Spacecraft thermal control system for lunar environment

Abstract

The invention relates to the technical field of spaceflight, in particular to a spacecraft thermal control system for the lunar environment, a spacecraft comprises a cabin and a single machine, the single machine is arranged in the cabin, and the spacecraft thermal control system comprises a heat source, a heat transmission system, a heat insulation system and a heat dissipation system; the heat transmission component is arranged in the spacecraft, makes contact with the heat source and the single machine and is used for transmitting heat. According to the spacecraft thermal control system used for the lunar environment, heat of the heat source can be guided into the cabin through the heat transmission system with the high heat conductivity coefficient and is matched with the heat insulation system, heat energy needed for heat preservation in the moon night period is provided, when the heat of a single machine is high, the heat can be evenly distributed through the heat transmission system, and the heat preservation efficiency is improved. Therefore, the stability of the internal environment of the spacecraft is ensured. Therefore, according to the thermal control system scheme, the thermal control design principle that passive thermal control serves as a main mode and active thermal control serves as an auxiliary mode is adopted, and the spacecraft can meet the index requirement of the working temperature of the spacecraft in the complex lunar surface environment.

Description

Technical Field

[0001] The present application relates to the field of aerospace technology, and in particular to a spacecraft thermal control system for a lunar environment. Background Art

[0002] At present, with the continuous advancement of aerospace technology, especially in the fields of manned spaceflight, deep space exploration, lunar and Mars missions, temperature equalization technology has gradually received attention. Spacecraft face extreme temperature changes in space, and uniform temperature control is crucial to protecting the normal operation of spacecraft and their internal equipment. Summary of the invention

[0003] The purpose of this application is to provide a spacecraft thermal control system for the lunar environment, which to a certain extent solves the technical problem in the prior art that spacecraft face extreme temperature changes in space, and uniform temperature control is crucial to protecting the normal operation of the spacecraft and its internal equipment.

[0004] The present application provides a spacecraft thermal control system for a lunar environment, which is applied to a spacecraft, and the spacecraft includes a cabin and a single machine, and the single machine is arranged in the cabin. The spacecraft thermal control system for a lunar environment includes: a heat transfer component and a heat source; wherein the heat transfer component is arranged in the cabin, and the heat transfer component is respectively in contact with the heat source and the single machine for transferring heat.

[0005] In the above technical solution, further, the single machine is fixed on the deck of the cabin, and at least a part of the structure of the heat transmission component is embedded in the deck of the cabin and is arranged corresponding to the single machine.

[0006] In any of the above technical solutions, further, the area of ​​the cabin plate where the single machine is installed is the single machine installation area, the cabin forms an open area connected to the outside, and the single machine installation area is arranged corresponding to the open area, the single machine is arranged away from the open area side, and a thermal control coating is provided on the side of at least part of the structure of the single machine installation area close to the open area.

[0007] In any of the above technical solutions, further, the inner wall and the outer wall of the cabin panel of the cabin are covered with multi-layer insulation components, and the multi-layer insulation components avoid the thermal control coating.

[0008] In any of the above technical solutions, further, the area of ​​the cabin panel where the single machine is installed is a single machine installation area, and a thermal control layer is provided on one side of the single machine installation area close to the single machine to enhance radiation heat exchange in the cabin.

[0009] In any of the above technical solutions, further, the thermal control layer is a thermal control coating with high infrared emissivity.

[0010] In any of the above technical solutions, further, the area of the cabin board where the single machine is installed is the single machine installation area, and the single machine installation area includes a first aluminum skin, a second aluminum skin, and a honeycomb panel core; wherein, the first aluminum skin and the second aluminum skin are respectively disposed on opposite sides of the honeycomb panel core, and the heat transfer component is embedded in the honeycomb panel core.

[0011] In any of the above technical solutions, further, the spacecraft thermal control system for the lunar environment further includes a temperature sensor, and the cabin board where the single machine is installed and / or the single machine is equipped with the temperature sensor.

[0012] In any of the above technical solutions, further, a heat conduction layer is provided between the cabin board and the single machine.

[0013] In any of the above technical solutions, further, the heat conduction layer is a heat conductive silica gel layer or a heat conductive silicone grease layer.

