A commercial aerospace energy storage and heat dissipation device

Abstract

This invention relates to the field of spacecraft thermal control technology and discloses a commercial aerospace energy storage and heat dissipation device. It includes a sealed, load-bearing outer shell. The inner side of the heat-conducting core has several phase change material (PCM) filling cavities, and several liquid metal flow channels are wound around the outside of the PCM filling cavities. Each liquid metal flow channel includes a main channel cavity at both ends and a high thermal conductivity extension cavity between two main channel cavities. Both the main channel cavity and the high thermal conductivity extension cavity have closed microchannels for liquid metal flow. This invention creates an extremely short and efficient heat transfer path by deeply coupling a high thermal conductivity liquid metal forced circulation loop with a high latent heat composite PCM material in an embedded structure. The extremely high thermal conductivity and latent heat of the liquid metal, combined with the high surface area and high thermal conductivity extension cavity outside the PCM filling cavity, enables rapid absorption and on-demand heat release of transient high heat flux.

Description

Technical Field

[0001] This invention relates to the field of spacecraft thermal control technology, specifically a commercial space energy storage and heat dissipation device. Background Technology

[0002] With the rapid development of commercial space missions, spacecraft (such as low-Earth orbit satellites and deep space probes) face increasingly severe thermal management challenges. During their on-orbit operation, they experience periodic alternations between sunny and shadowed areas, resulting in drastic fluctuations in external heat flow. Simultaneously, the intermittent operation of internal high-power electronic equipment generates significant pulsed thermal loads. If these drastic fluctuations in the thermal environment are not effectively controlled, the equipment temperature will exceed its safe operating range, severely impacting its performance and lifespan.

[0003] Current spacecraft thermal control systems primarily rely on a combination of passive and active technologies, such as radiant heat dissipation surfaces, heat pipes, and electric heaters. However, traditional solutions have inherent limitations: passive heat dissipation is constrained by the radiant area and cannot operate in shadowed regions; active electric heating consumes a large amount of valuable onboard electrical energy, increasing the weight and complexity of the power system. Therefore, developing phase change energy storage technologies capable of peak shaving and valley filling, and realizing heat transfer over time, has become a key direction for improving the efficiency and autonomy of spacecraft thermal control.

[0004] A Chinese patent (publication number CN209279746U) discloses a microchannel aluminum heat pipe based on a solid-liquid dual-working-fluid system for heat storage and dissipation. This patented technology alternately fills parallel microchannels with a liquid working fluid and a solid phase change material, and utilizes a thermally conductive framework to enhance heat transfer. While this solution combines heat storage and heat transfer functions, its "parallel isolation" structure results in significant solid-wall thermal resistance between the phase change material and the heat transfer working fluid. Furthermore, its reliance on a capillary-driven two-phase flow heat transfer mechanism presents uncertainties and even failure risks in the microgravity environment of space, making it difficult to guarantee heat transfer efficiency and reliability.

[0005] Chinese patent also discloses a liquid metal heat pipe and heat dissipation device for use in microgravity environments (CN119756039B). This patented technology focuses on optimizing the performance of liquid metal heat pipes under microgravity. Although it has extremely high heat transfer capacity, it lacks an effective phase change energy storage unit and cannot solve the problem of continuous heating during the shadow period, still requiring reliance on an external heat source. Summary of the Invention

[0006] The purpose of this invention is to provide a commercial aerospace energy storage and heat dissipation device to solve the problems mentioned in the background art.

[0007] To achieve the above objectives, the present invention provides the following technical solution: A commercial aerospace energy storage and heat dissipation device includes a sealed load-bearing outer shell, a heat-conducting core disposed on the inner side of the sealed load-bearing outer shell, a plurality of phase change material filling cavities formed on the inner side of the heat-conducting core, and a plurality of liquid metal flow channels wound around the outer side of the phase change material filling cavities on the heat-conducting core. Each liquid metal flow channel includes a main channel cavity at both ends and a high thermal conductivity extension cavity between the two main channel cavities. The main channel cavity and the high thermal conductivity extension cavity are both provided with closed microchannels for the flow of liquid metal.

[0008] As a further embodiment of the present invention: the phase change material filling cavity is filled with a composite phase change material, wherein the composite phase change material is a composite material formed with organic phase change material or molten salt phase change material as the matrix and porous expanded graphite or metal foam as the thermally conductive reinforcing skeleton.

[0009] As a further embodiment of the present invention: the liquid metal channels are distributed in an orthogonal or staggered array on the outside of the phase change material filling cavity.

[0010] As a further embodiment of the present invention: the high thermal conductivity extension cavity is one of a needle rib array, a fin array, or a columnar array, wherein the cross-section of the needle rib is circular, elliptical, or polygonal; the cross-section of the fin is plate-shaped, wavy, or perforated plate-shaped structure; and the cross-section of the columnar body is circular, square, or hexagonal.

[0011] As a further embodiment of the present invention: the sealed load-bearing outer shell and the heat-conducting core are integrally formed by 3D printing, precision casting or diffusion welding processes.

