Spacecraft structure plate and spacecraft

Abstract

The invention relates to the technical field of spaceflight equipment, in particular to a spacecraft structure plate and a spacecraft, the spacecraft structure plate comprises a shell and a plurality of thermal control units, each thermal control unit comprises a containing part and a phase change material, a containing cavity is formed in each containing part, the phase change materials are located in the containing cavities, a cavity is formed in the shell, and the phase change materials are located in the cavity. The thermal control units are located in the cavity so as to fill the cavity, and the thermal control units are used for transferring heat. The invention aims to provide a spacecraft structure plate and a spacecraft in order to solve at least one technical problem related in the background technology.

Description

Technical Field

[0001] This application relates to the field of aerospace equipment technology, and more specifically, to a spacecraft structural plate and a spacecraft. Background Technology

[0002] To ensure the successful completion of flight missions, spacecraft need to be equipped with not only the payload system to perform the mission, but also several support systems that perform different functions, such as thermal control system, structural system, attitude control system, orbit control system and power system.

[0003] The thermal control system and the structural system are two crucial components of a spacecraft. The former ensures the spacecraft and its onboard instruments operate within ideal ambient temperatures through a series of measures, while the latter provides the basic configuration of the spacecraft, supporting and bearing the instruments, equipment, and means of implementing the thermal control system (the installation of thermal control products). Furthermore, the spacecraft structure provides heat conduction channels for the thermal control system equipment and radiating surfaces for heat dissipation into space. Although inextricably linked, these two systems are typically designed independently during spacecraft development.

[0004] Compared to the thermal control systems of ground products, the thermal control systems of spacecraft also need to cope with the complex and ever-changing external heat flow in space: when the external heat flow fluctuates, the temperature of the spacecraft's structural plates will change accordingly, which will lead to temperature fluctuations in the instruments and equipment inside the spacecraft.

[0005] For temperature-sensitive equipment such as batteries, rubidium atomic clocks, and optical devices, temperature stability is crucial for ensuring equipment lifespan and performance. Currently, spacecraft thermal control systems employ three main design methods to suppress temperature fluctuations: covering spacecraft structural panels or equipment with multi-layered thermal insulation components; installing thermal insulation pads between the equipment and the spacecraft structural panels; and using active electric heating to control the equipment temperature at its highest point of fluctuation. While the first two methods can reduce the impact of external heat flow on thermal interference and suppress temperature fluctuations, they also affect the equipment's normal heat dissipation. The third method, while reducing the amplitude of temperature fluctuations, results in additional energy waste. Therefore, with the rapid and diversified development of space exploration missions, the adaptability of thermal control systems to the space environment urgently needs to be strengthened. Summary of the Invention

[0006] The purpose of this application is to provide a spacecraft structural plate and a spacecraft in response to at least one of the technical problems involved in the background art.

[0007] To achieve the above objectives, this application adopts the following technical solution: One aspect of this application provides a spacecraft structural plate including a shell and a plurality of thermal control units, each thermal control unit including a housing and a phase change material, a housing cavity being formed within the housing, the phase change material being located within the housing, a chamber being formed within the shell, each of the thermal control units being located within the chamber to fill the chamber, and the thermal control units being used to transfer heat between each other.

[0008] Optionally, the receiving element is a spherical element.

[0009] The beneficial effects of this technical solution are as follows: after each thermal control unit is sent into the chamber using spherical components, it can automatically arrange itself according to the shape and size of the chamber under the action of gravity, forming a stable arrangement structure. The arrangement process of each thermal control unit requires little or almost no intervention from personnel or equipment, which improves the production efficiency of spacecraft structural plates. At the same time, the spherical components also have good structural stability and compressive strength, giving the spacecraft structural plates good structural strength.

[0010] Optionally, each pair of adjacent thermal control units is in direct contact with each other.

[0011] The beneficial effects of this technical solution are: it enables better heat transfer between each thermal control unit and generates mutual supporting forces between each thermal control unit, giving the spacecraft structural plate better rigidity and strength.

[0012] Optionally, vacuum gaps are formed between each of the thermal control units and between each of the thermal control units and the housing.

[0013] The beneficial effects of this technical solution are: it can give the spacecraft structural plate better heat insulation capabilities, and the spacecraft structural plate with a vacuum gap inside can be specifically placed in the spacecraft at locations where heat insulation is required.

