A finned heat sink for spacecraft

Abstract

This utility model relates to a finned heat dissipation device for spacecraft. The finned heat dissipation device includes a base plate and multiple fins. The base plate is used to connect to the spacecraft's heat source mounting surface. The base plate includes a base plate body, a first mounting surface and a second mounting surface respectively disposed at the top and bottom of the base plate body. The spacecraft's heat source mounting surface is connected to the second mounting surface. The multiple fins are used for heat dissipation and are arranged on the first mounting surface to form a fin array. The fins are three-dimensional fins. This utility model designs a finned heat dissipation device based on the external heat flow distribution of the spacecraft, expanding the heat dissipation area and improving the overall heat dissipation capacity of the spacecraft. It solves the heat dissipation problem of high-heat-consuming spacecraft and addresses the problem that traditional ground-based finned heat dissipation devices are not suitable for vacuum conditions. Under the condition of limited spacecraft heat dissipation surface, it expands the spacecraft's heat dissipation area, increasing the spacecraft's heat dissipation capacity by no less than 35%.

Description

Technical Field

[0001] This utility model relates to the field of spacecraft heat dissipation technology, and to a finned heat dissipation device for spacecraft, and more particularly to a finned heat dissipation device for spacecraft that is suitable for operation in a vacuum state. Background Technology

[0002] Traditional planar communication satellites are compact in structure but limited in space, resulting in high overall power, limited heat dissipation area, and insufficient heat dissipation capacity. Conventional heat dissipation solutions include adding deployable extended radiators or heat dissipation fins. Deployable extended radiators are complex, heavy, and contain moving parts, posing a significant risk of failure and reducing the overall reliability of the satellite. Existing finned heat dissipation technology uses convection cooling, with structures mainly consisting of densely arranged multi-fin arrays. While increasing the heat dissipation surface area and optimizing the convective heat transfer coefficient improves heat dissipation capacity, it is only suitable for equipment cooling in terrestrial atmospheric environments (requiring fluid involvement). Spacecraft are generally in a vacuum state in space, without air or other fluids for convective heat transfer. They primarily radiate heat into the cold space environment through thermal radiation. The large radiation angle coefficient between adjacent fins means that most of the radiated heat is absorbed by adjacent fins, increasing the heat dissipation area but not the overall cooling capacity. Furthermore, when the fins are exposed to sunlight, multiple reflections and absorptions of sunlight occur between adjacent fins, increasing the absorbed solar heat flux and further reducing heat dissipation capacity. In other words, heat dissipation capacity is mainly affected by factors such as heat dissipation area, spatial angle coefficient and surface optical properties. Existing convection heat exchange fins are not suitable for heat dissipation in the vacuum state of spacecraft.

[0003] Chinese patent CN206993585U discloses an alternating toothed fin heat sink. This patent features vertical toothed fins on a heat sink mounting plate, utilizing an alternating design of trapezoidal grooves and protrusions to enhance airflow disturbance and improve convective heat transfer efficiency. Its core relies on fluid media such as air for heat transfer, making it suitable for ground environments with an atmosphere. However, this structure has inherent drawbacks: First, the heat dissipation principle is entirely based on convective heat transfer, which fails in the vacuum environment of spacecraft due to the lack of fluid participation; second, the planar toothed fin configuration does not consider the optimization of the angular coefficient for radiative heat dissipation, making it prone to thermal radiation interference between adjacent fins and hindering efficient heat radiation into the cold space; third, it is difficult to balance solar radiation absorption with its own heat radiation, making it unsuitable for the high-radiation, high-vacuum conditions of space. Utility Model Content

[0004] The purpose of this invention is to provide a simple, reliable, and easy-to-install finned heat sink for spacecraft, specifically a finned heat sink suitable for operation in a vacuum. This finned heat sink can improve the heat dissipation efficiency of spacecraft surfaces, solve the problem of insufficient heat dissipation surface in spacecraft with high heat consumption, and address the incompatibility of traditional ground-based finned heat sinks in a vacuum.

[0005] This utility model provides a finned heat dissipation device for spacecraft, comprising:

[0006] A substrate includes a substrate body, a first mounting surface and a second mounting surface respectively disposed at the top and bottom of the substrate body, wherein the spacecraft heat source mounting surface is connected to the second mounting surface; and

[0007] Multiple fins for heat dissipation are arranged on the first mounting surface to form a fin array, wherein the fins protrude from the first mounting surface.