[0014] In any of the above technical solutions, further, the heat source is disposed outside the cabin, and the structure of the heat transfer component extending outside the cabin is in contact with the heat source.

[0015] In any of the above technical solutions, further, the spacecraft thermal control system for the lunar environment further includes an electric heater, and the cabin board of the cabin and / or the single machine is equipped with the electric heater.

[0016] In any of the above technical solutions, further, the spacecraft further includes a solar wing, the solar wing is disposed outside the cabin, and except for the side where the battery cells are installed, the solar wing is covered with a multi-layer heat insulation component.

[0017] In any of the above technical solutions, further, the heat transfer component is pasted on the surface of the cabin board of the cabin.

[0018] In any of the above technical solutions, further, the heat transfer component is a flat heat pipe.

[0019] In any of the above technical solutions, further, the heat source is an isotope heat source or an electric heater.

[0020] Compared with the prior art, the beneficial effects of the present application are:

[0021] For the spacecraft thermal control system used in the lunar environment provided by this application, the heat of the heat source can be introduced into the cabin through the heat transfer component with high thermal conductivity to provide the thermal energy required for heat preservation during the lunar night. Moreover, when the heat of a single unit is relatively high, the heat can be evenly distributed through the heat transfer component, thus ensuring the stability of the internal environment of the spacecraft. It can be seen that the above thermal control system solution adopts the thermal control design principle of mainly passive thermal control and supplemented by active thermal control, which can enable the spacecraft to meet the index requirements of its working temperature in the complex lunar environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the specific embodiments of this application or the technical solutions in the prior art, the following will briefly introduce the drawings required for the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of this application. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0023] Figure 1 It is a schematic structural diagram of the cabin of the spacecraft provided by the embodiment of this application;

[0024] Figure 2 It is another schematic structural diagram of the cabin of the spacecraft provided by the embodiment of this application;

[0025] Figure 3 It is an assembly drawing of a single unit, a heat transfer component, and a cabin panel provided by the embodiment of this application;

[0026] Figure 4 It is another assembly drawing of a single unit, a heat transfer component, and a cabin panel provided by the embodiment of this application.

[0027] REFERENCE NUMERALS:

[0028] 1 - Cabin, 11 - Cabin panel, 111 - Single unit installation area, 2 - Single unit, 21 - Sensitive single unit, 3 - Heat transfer component, 4 - Heat source, 5 - Thermal control coating, 6 - Multi-layer insulation component, 7 - Thermal control layer. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0029] The following will clearly and completely describe the technical solutions of this application in conjunction with the drawings. Obviously, the described embodiments are some, but not all, of the embodiments of this application.

[0030] Generally, the components of the embodiments of this application described and shown in the drawings here can be arranged and designed in various different configurations. Therefore, the following detailed description of the embodiments of this application provided in the drawings is not intended to limit the scope of this application that is required to be protected, but only represents the selected embodiments of this application.

[0031] Based on the embodiments in this application, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the scope of protection of this application.

[0032] In the description of this application, it should be noted that the orientation or positional relationship indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing this application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this application. In addition, the terms "first", "second", "third" are only used for descriptive purposes and cannot be construed as indicating or implying relative importance.

[0033] In the description of this application, it should be noted that unless otherwise clearly specified and limited, the terms "installed", "connected", "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in this application can be understood according to specific situations.

[0034] The following refers to Figures 1 to 4 Describe a spacecraft thermal control system for lunar environment according to some embodiments of this application.

[0035] Refer to Figures 1 to 4 As shown, the embodiments of this application provide a spacecraft thermal control system for lunar environment, which is applied to a spacecraft, and the spacecraft includes a cabin 1 and a single machine 2. The single machine 2 is arranged in the cabin 1. The spacecraft thermal control system for lunar environment includes: a heat transfer component 3 and a heat source 4; wherein, the heat transfer component 3 is arranged in the cabin 1, and the heat transfer component 3 is in contact with the heat source 4 and the single machine 2 respectively for transferring heat. Among them, the heat source 4 mainly maintains the temperature level of the lunar night equipment, that is, the spacecraft under the non-operating condition.