[0014] Optionally, the spacecraft structural plate provided in this application further includes a heat dissipation connection layer disposed in the cavity, the heat dissipation connection layer filling the gaps between each of the thermal control units and between the thermal control units and the shell.

[0015] The beneficial effects of this technical solution are as follows: the heat dissipation connection layer filler has the function of connecting each thermal control unit and each thermal control unit to the shell as a whole, thereby improving the overall strength of the spacecraft structural plate. At the same time, the heat dissipation connection layer has a better heat transfer capacity than a vacuum, thereby enabling the spacecraft structural plate to have a better heat dissipation capacity. The spacecraft structural plate with the heat dissipation connection layer inside can be specifically placed in the spacecraft at locations with certain heat dissipation requirements.

[0016] Optionally, the outer wall of the accommodating member is polished.

[0017] The beneficial effect of this technical solution is that after the metal polishing process, the infrared hemispherical emissivity value of the outer surface of the container is lower, which means that it radiates and receives less heat, thereby reducing the mutual radiative heat transfer relationship between the containers.

[0018] Optionally, the receiving cavity is a spherical cavity.

[0019] The beneficial effects of this technical solution are: it enables the phase change material in the cavity to absorb heat uniformly from the surroundings and dissipate heat uniformly to the surroundings, thereby achieving a high efficiency in heat transfer. In addition, the spherical cavity has a good load-bearing capacity, effectively avoiding the stress concentration effect caused by the volume change of the phase change material in the cavity during phase change.

[0020] Optionally, the thermal control unit that contacts the housing is fixedly connected to the housing.

[0021] The beneficial effect of this technical solution is that it gives the spacecraft structural panels better integrity, and thus better structural strength.

[0022] Another aspect of this application provides a spacecraft including a heat-generating device and a spacecraft structural plate provided in this application, wherein the heat-generating device is mounted on one side of the spacecraft structural plate.

[0023] Optionally, the heating device is an optical remote sensing camera, and the phase change material is n-pentadecane or n-octadecane; Alternatively, the heating device may be a rubidium atomic clock, and the phase change material may be n-nonadecane or n-tetradecane; Alternatively, the heating device may be a storage battery, and the phase change material may be n-hexadecane.

[0024] The beneficial effect of this technical solution is that, based on the heating device and its ideal operating temperature, the thermal control unit uses a phase change material whose phase change point temperature is similar to or the same as the ideal operating temperature, thereby achieving an ideal heat dissipation effect.

[0025] The technical solution provided in this application can achieve at least one of the following beneficial effects: The spacecraft structural plate and spacecraft provided in this application have the functions of both a structural system and a thermal control system. They have strong adaptability to the space environment. Furthermore, while reducing the impact of external heat flow on the thermal interference of the equipment and suppressing the temperature fluctuation of the equipment, they also reduce the impact on the normal heat dissipation of the equipment.

[0026] The additional technical features and advantages of this application will become more apparent from the following description or from practical application. Attached Figure Description

[0027] To more clearly illustrate the technical solutions of the specific embodiments of this application, the accompanying drawings used in the description of the specific embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are some embodiments of this application. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0028] Figure 1 This is a partial structural schematic diagram of one embodiment of a spacecraft provided in this application, wherein arrow A indicates the flow direction of heat generated by the heat-generating device, and arrow B indicates the heat flow direction in the space environment; Figure 2 A schematic diagram of the front cross-sectional structure of one embodiment of the spacecraft structural plate provided in this application; Figure 3 A side sectional view of one embodiment of a spacecraft structural plate provided in this application; Figure 4 A top cross-sectional view of one embodiment of a spacecraft structural plate provided in this application; Figure 5 This is a partial structural schematic diagram of another embodiment of the spacecraft provided in this application, wherein arrow A indicates the flow direction of heat generated by the heat-generating device, and arrow B indicates the heat flow direction in the space environment; Figure 6 A front sectional view of another embodiment of the spacecraft structural plate provided in this application. Figure 7 A side sectional view of another embodiment of the spacecraft structural plate provided in this application. Figure 8 A top cross-sectional view of another embodiment of the spacecraft structural plate provided in this application. Figure 1 and Figure 5 The portion marked by numeral 03 in the attached figure represents a phase change material.

