed by aluminum alloy machining process, and the interface heat transfer resistance between the cabin panels is eliminated by integral processing, and the high thermal conductivity of aluminum alloy itself is used to achieve good overall temperature uniformity of the main structure;

[0021] Step 2: By spraying or attaching thermal control coatings with low solar absorption ratio and high infrared emissivity on each surface of the main structure, the surface thermal parameters (solar absorption ratio, infrared emissivity) of the main structure after industrial conductive oxidation treatment are changed, thereby controlling the combined temperature influence of the outer heat flow of the orbit and the extremely cold background of deep space on the celestial body, so that the overall temperature of the spacecraft is at a good temperature level.

[0022] As a preferred embodiment of the present invention, the main structure of the spacecraft is designed as a five-sided frame structure, which is integrally formed by aluminum alloy machining process. The thermal conductivity of this structure is greater than 121W / m / K, realizing the uniform temperature design of the entire spacecraft; the side of the spacecraft that is directly exposed to the sun for a long time (the side with the largest external heat flow) is the open cabin side, and a 15-unit multi-layer insulation component is used to construct a multi-layer flexible cabin panel (a layer of 6μm thick double-sided aluminum-plated polyester film and a layer of polyester mesh are laminated to form a unit, the polyester film and the polyester mesh are alternately laminated 15 times, and then a layer of double-layered polyester mesh is laminated on the top). Finally, a single-sided aluminum-coated polyimide film is covered on the side of the multi-layer insulation component facing the outside of the satellite cabin as a mask), so that the spacecraft is closed as a whole to prevent the internal instruments and equipment from being directly exposed to the sun; by designing the mounting ratio and area of ​​the F46 silver-coated secondary surface mirror on the industrial conductive oxide surface, the equivalent thermal parameters of the spacecraft structure surface (equivalent thermal control coating) are controlled, and then the overall spacecraft's ability to absorb external heat flux and exchange heat with the extremely cold background of deep space is controlled, so that the overall temperature level of the spacecraft is at a good temperature level.

[0023] Implementation principle: The overall temperature of the spacecraft is uniform, which is mainly achieved through the heat conduction capacity of the structure itself. This part is a machine-integrated structure, and its heat conduction and heat dissipation capacity is:

[0024]

[0025] Where: Q is the thermal conductivity and heat dissipation capacity of the integrated structure, unit: W;

[0026] K is the thermal conductivity of the integrated structure, in W / ㎡ / ℃;

[0027] L is the distance of the heat conduction path, in m;

[0028] S is the cross-sectional area of ​​the heat conduction path of the integrated structure, unit: m2;

[0029] ΔT is the temperature difference between the heat source position and the cold end of the integrated structure, in °C.

[0030] The temperature level of the spacecraft as a whole is balanced by controlling its own heat generation, the total amount of heat flux absorbed outside the orbit, and the radiation heat to the cold background of deep space. The energy balance formula is as follows:

[0031] Q 外 +Q 内 =Q 散

[0032] Where Q 外 is the extraorbital heat flux absorbed by the spacecraft, Q 内 is the heat generated by the spacecraft’s internal equipment, Q 散 is the amount of heat dissipated by the spacecraft to the cold background of deep space, and the unit is W.

[0033] Q 外 It includes the ability to absorb direct sunlight, the energy of the planet's reflection of sunlight, and the infrared radiation energy of the planet itself to the spacecraft.

[0034] Q 外 =q 直 ×α s ×S 直 +q 反 ×α s ×S 反 +q 红 ×ε×S 红

[0035] Where: q 直 is the heat flux density of solar radiation to the spacecraft, q 反 is the heat flux density of the solar radiation energy reflected by the planet to the spacecraft, q 红 is the heat flux density of infrared energy emitted by the planet itself reaching the spacecraft, in W / ㎡;

[0036] α s is the solar absorption ratio, the thermal parameters of the spacecraft surface under direct sunlight and planetary reflection, and the equivalent infrared emissivity of the thermal control coating surface is calculated based on the area ratio of industrial anodizing and thermal control coating;

[0037] ε is the infrared emissivity of the outer surface of the spacecraft (the thermal control coating surface is calculated based on the equivalent infrared emissivity of the industrial anodized and thermal control coating surfaces according to the area ratio), and the multi-layer flexible cabin panel uses the overall equivalent infrared emissivity;

[0038] S is the area of ​​each face.

