A satellite device with heat dissipation function for high-power, high-computing-power onboard chips.
By setting up heat dissipation zones and a stepped capillary wick structure on the satellite shell, the heat dissipation problem of high-performance chips in the space environment was solved, achieving efficient, stable, and low-cost heat dissipation, and improving the reliability and performance of the chips.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Utility models(China)
- Current Assignee / Owner
- SHANGHAI XUNTIAN QIANHE SPACE TECHNOLOGY CO LTD
- Filing Date
- 2025-07-31
- Publication Date
- 2026-06-30
Smart Images
Figure CN224439497U_ABST
Abstract
Description
Technical Field
[0001] This utility model belongs to the field of aerospace equipment technology, and in particular relates to a satellite device with heat dissipation function for high-power, high-computing-power onboard chips. Background Technology
[0002] As a core infrastructure of the digital economy, chip computing power has achieved leapfrog development in recent years, driven by explosive demand in fields such as artificial intelligence, autonomous driving, and cloud computing. The construction of navigation satellite constellations and mega-communication satellite constellations has ushered in a new boom in aerospace applications, giving rise to a new type of space-based resource—space-based computing systems—based on space-based high-performance computing chips. Over the past few decades, thanks to the development of communication internet and computing resources, large-scale, real-time knowledge sharing, communication interconnection, and even collaborative work have gradually become the norm on Earth. However, in space, the traditional working mode of passive operation by a single satellite, data transit and relay, and post-processing on the ground means that the capabilities of space-based high-performance computing are far from being fully realized.
[0003] Correspondingly, introducing high-performance computing chips into space-based platforms has greatly enhanced processing capabilities, but it has also brought severe challenges in heat flux density. These chips generate a large amount of heat during operation, and how to efficiently and reliably dissipate this heat in the special space environments such as vacuum and microgravity has become a core bottleneck limiting the performance and long-term stable operation of high-performance computing payloads on low-Earth orbit satellites. The heat flux density of space-based high-performance computing chips has increased from approximately 10 W / cm² in the early 2000s. 2 Climbing to the current 200-500W / cm 2 (Some advanced GPU / AI chips even exceed 1000W / cm²) 2 Traditional heat dissipation technologies face severe challenges. Excessively high temperatures reduce chip stability and increase error rates. Furthermore, thermal stress generated within the module and by the external environment directly impacts chip performance, operating frequency, mechanical strength, and reliability. Heat dissipation capacity has become a critical bottleneck restricting the development and performance improvement of high-performance chips.
[0004] Under current technology, traditional heat dissipation solutions include: using pre-embedded heat pipes for heat exchange, which is technically mature and structurally simple, but has limited heat dissipation capacity and is difficult to effectively exchange heat exceeding 100 watts, failing to meet the heat dissipation requirements of high-performance computing chips. The orthogonal heat pipe network + deployable thermal radiator solution has strong heat dissipation capacity, but the deployable radiator design is relatively complex, and the on-orbit deployment of such mechanical products carries certain risks. Failure to deploy properly will lead to heat dissipation failure, and the external radiators are often quite heavy, adding to the overall launch cost of the satellite. The heat dissipation solution using pre-embedded heat pipes + a mechanically pumped single-phase fluid loop is also complex, especially in key technologies such as long-life circulating pumps, working fluid compatibility, and fluid loop operation and control, which still require breakthroughs. The stability and reliability of long-term on-orbit operation remain to be verified. Utility Model Content
[0005] This invention provides a satellite device with heat dissipation function for high-power, high-computing-power onboard chips, in order to solve the technical problem that under the existing technology, high-power, high-computing-power onboard chips in conventional satellite devices cannot achieve efficient, stable, reliable, and low-cost heat dissipation.
[0006] To solve the above problems, the technical solution of this utility model is: a satellite device with heat dissipation function for high-power, high-computing-power onboard chips, comprising:
[0007] The satellite shell has a heat dissipation area, which includes a first side and a second side. The first side of the heat dissipation area faces the internal accommodating chamber of the satellite equipment, and the second side of the heat dissipation area faces the external space of the satellite equipment. Under the principle of temperature gradient drive, the heat dissipation area is used to conduct heat from inside the satellite equipment to the external space of the satellite equipment.
[0008] The satellite-borne chip is disposed in the internal cavity of the satellite equipment. The heat source surface of the satellite-borne chip is attached to the first side of the heat dissipation area, and a thermally conductive medium layer is provided between the heat source surface of the satellite-borne chip and the first side of the heat dissipation area.