[0029] Figure label: 01. Heating device; 02. Housing; 03. Phase change materials; 04. Receptacle components; 05. Vacuum gap; 06. Thermal control unit; 07. Spacecraft structural panels. Detailed Implementation

[0030] The technical solutions of this application will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0031] In the description of this application, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0032] In the description of this application, it should be noted that, unless otherwise expressly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection between two components. Those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0033] like Figures 1 to 8 As shown, one aspect of this application provides a spacecraft structural plate 07, including a shell 02 and a plurality of thermal control units 06. Each thermal control unit 06 includes a receiving member 04 and a phase change material 03. A receiving cavity is formed in the receiving member 04, and the phase change material 03 is located in the receiving cavity. A chamber is formed in the shell 02, and each thermal control unit 06 is located in the chamber to fill the chamber. The thermal control units 06 are used to transfer heat between each other.

[0034] In this embodiment, each of the thermal control units 06 is located within the cavity to fill it, allowing gaps between the thermal control unit 06 and the housing 02, as well as between the thermal control units 06 themselves. In this embodiment, the housing 02 is preferably a metal skin. It is understood that the housing 02 and the thermal control units 06 are manufactured separately and then assembled into the spacecraft structural panel 07.

[0035] The spacecraft structural plate 07 provided in this application, when in use, has one side for mounting the heating device 01, and the other side facing the space environment. When the external heat flow to the space environment is large, the temperature of the phase change material 03 in the thermal control unit 06 located near the space environment rises to the phase change point temperature. During the phase change process, the phase change material 03 maintains a constant temperature in the form of latent heat of fusion. When the external heat flow to the space environment is small, the molten heat energy of the phase change material 03 in the thermal control unit 06 located near the space environment is dissipated into the space environment through radiation, and the phase change material 03 slowly cools to solidify. During the alternating melting and solidification process of the phase change material 03, the heat of fusion of the phase change material 03 can be effectively suppressed by the space environment. External heat flow fluctuations cause thermal interference to the heating device 01 located on the other side of the spacecraft structural plate 07. When the heating device 01 is an intermittently powered device or a device with frequently changing operating heat consumption, as the heating device 01 heats up and stops heating up, the phase change material 03 in the thermal control unit 06 corresponding to the heating device 01 melts and solidifies alternately. The latent heat of phase change of the phase change material 03 in the thermal control unit 06 can directly suppress large temperature fluctuations of the heating device 01. Furthermore, since each of the thermal control units 06 is located in the cavity to fill the cavity, each thermal control unit 06 and the shell 02 together bear the force as part of the structural system.

[0036] The spacecraft structural plate 07 provided in this application combines the functions of a structural system and a thermal control system. It has a strong adaptability to the space environment. Furthermore, while reducing the impact of external heat flow on the thermal interference of the equipment and suppressing the temperature fluctuation of the equipment, it also reduces the impact on the normal heat dissipation of the equipment.

[0037] Enhancing the adaptability of thermal control systems to the space environment is a crucial prerequisite for enabling spacecraft to autonomously plan for diverse missions. Furthermore, fully leveraging the close relationship between structural and thermal control systems to improve both systems is key to enhancing general-purpose spacecraft platforms. The spacecraft structural plate 07 provided in this application exhibits high adaptability to general-purpose spacecraft platforms, demonstrating strong temperature stability and passive thermal control capabilities against external heat flow fluctuations under different orbital conditions and variations in instrument and equipment heat dissipation in different application scenarios.

[0038] Optionally, the receiving component 04 is a spherical component. Because the spherical component allows the thermal control units 06 to automatically align themselves according to the shape and size of the chamber under gravity, forming a stable arrangement structure, the arrangement process of each thermal control unit 06 requires little or no intervention from personnel or equipment, thus improving the production efficiency of the spacecraft structural plate 07. Simultaneously, the spherical component also possesses better structural stability and compressive strength, giving the spacecraft structural plate 07 better structural strength. Of course, the receiving component 04 can also be a rugby ball-shaped component or other devices with curved surfaces; or, as... Figures 1 to 4As shown, the multiple receiving components 04 include spherical components and incomplete spherical components. The outer surface of the incomplete spherical component consists of an incomplete spherical portion and a planar portion. The incomplete spherical portion of the incomplete spherical component can occupy 1 / 10 to 9 / 10 of the outer surface of the spherical component. For example, the incomplete spherical portion of the incomplete spherical component can occupy 3 / 4, 1 / 2, or 1 / 3 of the outer surface of the spherical component. The incomplete spherical component is located close to the inner wall of the shell 02, and the planar portion of the incomplete spherical component is in contact with the inner wall of the shell 02. In the embodiments of this application, the number of receiving components 04 can be flexibly selected according to the shape and size of the cavity. In this way, it is not necessary to specially design and manufacture an internal filling structure of a corresponding size for a spacecraft structural plate 07 of a specific size. Instead, it is only necessary to produce a shell 02 of a corresponding size and then arrange a corresponding number of thermal control units 06 in the cavity, saving design and production costs and improving production efficiency. Figures 2 to 4 As shown, and Figures 6 to 8 As shown, at least two rows of thermal control units 06 should be provided in both the X direction (width direction of spacecraft structural plate 07) and the Y direction (length direction of spacecraft structural plate 07), and at least one layer of thermal control units 06 should be provided in the Z direction (rear direction of spacecraft structural plate 07).