[0008] Furthermore, the fins are multi-angle three-dimensional fins.

[0009] Furthermore, the substrate is a planar substrate or a curved substrate.

[0010] Furthermore, a thermally conductive filler layer is provided on the second mounting surface. This further reduces thermal resistance and improves heat dissipation efficiency.

[0011] Furthermore, the thermally conductive filler layer includes any one of a thermally conductive silicone grease layer, a thermally conductive silicone rubber layer, a metal filler layer, a thermally conductive gel layer, or a thermally conductive pad; the metal filler layer is an indium foil layer.

[0012] Furthermore, the fins are pyramidal in shape to form pyramidal fins. The pyramidal fins increase the heat dissipation area while reducing mutual radiative heat transfer between fins, and also reduce multiple reflections and absorptions of sunlight.

[0013] Furthermore, the fin comprises multiple fin bodies forming a pyramidal fin, and the angle between the fin body and the substrate is 45° to 60°. The angle between the fin body and the substrate is not limited to the above angle and can be adjusted and determined according to the angle of solar incidence.

[0014] Furthermore, the pyramidal fin is a pyramidal fin with a cavity, and the corresponding middle area of ​​the substrate and the fin is cut off to further reduce the weight of the fin.

[0015] Furthermore, the larger the angle between the incident sunlight and the substrate, the smaller the angle between the fin and the substrate should be. When the angle between the incident sunlight and the substrate is 90°, the angle between the fin and the substrate should be 45°; when the angle between the incident sunlight and the substrate is 0°...

[0016] At that time, the angle between the fin body and the substrate is 60°.

[0017] Furthermore, the angle between the fins and the substrate is designed to vary depending on the external heat flow conditions of the spacecraft in the fin mounting area. This facilitates the reflection of solar radiation heat while simultaneously reducing the angular coefficient between the fins, thus reducing heat radiation between adjacent fins, and increasing the angular coefficient between the fins and the cold space, which is beneficial for heat radiation to the cold space.

[0018] Furthermore, it also includes:

[0019] A top plate is disposed on top of the fins to form a closed cavity between the fins, the first mounting surface, and the top plate.

[0020] Furthermore, the fins are also provided with vent holes to facilitate venting under vacuum conditions.

[0021] Furthermore, the fins are any one of aluminum fins, copper fins, or stainless steel fins.

[0022] Furthermore, the plurality of fins are arranged at intervals.

[0023] Furthermore, the substrate, fins, and top plate are welded or integrally formed, reducing the thermal resistance between different components.

[0024] Furthermore, the surfaces of the substrate, fins, and top plate are provided with a thermal control coating or an OSR (Optical Surface Reflector). This reduces solar radiation heat and increases heat dissipation into the cool space.

[0025] Furthermore, the secondary surface mirror is a cerium glass type secondary surface mirror.

[0026] Furthermore, the thermal control coating is a low-absorption, high-heat-generating thermal control white paint layer, which includes an organic heat-dissipating white paint layer or an inorganic heat-dissipating white paint layer.

[0027] Furthermore, the low-absorption, high-heat-generating white paint includes any one of S781 white paint, SR107-ZK white paint, KS-ZA white paint, or KS-1 white paint. The low-absorption, high-heat-generating white paint layer is formed from the low-absorption, high-heat-generating white paint.

[0028] This invention has at least the following beneficial effects: 1) Based on the distribution of external heat flow in the spacecraft, this invention designs a finned heat dissipation device to expand the heat dissipation area and improve the overall heat dissipation capacity of the spacecraft. Specifically, by using pyramidal fins, the heat dissipation area is increased while reducing mutual radiation heat transfer between fins and repeated reflection of sunlight. At the same time, a low-absorption, high-heat-controlling white paint layer is sprayed on the fin surface to further reduce the solar-absorbed heat flow and improve the net heat dissipation capacity of the heat dissipation fins. This solves the heat dissipation problem of spacecraft with high heat consumption and also solves the problem that traditional ground-based finned heat dissipation devices are not suitable for vacuum conditions; 2) Based on the different external heat flow on the mounting surface, this invention designs heat dissipation fins at different angles, which not only reflect sunlight and reduce external heat flow, but also increase the heat dissipation of the fins to the cold space. Under the condition of limited heat dissipation surface of the spacecraft, the heat dissipation area of ​​the spacecraft is expanded, and the heat dissipation capacity of the spacecraft is increased by no less than 35%; 3) Compared with complex active thermal control systems or fluid loop heat dissipation systems, this invention has a high degree of integration, is easy to arrange on the surface of the spacecraft, and has no complex moving parts, resulting in high reliability. Attached Figure Description

[0029] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the various embodiments of the present invention will be presented with reference to the accompanying drawings. It is understood that these drawings depict only typical embodiments of the present invention and are therefore not intended to limit its scope. In the drawings, for clarity, the same or corresponding parts will be indicated by the same or similar reference numerals.