[0009] The satellite equipment is configured such that the heat generated by the onboard chip is conducted through the thermally conductive medium layer and radiated outward into space through the heat dissipation area of the satellite shell.
[0010] Preferably, the heat dissipation area is provided with a first cover plate located on the first side and a second cover plate located on the second side. The first cover plate and the second cover plate are spliced together by solid diffusion welding to form a sealed cavity inside. The sealed cavity is provided with a stepped capillary core structure that fits the inner walls of the first cover plate and the second cover plate and is connected in an annular manner. A gaseous or liquid heat transfer working fluid circulates in the stepped capillary core structure.
[0011] The first cover plate and the second cover plate are made of metal.
[0012] Preferably, one end of the sealed cavity is provided with a working fluid filling port that communicates with the external space. The working fluid filling port is connected to the internal channel of the stepped capillary structure and is used to replenish the heat transfer working fluid to the stepped capillary structure.
[0013] The working fluid filling port is sealed by brazing.
[0014] Preferably, the sealed cavity is provided with a plurality of copper pillars at both ends that are fixedly connected to the inner walls of the first cover plate and the second cover plate, respectively, to provide mechanical support for the heat dissipation area.
[0015] Preferably, the thermally conductive medium layer includes a first thermally conductive unit and a second thermally conductive unit, one end face of the first thermally conductive unit is in contact with the heat source surface of the spaceborne chip, and one end face of the second thermally conductive unit is in contact with the surface of the heat dissipation area.
[0016] A connecting unit is provided between the first heat-conducting unit and the second heat-conducting unit for fixing the first heat-conducting unit and the second heat-conducting unit together.
[0017] Preferably, the first heat-conducting unit and the second heat-conducting unit are made of graphene three-dimensional skeleton or mesophase pitch-based carbon fiber three-dimensional skeleton, and the connecting unit is made of silicone rubber.
[0018] Preferably, the satellite shell includes a first shell, a second shell, a third shell, a fourth shell, a fifth shell, and a sixth shell, and the satellite shell is made of an aluminum honeycomb panel structure with a thickness of 20mm.
[0019] Preferably, when the satellite equipment is in a stable three-axis attitude in orbit, the first shell is in orbit flying relative to the ground, and the first shell is equipped with an emergency telemetry and control antenna, an X telemetry and control antenna, an X data transmission antenna and a first simulated solar angle calculation device.
[0020] The second housing is for the on-orbit flight of the sky surface, and the second housing is equipped with an X-ray telemetry and control antenna, a GNSS antenna and a second simulated solar angle calculation device;
[0021] A thruster is arranged on the third housing.
[0022] Preferably, the fourth housing and the fifth housing are provided with deployable solar panels. When the deployable solar panels are deployed, the fourth housing or the fifth housing forms a shaded side, and the heat dissipation area is located in the fourth housing or the fifth housing.
[0023] Preferably, the satellite equipment has a partition inside, which is used to support and fix the satellite shell.
[0024] Because of the adoption of the above technical solution, this utility model has the following advantages and positive effects compared with the prior art:
[0025] This utility model provides a satellite device with heat dissipation function for high-power and high-computing-power onboard chips. A heat dissipation area is provided on the satellite shell made of metal material. The heat dissipation area has a stepped capillary core structure inside, which makes the heat dissipation area have good heat dissipation performance. The heat source surface of the onboard chip is attached to the inner wall surface of the heat dissipation area, and a thermally conductive medium layer is provided between the onboard chip and the heat dissipation area. The thermally conductive medium layer is made of a graphene three-dimensional skeleton or a mesophase pitch-based carbon fiber three-dimensional skeleton that is easy to conduct heat. In this embodiment, when the high-power, high-computing-power onboard chip is operating, the heat from the heat source surface of the onboard chip is directly conducted through the thermally conductive medium layer and radiated outwards into space through the heat dissipation area of the satellite shell. This effectively reduces intermediate heat transfer components, shortens the heat transfer path, and lowers the thermal resistance, thereby reducing the temperature difference between the onboard chip and the heat sink and improving the heat dissipation efficiency of the onboard chip. At the same time, based on the integrated design of the heat dissipation area of the satellite shell, the need for additional radiator placement is reduced, effectively reducing the overall weight and size of the satellite equipment and further reducing the manufacturing cost of the satellite equipment. This achieves efficient, stable, reliable, and low-cost heat dissipation for the onboard chip, providing strong technical support for the "launch" and on-orbit application of high-power, high-computing-power onboard chips. Attached Figure Description
[0026] Figure 1 This utility model provides a schematic diagram of the contact area between the spaceborne chip and the heat dissipation area;
[0027] Figure 2 This utility model provides a schematic diagram of the structure of the satellite equipment;
[0028] Figure 3 This utility model provides a schematic diagram of the satellite shell structure;
[0029] Figure 4 A schematic diagram of the heat dissipation area provided by this utility model;
[0030] Figure 5 This utility model provides a schematic diagram of the structure of the heat-conducting medium layer.