[0039] Optionally, each pair of adjacent thermal control units 06 is in direct contact. This results in good heat transfer between the thermal control units 06 and generates mutual supporting forces between them, giving the spacecraft structural plate 07 good rigidity and strength. Of course, isolation components can also be provided between each pair of adjacent thermal control units 06 as needed.

[0040] Optionally, vacuum gaps 05 are formed between each of the thermal control units 06, and between each of the thermal control units 06 and the housing 02. The vacuum gaps 05 can also be formed by the thermal control units 06 and the housing 02. That is, since the receiving member 04 is a spherical member, after direct contact between two adjacent thermal control units 06, gaps will inevitably be generated between each of the thermal control units 06, and between each of the thermal control units 06 and the housing 02. In this embodiment, these gaps are set as vacuum gaps 05. This allows the spacecraft structural plate 07 to have better thermal insulation capabilities, and the spacecraft structural plate 07 with the internal vacuum gaps 05 can be specifically placed in locations on the spacecraft where thermal insulation is required.

[0041] Optionally, the spacecraft structural plate 07 provided in this embodiment further includes a heat dissipation connecting layer disposed within the cavity. This heat dissipation connecting layer fills the gaps between each of the thermal control units 06 and between the thermal control units 06 and the shell 02. The gaps can also be formed by the thermal control units 06 and the shell 02. The heat dissipation connecting layer serves to connect each thermal control unit 06 and its connection to the shell 02, thereby improving the overall strength of the spacecraft structural plate 07. Simultaneously, the heat dissipation connecting layer has good heat transfer capabilities relative to a vacuum, thus enabling the spacecraft structural plate 07 to have good heat dissipation capabilities. This spacecraft structural plate 07, with its internal heat dissipation connecting layer, can be strategically placed in locations on the spacecraft where heat dissipation is required. Of course, the gaps can be partially filled by the heat dissipation connecting layer, with the remaining portion being a vacuum gap 05.

[0042] Optionally, the outer wall of the receiving element 04 is polished. After metal polishing, the infrared hemispherical emissivity value of the outer surface of the receiving element 04 is lower, meaning it radiates and receives less heat, thereby reducing the mutual radiative heat transfer between the receiving elements 04. After metal polishing, the infrared hemispherical emissivity value of the outer surface of the receiving element 04 is typically below 0.1.

[0043] Optionally, the receiving cavity is a spherical cavity. This allows the phase change material 03 inside the receiving cavity to absorb heat uniformly from its surroundings and dissipate heat uniformly to its surroundings, thereby achieving high heat transfer efficiency. Furthermore, the spherical cavity has good load-bearing capacity, effectively avoiding stress concentration caused by volume changes during phase change of the phase change material inside the cavity. Of course, the receiving cavity can also be rugby ball-shaped or cubic, etc.

[0044] Optionally, the thermal control unit 06, which contacts the housing 02, is fixedly connected to the housing 02. In this embodiment, adhesive is filled between the thermal control unit 06 and the housing 02, and the composite is performed under high temperature and pressure. This gives the spacecraft structural plate 07 better integrity and thus better structural strength. Of course, the thermal control units 06 that contact the housing 02 can also only make pressure contact with the housing 02 to define their positions. This pressure can be generated by the interaction between the thermal control units 06 after they fill the chamber, and by the interaction between the thermal control units 06 and the housing 02.

[0045] Another aspect of this application provides a spacecraft, including a heat-generating device 01 and a spacecraft structural plate 07 provided in the embodiments of this application, wherein the heat-generating device 01 is mounted on one side of the spacecraft structural plate 07.