[0030] Figure 1 The following is a schematic diagram of the structure of a finned heat sink for spacecraft in some embodiments of the present invention;

[0031] Figure 2 The following is a front view of a finned heat sink for spacecraft in some embodiments of the present invention;

[0032] Figure 3 It shows Figure 2 A magnified view of a portion of point A in the middle;

[0033] Figure 4 A top view of a finned heat sink for spacecraft is shown in some embodiments of the present invention;

[0034] Figure 5 The following is a schematic diagram of the structure of a finned heat sink for spacecraft in some embodiments of the present invention;

[0035] Figure 6 The diagram shows a schematic representation of the cut-off region in some embodiments of the present invention;

[0036] Figure label:

[0037] 1-Substrate, 101-Substrate body, 102-First mounting surface, 103-Second mounting surface, 2-Fin, 3-Top plate, 4-Vent hole, 5-Thermal conductive filler layer, 6-Thermal control coating, 7-Removed area. Detailed Implementation

[0038] It should be noted that the components in the accompanying drawings may be shown exaggerated for illustrative purposes and may not be to scale.

[0039] In this utility model, the various embodiments are merely intended to illustrate the solution of this utility model and should not be construed as limiting.

[0040] In this utility model, unless otherwise specified, the quantifiers “one” and “one” do not exclude scenarios involving multiple elements.

[0041] It should also be noted that in the embodiments of this utility model, only a portion of the parts or components may be shown for clarity and simplicity. However, those skilled in the art will understand that, under the teachings of this utility model, the required parts or components can be added according to the specific scenario.

[0042] It should also be noted that within the scope of this utility model, the terms "same", "equal", and "equal to" do not mean that the two values ​​are absolutely equal, but allow for a certain reasonable error. In other words, the terms also cover "substantially the same", "substantially equal", and "substantially equal to".

[0043] It should also be noted that in the description of this utility model, the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "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 utility model and for simplifying the description, and do not explicitly or implicitly suggest 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 utility model. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0044] Furthermore, the embodiments of this utility model describe the process steps in a specific order. However, this is only for the convenience of distinguishing each step, and is not a limitation on the order of each step. In different embodiments of this utility model, the order of each step can be adjusted according to the process.

[0045] In the following embodiments, the fin 2 is any one of aluminum fin, copper fin, or stainless steel fin; the thermal control coating 6 is a low-absorption, high-heat-generating thermal control white paint layer, which is formed by low-absorption, high-heat-generating thermal control white paint, including any one of S781 white paint, SR107-ZK white paint, KS-ZA white paint, or KS-1 white paint; the thermally conductive filler layer includes any one of thermally conductive silicone grease layer, thermally conductive silicone rubber layer, indium foil layer, thermally conductive gel layer, or thermally conductive pad.

[0046] The following embodiment provides a finned heat sink for spacecraft. Figure 1 This diagram shows the structure of a finned heat sink for spacecraft. Figure 2 The front view of a finned heat sink for spacecraft is shown. Figure 3 It shows Figure 2 A magnified view of a portion of point A in the diagram. Figure 4 A top view of a finned heat sink for spacecraft is shown; the finned heat sink includes:

[0047] Substrate 1, substrate 1 is a planar substrate 1, substrate 1 includes substrate body 101, first mounting surface 102 and second mounting surface 103 respectively disposed on the top and bottom of substrate body 101, wherein the spacecraft heat source mounting surface is connected to the second mounting surface 103, and a thermally conductive filler layer 5 is provided on the second mounting surface 103 to further reduce the thermal resistance and improve the heat dissipation efficiency.