[0031] Explanation of reference numerals in the attached drawings: 1: Satellite shell; 11: Heat dissipation area; 111: First cover plate; 112: Second cover plate; 113: Stepped capillary wick structure; 114: Working fluid filling port; 115: Copper pillar; 101: First shell; 102: Second shell; 103: Third shell; 104: Fourth shell; 105: Fifth shell; 106: Sixth shell; 107: Partition; 108: Deployable solar array; 2: Onboard chip; 3: Thermal conductive medium layer; 31: First thermal conductive unit; 32: Second thermal conductive unit; 33: Connection unit. Detailed Implementation
[0032] The following detailed description, in conjunction with the accompanying drawings and specific embodiments, provides a satellite device with high-power, high-computing-power onboard chip heat dissipation functionality according to this invention. The advantages and features of this invention will become clearer from the following description and claims.
[0033] See Figures 1-5 This embodiment provides a satellite device with a high-power, high-computing-power onboard chip heat dissipation function, which is used to achieve efficient, stable, reliable and low-cost heat dissipation of the onboard chip 2. The main structure of the satellite device includes a satellite shell 1 and an onboard chip 2.
[0034] The satellite shell 1 is provided with a heat dissipation area 11. The heat dissipation area 11 forms a first side and a second side on the satellite shell 1. The first side of the heat dissipation area 11 faces the internal accommodating chamber of the satellite equipment, and the second side of the heat dissipation area 11 faces the external space of the satellite equipment. Under the principle of temperature gradient drive, the heat dissipation area 11 is used to conduct heat from inside the satellite equipment to the external space of the satellite equipment.
[0035] The onboard chip 2 is housed within the internal cavity of the satellite equipment. The heat source surface of the onboard chip 2 is in contact with the first side of the heat dissipation area 11, and the heat source surface of the onboard chip 2 is defined as the end face of the onboard chip 2 away from the PCB board. Furthermore, a thermally conductive medium layer 3 is provided between the heat source surface of the onboard chip 2 and the first side of the heat dissipation area 11, which is used to further improve the thermal conductivity between the onboard chip 2 and the heat dissipation area 11.
[0036] In this embodiment, the satellite equipment is configured such that the heat generated by the onboard chip 2 is conducted through the thermally conductive medium layer 3 and directly radiated into the external space through the heat dissipation area 11 of the satellite shell 1.
[0037] Therefore, in this embodiment, the efficient heat transfer scheme of the onboard chip 2-satellite shell 1 heat dissipation area 11 is adopted, which is different from the complex heat transfer scheme of the traditional onboard chip-PCB board-satellite shell-satellite compartment board / radiator board. This embodiment effectively reduces the intermediate transition / transfer heat transfer components, shortens the heat transfer path, and reduces the thermal resistance, thereby reducing the temperature difference between the onboard chip 2 and the heat sink and improving the heat dissipation efficiency of the onboard chip 2. At the same time, based on the integrated design of the heat dissipation area 11 of the satellite shell 1, the additional radiator arrangement requirements are reduced, effectively reducing the overall weight and size of the satellite equipment, and further reducing the manufacturing cost of the satellite equipment. Thus, the efficient, stable, reliable and low-cost heat dissipation function of the onboard chip 2 is achieved.
[0038] The following will provide a more detailed description of the specific structure and implementation functions of a satellite device with high power consumption and high computing power onboard chip heat dissipation function provided in this embodiment:
[0039] Preferably, in one embodiment, the heat dissipation area 11 is provided with a first cover plate 111 located on the first side and a second cover plate 112 located on the second side. After the first cover plate 111 and the second cover plate 112 are spliced and fixed by solid diffusion welding, a sealed cavity is formed inside the heat dissipation area 11 constructed by the first cover plate 111 and the second cover plate 112. The sealed cavity is provided with a stepped capillary core structure 113 that fits the inner wall of the first cover plate 111 and the second cover plate 112 and is connected in an annular manner. A heat transfer working fluid circulates in the stepped capillary core structure 113.