[0046] The spacecraft provided in this application utilizes the spacecraft structural plate 07 provided in this application. During use, one side is used to install the heating device 01, and the other side faces the space environment. When the external heat flow to the space environment is large, the temperature of the phase change material 03 in the thermal control unit 06 located near the space environment rises to the phase change point temperature. During the phase change process, the phase change material 03 maintains a constant temperature in the form of latent heat of fusion. When the external heat flow to the space environment is small, the molten heat energy of the phase change material 03 in the thermal control unit 06 located near the space environment is dissipated into the space environment through radiation. The phase change material 03 slowly cools to solidification. During the alternating melting and solidification process of the phase change material 03, [there is a possibility of...]. The system effectively suppresses the thermal interference caused by external heat flow fluctuations in the space environment to the heating device 01 located on the other side of the spacecraft structural plate 07. When the heating device 01 is an intermittently powered device or a device with frequently changing operating heat consumption, as the heating device 01 heats up and stops heating, the phase change material 03 in the thermal control unit 06 corresponding to the heating device 01 alternately melts and solidifies. The latent heat of phase change of the phase change material 03 in the thermal control unit 06 can directly suppress large temperature fluctuations of the heating device 01. Furthermore, since each of the thermal control units 06 is located in the cavity to fill the cavity, each thermal control unit 06 and the shell 02 together bear the force as part of the structural system. The spacecraft structural plate 07 provided in this application combines the functions of a structural system and a thermal control system, has a strong adaptability to the space environment, and while reducing the impact of external heat flow on the thermal interference factors of the equipment and suppressing the temperature fluctuations of the equipment, it also reduces the impact on the normal heat dissipation of the equipment.

[0047] Optionally, the heating device 01 is an optical remote sensing camera, and the phase change material 03 is n-pentadecane or n-octadecane; Alternatively, the heating device 01 is a rubidium atomic clock, and the phase change material 03 is n-nonadecane or n-tetramone; Alternatively, the heating device 01 may be a storage battery, and the phase change material 03 may be n-hexadecane.

[0048] Thus, based on the heating device 01 and its ideal operating temperature, the thermal control unit 06 uses a phase change material 03 with a phase change point temperature similar to or the same as the ideal operating temperature, thereby achieving an ideal heat dissipation effect. Preferably, when using the same type of phase change material 03, the thermal control unit 06 can have multiple specifications, with different specifications of thermal control unit 06 having different volume ratios of phase change material 03 in the receiving cavity; or, the receiving component 04 can be designed to allow the addition of phase change material 03 into the receiving cavity and to allow the phase change material 03 to be poured out of the receiving cavity. For example, the receiving component 04 can have a sealing end cap, which can be removed to allow the addition or removal of phase change material 03.

[0049] 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 skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features therein. Such modifications or substitutions 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 structural plate, characterized in that, The device includes a housing and multiple thermal control units. Each thermal control unit includes a housing and a phase change material. A housing cavity is formed within the housing, and the phase change material is located within the housing cavity. A chamber is formed within the housing, and each thermal control unit is located within the chamber to fill the chamber. The thermal control units are used to transfer heat between each other.

2. The spacecraft structural plate according to claim 1, characterized in that, The receiving element is a spherical element.

3. The spacecraft structural plate according to claim 2, characterized in that, The two adjacent thermal control units are in direct contact with each other.

4. The spacecraft structural plate according to claim 3, characterized in that, Vacuum gaps are formed between each of the thermal control units and between each of the thermal control units and the housing.

5. The spacecraft structural plate according to claim 3, characterized in that, It also includes a heat dissipation connection layer disposed in the cavity, the heat dissipation connection layer filling the gaps between each of the thermal control units and between the thermal control unit and the housing.

6. The spacecraft structural plate according to claim 2, characterized in that, The outer wall of the receiving component is polished.

7. The spacecraft structural plate according to claim 2, characterized in that, The cavity is spherical.

8. The spacecraft structural plate according to any one of claims 1 to 7, characterized in that, The thermal control unit, which is in contact with the housing, is fixedly connected to the housing.

9. A spacecraft, characterized in that, It includes a heating device and a spacecraft structural plate as described in any one of claims 1 to 7, wherein the heating device is mounted on one side of the spacecraft structural plate.

10. The spacecraft according to claim 9, characterized in that, The heating device is an optical remote sensing camera, and the phase change material is n-pentadecane or n-octadecane; Alternatively, the heating device may be a rubidium atomic clock, and the phase change material may be n-nonadecane or n-tetradecane; Alternatively, the heating device may be a storage battery, and the phase change material may be n-hexadecane.

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