[0036] According to the above-described structure, it can be seen that the heat of the heat source 4 can be introduced into the cabin through the heat transfer component 3 with high thermal conductivity to provide the thermal energy required for heat preservation during the lunar night. Moreover, when the heat of the single machine 2 is relatively high, the heat can be evenly distributed through the heat transfer component 3, thereby ensuring the stability of the internal environment of the spacecraft.

[0037] It can be seen that the above thermal control system solution adopts the thermal control design principle of mainly passive thermal control and supplemented by active thermal control, which can enable the spacecraft to meet the index requirements of its working temperature in the complex lunar environment.

[0038] Furthermore, preferably, the heat transfer component 3 can be a flat heat pipe, which has the ability to operate against gravity and can adapt to climbing operations in the lunar g / 6 gravity environment. Of course, the type of the heat transfer component 3 is not limited to the above, and other types of heat conduction structures can also be used.

[0039] Furthermore, preferably, a part of the structure of the heat transfer component 3 extends to the outside of the chamber 1. During lunar day, the heat transfer component 3 is used to evenly distribute the uneven heat dissipation in the area of the single machine 2, so as to make the heat distribution uniform, ensure the stability of the internal environment of the spacecraft, and at the same time transfer the heat dissipation of the single machine 2 to the heat dissipation surface outside the chamber 1 for dissipation. Of course, it is not limited to this, and the whole of the heat transfer component 3 can also be only arranged inside the chamber 1 and not exposed to the outside of the chamber 1 at all.

[0040] Furthermore, preferably, the heat source 4 is arranged outside the chamber 1, which can avoid occupying the space inside the chamber 1 and can be in contact with the heat transfer component 3 extending outside the chamber 1, so as to provide heat to the inside of the chamber 1. This will also be taken as an example for explanation later. Of course, the heat source 4 can be installed inside the chamber 1. In this case, the heat transfer component 3 can be completely located inside the chamber 1, or part of it can be located outside the chamber 1 and part of it can be located inside the chamber 1, which is specifically selected according to actual needs.

[0041] Furthermore, preferably, the heat source 4 is an isotope heat source. Of course, it is not limited to this, and it can also be selected according to actual needs.

[0042] Furthermore, preferably, the number of the single machines 2 is multiple. Of course, it is not limited to this, and the number of the single machines 2 can also be one, which is specifically selected according to actual needs.

[0043] In this embodiment, preferably, as Figure 1 and Figure 2 shown, the single machine 2 is fixed on the cabin board 11 of the chamber 1, and at least part of the structure of the heat transfer component 3 is embedded in the cabin board 11 of the chamber 1 and is arranged corresponding to the single machine 2.

[0044] According to the structure described above, using the inner wall of the chamber 1 as the structure to support the single machine 2 can avoid introducing additional support structures, which is helpful for lightweight and miniaturized design, and at the same time helps to reduce costs.

[0045] Furthermore, preferably, a part of the structure of the heat transfer component 3 is embedded in the cabin board 11 of the chamber 1, and the other part of its structure extends to the outside of the chamber 1 and is in contact with the heat source 4 outside the chamber 1. And preferably, the heat transfer component 3 can be connected to the heat source 4 together. Of course, it is not limited to this.

[0046] In this embodiment, preferably, as Figure 3As shown, the area of ​​the cabin panel 11 where the single machine 2 is installed is the single machine installation area 111, the cabin 1 forms an open area connected to the outside, and the single machine installation area 111 is arranged corresponding to the open area, the single machine 2 is arranged away from the open area side, and the single machine installation area 111 has at least part of the structure thereof and is close to the open area. A thermal control coating 5 is provided on the skyward surface of at least part of the structure thereof, and the thermal control coating 5 has the characteristics of low solar absorption ratio and high infrared emissivity. In addition, it should be noted that structures such as solar wings can also close this open area. This is a prior art and will not be described in detail herein.

[0047] According to the structure described above, a thermal control coating 5 is provided in the single-machine installation area 111 near the opening, that is, away from the single-machine 2 side. The thermal control coating 5 has high reflectivity and low absorption to the solar spectrum and high absorption to the infrared spectrum. The heat consumption of the single-machine 2 can be dissipated to the ceiling through the cabin panel 11, thereby ensuring the normal operation of the single-machine 2 and helping to extend the service life of the single-machine 2.