[0040] Furthermore, the first cover plate 111 and the second cover plate 112 are made of metal.
[0041] Among them, the stepped capillary wick structure 113 is a heat pipe device with an evaporation section and a condensation section to realize heat absorption and conduction. By designing microribs or pores of different levels, the stepped capillary wick structure 113 can provide stronger capillary force through a progressively changing structure, thereby promoting the flow of heat transfer working fluid in the capillary wick and enabling the heat pipe to transport heat more efficiently.
[0042] Heat transfer fluids include gaseous or liquid fluids, such as water, acetone, methanol, alcohol mixtures, Freon, ammonia, ethylene glycol, etc.
[0043] Preferably, in one embodiment, one end of the sealed cavity is provided with a working fluid filling port 114 that communicates with the external space. The working fluid filling port 114 is connected to the internal channel of the stepped capillary structure 113. The working fluid filling port 114 can be used to supplement the heat transfer working fluid to the stepped capillary structure 113.
[0044] The working fluid filling port 114 is sealed by brazing.
[0045] Preferably, in one embodiment, the sealed cavity is provided with a plurality of copper pillars 115 at both ends which are fixedly connected to the inner walls of the first cover plate 111 and the second cover plate 112, respectively. The copper pillars 115 are used to provide mechanical support for the heat dissipation area 11 and improve the stability of the heat dissipation area 11.
[0046] Preferably, in one embodiment, the thermally conductive medium layer 3 has a multi-layer structure, including a first thermally conductive unit 31 and a second thermally conductive unit 32. One end face of the first thermally conductive unit 31 is attached to the heat source surface of the onboard chip 2, and one end face of the second thermally conductive unit 32 is attached to the surface of the heat dissipation area 11.
[0047] A connecting unit 33 is provided between the first heat-conducting unit 31 and the second heat-conducting unit 32 for fixing the first heat-conducting unit 31 and the second heat-conducting unit 32.
[0048] Specifically, the first heat-conducting unit 31 and the second heat-conducting unit 32 are made of graphene three-dimensional skeleton or mesophase pitch-based carbon fiber three-dimensional skeleton, and the connecting unit 33 is made of silicone rubber.
[0049] In this embodiment, both the graphene three-dimensional framework and the mesophase pitch-based carbon fiber three-dimensional framework have extremely high thermal conductivity. Through the first thermally conductive unit 31 and the second thermally conductive unit 32, rapid heat transfer between the heat source side of the spaceborne chip 2 and the heat dissipation area 11 can be achieved. Simultaneously, silicone rubber is disposed between the first thermally conductive unit 31 and the second thermally conductive unit 32. Silicone rubber has excellent flexibility and elasticity, which can maintain high thermal conductivity while imparting good flexibility and mechanical properties to the thermally conductive medium layer 3. Furthermore, silicone rubber can improve the contact effect between the two thermally conductive units and reduce interfacial thermal resistance.
[0050] Preferably, in one embodiment, the satellite shell 1 includes a first shell 101, a second shell 102, a third shell 103, a fourth shell 104, a fifth shell 105 and a sixth shell 106. The satellite shell 1 is made of an aluminum honeycomb panel structure with a thickness of 20mm. The non-load-bearing areas of the satellite shell 1 are locally weight-reduced to maximize the reduction of the structural weight of the satellite shell 1.
[0051] Preferably, in one embodiment, when the satellite equipment is in a stable three-axis attitude in orbit, the first housing 101 is in orbit flying relative to the ground, and the first housing 101 is equipped with an emergency telemetry and control antenna, an X telemetry and control antenna, an X data transmission antenna, and a first simulated solar angle calculation device.
[0052] The second housing 102 is for the on-orbit flight of the celestial surface. The second housing 102 is equipped with an X-ray telemetry and control antenna, a GNSS antenna, and a second simulated solar angle calculation device.
[0053] The third housing 103 is equipped with a thruster, which is used to adjust the on-orbit attitude of the satellite equipment.
[0054] Preferably, in one embodiment, the fourth housing 104 and the fifth housing 105 are provided with deployable solar panels 108. The surface of the deployable solar panels 108 is provided with a solar cell array. Before launch, the deployable solar panels 108 are in a retracted state. After the satellite enters orbit, the deployable solar panels 108 are in an extended state, providing energy to the satellite through the solar cell array. Furthermore, when the deployable solar panels 108 are deployed, they act as a light barrier, causing the fourth housing 104 or the fifth housing 105 to form a shaded side. A heat dissipation area 11 is located within the fourth housing 104 or the fifth housing 105, improving the heat dissipation efficiency of the satellite housing 1.