[0039] Q 散 This is the amount of heat radiated by each surface of the spacecraft to the cold background of deep space. If the surface infrared emissivity of each surface is different, it is necessary to calculate each surface separately and then sum them up.

[0040]

[0041] Where: ε is the infrared emissivity of each surface of the spacecraft;

[0042] σ is the radiation constant of a black body (Stefan-Boltzmann constant);

[0043] S is the radiation surface area of ​​each surface of the spacecraft;

[0044] T is the absolute temperature of the spacecraft surface, in K.

[0045] Through analysis, the present invention adopts a mechanically integrated open cabin main body structure, and utilizes its dual functions of high thermal conductivity and load-bearing to achieve the uniform temperature design of the spacecraft as a whole. By reasonably designing the equivalent solar absorption ratio and infrared emissivity of the spacecraft surface and using multi-layer flexible cabin panels to isolate the direct solar heat flux, the purpose of temperature control of the spacecraft as a whole can be achieved.

[0046] In summary, the present invention utilizes the high thermal conductivity and high load-bearing characteristics of the metal main structure to construct a satellite thermal bus, and then combines low-infrared emissivity and high-infrared emissivity thermal control coatings with low-cost industrial conductive oxide coatings. By controlling the area of ​​the thermal control coating, the expected equivalent solar absorption ratio and infrared emissivity are achieved. The open cabin side of the spacecraft is covered with a whole multi-layer thermal insulation component to construct a flexible closed cabin, so that the equipment inside and outside the spacecraft is within a good temperature range, avoiding the complicated installation of the spacecraft multi-layer thermal insulation component, thereby better adapting to the short-cycle and mass production of spacecraft.

[0047] The above contents are further detailed descriptions of the present invention in combination with specific preferred embodiments, and it cannot be determined that the specific implementation of the present invention is limited to these descriptions. For ordinary technicians in the technical field to which the present invention belongs, several simple deductions or substitutions can be made without departing from the concept of the present invention, which should be regarded as falling within the protection scope of the present invention.

Claims

1. A thermal control design method for an open-cabin spacecraft, characterized in that: The steps include: Step 1: The main structure of the spacecraft is integrally formed by aluminum alloy machining process; Step 2: Spray or mount thermal control coatings with low solar absorption ratio and high infrared emissivity on all surfaces of the main structure.

2. The thermal control design method for an open-cabin spacecraft according to claim 1, characterized in that: In step 1, the main structure of the spacecraft is designed as a five-sided frame structure and is integrally formed by aluminum alloy machining process.

3. The thermal control design method for an open-cabin spacecraft according to claim 1, characterized in that: Multi-layered thermal insulation components are used to construct multi-layered flexible cabin panels.

4. The thermal control design method for an open-cabin spacecraft according to claim 3, characterized in that: The multi-layer flexible cabin panel is composed of a layer of double-sided aluminized polyester film and a layer of polyester mesh. The double-sided aluminized polyester film and the polyester mesh are alternately laminated multiple times, and then a layer of double-sided aluminized polyester film is laminated on the top. Finally, a single-sided aluminized polyimide film is covered on the side of the multi-layer insulation component facing the outside of the star cabin as a mask.

5. The thermal control design method for an open-cabin spacecraft according to claim 4, characterized in that: The multi-layer flexible cabin panel is composed of a unit consisting of a 6μm thick double-sided aluminum-coated polyester film and a layer of polyester mesh. The double-sided aluminum-coated polyester film and the polyester mesh are alternately laminated 15 times, and then a double-sided aluminum-coated polyester film is laminated on the top. Finally, a single-sided aluminum-coated polyimide film is covered on the side of the multi-layer insulation component facing the outside of the star cabin as a mask.

6. The spacecraft thermal control design method according to claim 1, characterized in that: The thermal conductivity of the main structure is greater than 121 W / m / K.

7. The thermal control design method for an open-cabin spacecraft according to claim 1, characterized in that: In step 2, the equivalent thermal parameters of the spacecraft structure surface are controlled by designing the mounting ratio and area of ​​the F46 silver-plated secondary surface mirror on the industrial conductive oxide surface.

8. The thermal control design method for an open-cabin spacecraft according to claim 1, characterized in that: The equivalent infrared emissivity of the thermal control coating surface is calculated based on the area ratio of industrial anodizing and thermal control coating.

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