[0048] Further, preferably, the thermal control coating 5 can be a secondary surface mirror. Of course, it is not limited thereto and can be selected according to actual needs.

[0049] In this embodiment, preferably, Figure 3 As shown, the inner wall and the outer wall of the cabin panel 11 of the cabin 1 are both covered with a multi-layer heat insulation component 6, and the multi-layer heat insulation component 6 avoids the thermal control coating 5.

[0050] According to the structure described above, the inner and outer walls of the cabin plate 11 of the entire cabin 1 are covered with multi-layer insulation components 6 to reduce heat leakage, and the multi-layer insulation components 6 avoid the thermal control coating 5 to avoid interference, and the two do not affect each other.

[0051] Further, preferably, the multi-layer thermal insulation component 6 may be a 20-unit multi-layer thermal insulation component 6. Of course, it is not limited thereto and may be selected according to actual needs.

[0052] In this embodiment, preferably, Figure 3 and Figure 4 As shown, when the standard operating temperature of the single machine 2 is above the preset temperature, the single machine 2 is a sensitive single machine 21, and the single machine installation area 111 corresponding to the sensitive single machine 21 is wrapped with a multi-layer thermal insulation component 6. That is to say, the thermal control coating 5 is not arranged at the sensitive single machine 21, but a corresponding multi-layer thermal insulation component 6 is arranged, or most of the sensitive single machine 21 is equipped with a multi-layer thermal insulation component 6, and only a small part is equipped with a thermal control coating 5, which is selected according to actual needs.

[0053] According to the structure described above, it can be seen that the single-machine installation area 111 corresponding to some single-machines 2 with higher standard operating temperature requirements, that is, the inner wall and outer wall of the corresponding cabin plate 11 are wrapped with multiple layers of thermal insulation components 6 to play a role in heat preservation and ensure the operating temperature requirements of such single-machines 2.

[0054] Further, preferably, the preset temperature may be 0°C, but is not limited thereto and may be designed according to actual needs.

[0055] In this embodiment, preferably, Figure 4 As shown, the structure of the heat transmission component 3 exposed to the cabin 1 is covered with a multi-layer heat insulation component 6 to reduce heat leakage.

[0056] Further, preferably, the multi-layer thermal insulation component 6 may be a 20-unit multi-layer thermal insulation component 6. Of course, it is not limited thereto and may be selected according to actual needs.

[0057] In this embodiment, preferably, Figure 4 As shown, the area of ​​the cabin panel 11 where the single machine 2 is installed is the single machine installation area 111, and a thermal control layer 7 is provided on one side of the single machine installation area 111 close to the single machine 2 to enhance the radiation heat exchange in the cabin.

[0058] According to the structure described above, the thermal control layer 7 is used to enhance the radiation heat exchange in the cabin and to even out the overall temperature. Of course, the thermal control layer 7 may not be provided on the side of the single-machine installation area 111 close to the single-machine 2, depending on actual needs.

[0059] Further, preferably, the thermal control layer 7 is formed by spraying thermal control black paint on the side of the single-machine installation area 111 close to the single machine 2, that is, the thermal control layer 7 is a thermal control coating with high infrared emissivity. Of course, the material of the thermal control layer 7 is not limited to this, and other materials can also be selected.

[0060] In this embodiment, preferably, the area of ​​the cabin panel 11 where the single machine 2 is installed is the single machine installation area 111, and the single machine installation area 111 includes a first aluminum skin, a second aluminum skin and a honeycomb panel core; wherein the first aluminum skin and the second aluminum skin are respectively arranged on opposite sides of the honeycomb panel core, and the heat transmission component 3 is pre-embedded in the honeycomb panel core.

[0061] According to the structure described above, the single-machine installation area 111, that is, at least a part of the cabin 11, adopts a honeycomb structure as the core, which not only meets the strength requirements but also contributes to lightweight design. In addition, the heat transmission component 3 is pre-buried in the single-machine installation area 111, that is, at least a part of the cabin 11, which not only meets the heat conduction requirements but also serves as a supporting structure for the single machine 2.

[0062] Further, preferably, the honeycomb panel core is an aluminum honeycomb panel, and the heat transfer component 3, such as a flat heat pipe, is disposed inside the aluminum honeycomb panel. And preferably, the above components can be formed into an integral structure by hot pressing, so as to embed the heat transfer component 3, such as a flat heat pipe, inside the cabin panel 11.