[0055] Preferably, in one embodiment, the satellite equipment is provided with a partition 107 inside, which is used to support and fix the satellite shell 1, thereby improving the structural stability of the satellite equipment during takeoff and in-orbit flight.
[0056] The embodiments of the present invention have been described in detail above with reference to the accompanying drawings, but the present invention is not limited to the above embodiments. Even if various changes are made to the present invention, if these changes fall within the scope of the claims of the present invention and their equivalents, they shall still fall within the protection scope of the present invention.
Claims
1. A satellite device with high power consumption and high computing power satellite chip heat dissipation function, characterized by, include: The satellite shell has a heat dissipation area, which includes a first side and a second side. The first side of the heat dissipation area faces the internal accommodating chamber of the satellite equipment, and the second side of the heat dissipation area faces the external space of the satellite equipment. Under the principle of temperature gradient drive, the heat dissipation area is used to conduct heat from inside the satellite equipment to the external space of the satellite equipment. The satellite-borne chip is disposed in the internal cavity of the satellite equipment. The heat source surface of the satellite-borne chip is attached to the first side of the heat dissipation area, and a thermally conductive medium layer is provided between the heat source surface of the satellite-borne chip and the first side of the heat dissipation area. The satellite equipment is configured such that the heat generated by the onboard chip is conducted through the thermally conductive medium layer and radiated outward into space through the heat dissipation area of the satellite shell.
2. The satellite device with high power consumption and high computing power on-board chip heat dissipation function according to claim 1, characterized in that, The heat dissipation area is provided with a first cover plate located on the first side and a second cover plate located on the second side. The first cover plate and the second cover plate are spliced together by solid diffusion welding to form a sealed cavity inside. The sealed cavity is provided with a stepped capillary core structure that fits the inner wall of the first cover plate and the second cover plate and is connected in an annular manner. A gaseous or liquid heat transfer working fluid circulates in the stepped capillary core structure. The first cover plate and the second cover plate are made of metal.
3. The satellite device with high power consumption and high computing power on-board chip heat dissipation function according to claim 2, characterized in that, One end of the sealed cavity is provided with a working fluid filling port that connects to the external space. The working fluid filling port is connected to the internal channel of the stepped capillary structure and is used to replenish the heat transfer working fluid to the stepped capillary structure. The working fluid filling port is sealed by brazing.
4. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 2, characterized in that, The sealed cavity is provided with several copper pillars at both ends that are fixedly connected to the inner walls of the first cover plate and the second cover plate, respectively, to provide mechanical support for the heat dissipation area.
5. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 1, characterized in that, The thermally conductive medium layer includes a first thermally conductive unit and a second thermally conductive unit. One end face of the first thermally conductive unit is in contact with the heat source surface of the spaceborne chip, and one end face of the second thermally conductive unit is in contact with the surface of the heat dissipation area. A connecting unit is provided between the first heat-conducting unit and the second heat-conducting unit for fixing the first heat-conducting unit and the second heat-conducting unit together.
6. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 5, characterized in that, The first and second heat-conducting units are made of graphene three-dimensional skeleton or mesophase pitch-based carbon fiber three-dimensional skeleton, and the connecting unit is made of silicone rubber.
7. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 1, characterized in that, The satellite shell includes a first shell, a second shell, a third shell, a fourth shell, a fifth shell, and a sixth shell, and the satellite shell is made of an aluminum honeycomb panel structure with a thickness of 20mm.
8. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 7, characterized in that, When the satellite equipment is set to be in a stable three-axis attitude in orbit, the first shell is the ground in orbit, and the first shell is equipped with an emergency telemetry and control antenna, an X telemetry and control antenna, an X data transmission antenna and a first simulated solar angle calculation device. The second housing is for the on-orbit flight of the sky surface, and the second housing is equipped with an X-ray telemetry and control antenna, a GNSS antenna and a second simulated solar angle calculation device; A thruster is arranged on the third housing.
9. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 7, characterized in that, The fourth and fifth housings are provided with deployable solar panels. When the deployable solar panels are deployed, the fourth or fifth housing forms a shaded side, and the heat dissipation area is located in the fourth or fifth housing.
10. The satellite equipment with high-power, high-computing-power onboard chip heat dissipation function as described in claim 7, characterized in that, The satellite equipment has a partition inside, which is used to support and fix the satellite shell.