[0048] Multiple fins 2 are used for heat dissipation. These fins 2 are spaced apart on the first mounting surface 102 to form a fin array. The fins 2 are pyramidal in shape to form pyramidal fins. The pyramidal fins are pyramidal fins with cavities, meaning that the corresponding middle area between the substrate 1 and the fins 2 is cut off. The cut-off area 7 is as follows: Figure 6 As shown, to further reduce the weight of fin 2, fin 2 includes multiple fin bodies forming a pyramidal fin 2. The angle between the fin body and the substrate 1 is 45° to 60°, which facilitates the reflection of solar radiation heat, while also reducing the angular coefficient between fins 2, reducing heat radiation between adjacent fins 2, and increasing the angular coefficient between fin 2 and the cold space, which is beneficial for heat radiation to the cold space; and

[0049] A top plate 3 is located on top of the fins 2, forming a closed cavity between the fins 2, the first mounting surface 102, and the top plate 3. Vent holes 4 are also provided on the fins 2 for venting under vacuum conditions. The base plate 1, fins 2, and top plate 3 are welded or integrally formed to reduce thermal resistance between different components. A thermal control coating 6 is provided on the surfaces of the base plate 1, fins 2, and top plate 3 to adapt to the space radiation heat dissipation environment, reduce solar radiation heat, and increase heat dissipation into the cold space. After using this finned-2 type heat dissipation device for spacecraft, under the condition of limited spacecraft heat dissipation surface, the spacecraft's heat dissipation area can be expanded, increasing the spacecraft's heat dissipation capacity by no less than 35%.

[0050] In some embodiments, substrate 1 is a curved substrate 1, such as... Figure 5 As shown. Fin 2 operates in the shaded area, where there is no solar heat flow, and increasing the heat dissipation area is more conducive to providing heat dissipation area; when fin 2 operates in the illuminated area, it is affected by solar heat flow, and the pyramidal fins can increase the heat dissipation area, reduce mutual radiative heat transfer between fins, and at the same time reduce multiple reflections and absorption of sunlight.

[0051] In some embodiments, the surfaces of the substrate 1, fins 2 and top plate 3 are provided with secondary surface mirrors, which are cerium glass type secondary surface mirrors.

[0052] In some embodiments, the top plate 3 and the vent 4 can be omitted depending on the application scenario.

[0053] In some embodiments, the angle between the fin and the substrate 1 is designed differently based on the external heat flow of the spacecraft in the fin installation area. That is, the angle of the extended fin 2 can be adaptively adjusted according to the external heat flow distribution of the spacecraft, facilitating the reflection of solar radiation heat, while simultaneously reducing the radiation angle coefficient between the fins 2 and increasing the angle coefficient between the fins 2 and the cold space, thereby improving the heat dissipation of the cold space. Specifically, the larger the angle between the incident solar light and the substrate, the smaller the angle between the fin and the substrate should be. When the angle between the incident solar light and the substrate is 90°, the angle between the fin and the substrate should be 45°.

[0054] When the angle between the solar incident light and the substrate is 0°, the angle between the fin and the substrate is 60°, the heat dissipation area increases by 36.7%, and the net heat dissipation capacity increases by 35.3%.

[0055] In some embodiments, the heat dissipation fins 2 can be integrally formed, which has a simple structure, high reliability, and is easy to process and implement.

[0056] While some embodiments of this invention have been described in this application, those skilled in the art will understand that these embodiments are merely illustrative. Numerous variations, alternatives, and improvements will arise in those skilled in the art under the teachings of this invention without departing from its scope. The appended claims are intended to define the scope of this invention and thereby cover the methods and structures within the scope of the claims themselves and their equivalents.

Claims

1. A finned heat dissipation device for spacecraft, characterized in that, include: The substrate includes a substrate body, a first mounting surface and a second mounting surface respectively disposed on the top and bottom of the substrate body, wherein the spacecraft heat source mounting surface is connected to the second mounting surface. as well as Multiple fins for heat dissipation are arranged on the first mounting surface to form a fin array, wherein the fins protrude from the first mounting surface.

2. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The substrate is a planar substrate or a curved substrate.

3. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, A thermally conductive filler layer is provided on the second mounting surface.

4. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The fins are pyramidal in shape to form pyramidal fins.

5. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, Also includes: A top plate is disposed on top of the fins to form a closed cavity between the fins, the first mounting surface, and the top plate.

6. The finned heat dissipation device for spacecraft according to claim 5, characterized in that, The fins are also provided with vent holes.

7. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The fins are any one of aluminum fins, copper fins, or stainless steel fins.

8. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The multiple fins are arranged at intervals.

9. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The substrate, fins, and top plate are formed by welding or integral processing.

10. The finned heat dissipation device for spacecraft according to claim 1, characterized in that, The surfaces of the substrate, fins, and top plate are provided with a thermal control coating or a secondary surface mirror.

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