[0063] It should be noted that: the heat transfer component 3 and the single machine installation area 111, that is, at least part of the cabin panel 11, are not limited to the above-mentioned embedded structure. The heat transfer component 3 can also be directly installed on the surface of the single machine installation area 111, such as by pasting, or an installation groove can be opened on the single machine installation area 111, and the heat transfer component 3 can be installed in the installation groove, etc., which is specifically selected according to actual needs.

[0064] In addition, it should also be noted that: the cabin panels 11 of the entire cabin 1, that is, all the cabin panels 11, can be designed into the above-mentioned honeycomb panel structure with skins, or only part of the cabin panels 11 in the area where the single machine 2 is installed can be designed into the honeycomb panel structure with skins, which is specifically selected according to actual needs.

[0065] In this embodiment, preferably, the spacecraft thermal control system for the lunar environment further includes a temperature sensor (not shown in the figure), and the cabin panel 11 where the single machine 2 is installed and / or the single machine 2 is provided with a temperature sensor. That is to say, there are three cases: the cabin panel 11 where the single machine 2 is installed is provided with a temperature sensor, only the single machine 2 is provided with a temperature sensor, or both the cabin panel 11 where the single machine 2 is installed and the single machine 2 are provided with temperature sensors for temperature measurement and control of the equipment.

[0066] Further, preferably, the temperature sensor is pasted on the surface of the single machine installation area 111 or the surface of the single machine 2, or multiple temperature sensors are set at the same time and are respectively installed on the surface of the single machine installation area 111 and the surface of the single machine 2.

[0067] In this embodiment, preferably, a heat conduction layer (not shown in the figure) is provided between the cabin panel 11 and the single machine 2.

[0068] According to the structure described above, a heat conduction layer is provided between the cabin panel 11 and the single machine 2, which plays a role in strengthening heat transfer. Of course, this heat conduction layer can also not be provided, which is specifically selected according to actual needs.

[0069] Further, preferably, the heat conduction layer is a heat conduction silicone layer or a heat conduction silicone grease layer, both of which can play a role in strengthening heat transfer. Of course, the material of the heat conduction layer is not limited to this, and can also be selected according to actual needs.

[0070] In this embodiment, preferably, the spacecraft further includes a solar wing (not shown in the figure). The solar wing is disposed outside the cabin 1, and a multi-layer heat insulation component 6 is coated on all sides of the solar wing except the side where the battery cells are installed.

[0071] According to the structure described above, most of the outer surface of the solar panel is covered with a multi-layer thermal insulation component 6, thereby reducing heat leakage. Of course, it is not limited to this, and this multi-layer thermal insulation component 6 may not be provided.

[0072] Furthermore, preferably, this multi-layer thermal insulation component 6 can be a 20-unit multi-layer thermal insulation component 6. Of course, it is not limited to this, and it can also be selected according to actual needs.

[0073] In this embodiment, preferably, as Figure 1 shown, the number of heat transfer components 3 is two. Of course, it is not limited to this, and the number can also be one or more than two, such as three or four, etc., and is specifically arranged reasonably according to the layout of the single machine 2.

[0074] In this embodiment, preferably, the spacecraft thermal control system for the lunar environment further includes an electric heater (not shown in the figure). The cabin panel 11 of the cabin 1 and / or the single machine 2 are provided with electric heaters. That is to say, there are three cases: only the cabin panel 11 of the cabin 1 is provided with an electric heater, only the single machine 2 is provided with an electric heater, or both the cabin panel 11 of the cabin 1 and the single machine 2 are provided with electric heaters. The electric heater is used as a compensation heater to maintain the temperature level under the non-operating condition of the equipment, so as to improve the environmental adaptability of the thermal control system.

[0075] Furthermore, preferably, the electric heater can be pasted on the surface of the cabin panel 11 or the surface of the single machine 2.

[0076] In summary, for the spacecraft thermal control system for the lunar environment provided by this application, the overall thermal design passive thermal control measures mainly include heat insulation, heat conduction, radiation heat dissipation, etc., and the active thermal control measure is an electric heater. The working principle of the thermal control system is mainly to conduct the heat generated by each platform device, that is, the single machine 2, to the robot body, that is, the cabin 1 and the extended heat dissipation surface, through various heat transfer methods, and dissipate it into space through space radiation. At the same time, through certain heat insulation measures, it blocks the external heat fluxes such as direct sunlight and lunar infrared from entering the robot interior. When the equipment is at a low temperature, the active thermal control mainly relies on electric heating and isotope heat sources to ensure that each equipment operates within the required temperature range.

[0077] Finally, it should be noted that: the above embodiments are only used to illustrate the technical solutions of this application, and are not intended to limit them; although this application has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions recorded in the foregoing embodiments, or perform equivalent replacements on some or all of the technical features; and these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of this application.

Claims

1. A spacecraft thermal control system for a lunar environment, applied to a spacecraft, wherein the spacecraft comprises a cabin and a single machine, wherein the single machine is arranged in the cabin, and wherein: The spacecraft thermal control system for the lunar environment includes: a heat transmission component and a heat source; wherein the heat transmission component is arranged in the cabin, and the heat transmission component is respectively in contact with the heat source and the single machine for transferring heat.

2. The spacecraft thermal control system for lunar environment according to claim 1, characterized in that: The single unit is fixed on the deck of the cabin, and at least a part of the structure of the heat transmission component is pre-buried in the deck of the cabin and is arranged corresponding to the single unit.

3. The spacecraft thermal control system for lunar environment according to claim 2, characterized in that: The area of ​​the cabin plate where the single machine is installed is the single machine installation area, the cabin forms an open area connected to the outside, and the single machine installation area is arranged corresponding to the open area, the single machine is arranged away from the open area side, and a thermal control coating is provided on the side of at least part of the structure of the single machine installation area close to the open area.

4. The spacecraft thermal control system for lunar environment according to claim 3, characterized in that: The inner wall and the outer wall of the cabin panel are both covered with multi-layer heat insulation components, and the multi-layer heat insulation components avoid the thermal control coating.

5. The spacecraft thermal control system for lunar environment according to claim 2, characterized in that: The area of ​​the cabin panel where the single machine is installed is a single machine installation area, and a thermal control layer is provided on one side of the single machine installation area close to the single machine to enhance radiation heat exchange in the cabin.

6. The spacecraft thermal control system for lunar environment according to claim 5, characterized in that: The thermal control layer is a thermal control coating with high infrared emissivity.

7. The spacecraft thermal control system for lunar environment according to claim 2, characterized in that: The area of ​​the cabin panel where the single machine is installed is a single machine installation area, and the single machine installation area includes a first aluminum skin, a second aluminum skin and a honeycomb panel core; wherein the first aluminum skin and the second aluminum skin are respectively arranged on opposite sides of the honeycomb panel core, and the heat transmission component is pre-buried in the honeycomb panel core; and / or The spacecraft thermal control system for the lunar environment also includes a temperature sensor, and the cabin panel on which the single unit is installed and / or the single unit is equipped with the temperature sensor.

8. The spacecraft thermal control system for lunar environment according to claim 2, characterized in that: A heat-conducting layer is provided between the cabin plate and the single machine; The heat-conducting layer is a heat-conducting silica gel layer or a heat-conducting silicone grease layer.

9. The spacecraft thermal control system for lunar environment according to claim 2, characterized in that: A portion of the heat transfer component extends to the exterior of the chamber; The heat source is disposed outside the cabin, and the structure of the heat transmission component extending to the outside of the cabin is in contact with the heat source; The structure of the heat transmission component exposed to the cabin is covered with multiple layers of heat insulation components.

10. The spacecraft thermal control system for lunar environment according to any one of claims 1 to 9, characterized in that: The spacecraft thermal control system for the lunar environment further comprises an electric heater, and the cabin panel and / or the single unit are equipped with the electric heater; and / or The spacecraft further comprises a solar wing, which is arranged outside the cabin, and the solar wing is covered with a multi-layer heat insulation component except for one side where the battery sheet is installed; and / or The heat transmission component is attached to the surface of the cabin panel; and / or The heat transfer component is a flat heat pipe; and / or The heat source is an isotope heat source or an electric heater.

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