[0012] As a further aspect of the present invention: in each of the liquid metal channels, one main channel cavity is an evaporation end, used to absorb external heat, and the liquid metal in the evaporation end is in a gaseous state; the other main channel cavity is a condensation end, used to release the heat absorbed in the high thermal conductivity extension cavity, and the liquid metal in the condensation end is in a liquid state.

[0013] As a further aspect of the present invention: the liquid metal in the liquid metal channel is a gallium indium tin alloy.

[0014] As a further aspect of the present invention: the inlet and outlet of the liquid metal flow channel are connected and equipped with a space drive unit for driving the liquid metal to circulate in a microgravity environment.

[0015] Compared with the prior art, the beneficial effects of the present invention are as follows: This invention creates an extremely short and efficient heat transfer path by deeply coupling a high thermal conductivity liquid metal forced circulation loop with a high latent heat composite phase change material in an embedded structure. The extremely high thermal conductivity and latent heat of the liquid metal, combined with the high thermal conductivity extension cavity with a high surface area on the outside of the phase change material filling cavity, enable the rapid absorption of transient high heat flux and on-demand heat release.

[0016] This device, acting as a highly efficient thermal battery, can precisely transfer heat in and out of time. It rapidly stores heat during peak heat generation and releases heat smoothly during periods of low demand, thus significantly narrowing the operating temperature fluctuation range of the equipment. This effectively solves the problem of large-scale temperature cycling caused by orbital period changes in spacecraft, greatly improving the reliability and lifespan of electronic equipment. Attached Figure Description

[0017] Figure 1 This is a front view schematic diagram of a commercial aerospace energy storage and heat dissipation device; Figure 2 This is a top view schematic diagram of a commercial aerospace energy storage and heat dissipation device.

[0018] In the diagram: 1. Sealed load-bearing outer shell; 2. Thermally conductive core; 3. Phase change material filling cavity; 4. Main channel cavity; 5. High thermal conductivity extension cavity. Detailed Implementation

[0019] Please see Figures 1-2 In this embodiment of the invention, a commercial aerospace energy storage and heat dissipation device includes a sealed load-bearing outer shell 1 located on the outer layer, and a heat-conducting core 2 disposed on the inner side of the sealed load-bearing outer shell 1. The sealed load-bearing outer shell 1 and the heat-conducting core 2 are integrally formed by 3D printing, precision casting or diffusion welding process. The inner side of the heat-conducting core 2 is provided with several phase change material filling cavities 3, which are filled with composite phase change materials. The composite phase change material is a composite material formed by organic phase change material or molten salt phase change material as the matrix and porous expanded graphite or metal foam as the heat-conducting reinforcement skeleton. It has high latent heat density and good thermal conductivity. Furthermore, several liquid metal channels are formed on the heat-conducting core 2, winding around the outside of the phase change material filling cavity 3. The liquid metal channels are distributed in an orthogonal or staggered array on the outside of the phase change material filling cavity 3. Each liquid metal channel includes a main channel cavity 4 at both ends and a high thermal conductivity extension cavity 5 between the two main channel cavities 4. The main channel cavity 4 and the high thermal conductivity extension cavity 5 are both provided with closed microchannels for the flow of liquid metal.

[0020] Preferably, the high thermal conductivity extension cavity 5 is one of a needle rib array, a fin array, or a columnar array, wherein the cross-section of the needle rib is circular, elliptical, or polygonal; the cross-section of the fin is plate-shaped, wavy, or perforated plate-shaped structure; and the cross-section of the columnar body is circular, square, or hexagonal.

[0021] Preferably, in each liquid metal flow channel, one main channel cavity 4 is an evaporation end, used to absorb external heat, and the liquid metal in the evaporation end is in a gaseous state; the other main channel cavity 4 is a condensation end, used to release the heat absorbed in the high thermal conductivity extension cavity 5, and the liquid metal in the condensation end is in a liquid state. Energy storage mode (heat charging): When the spacecraft is in a sunny area or the equipment is running at high power, the evaporation end absorbs heat from the heat source, and the liquid metal absorbs heat and evaporates, causing the gaseous liquid metal to enter the high thermal conductivity extension cavity 5 and condense at the condensation end. The gaseous liquid metal condenses into liquid and releases heat. After heat transfer through the heat-conducting core 2, the composite phase change material in the phase change material filling cavity 3 undergoes a phase change (melting), storing the heat in the form of latent heat of chemical reaction. This creates a pressure difference between the evaporation end and the condensation end. Under the action of the pressure difference, the gaseous liquid metal automatically flows from the evaporation end to the condensation end along the high thermal conductivity extension cavity 5. Under the capillary action of the liquid metal flow channel, the liquid metal returns from the condensation end to the evaporation end along the high thermal conductivity extension cavity 5 and is heated and evaporated again under the action of the heat source, thus completing one phase change process of the liquid metal, and can flow back and forth along this circulation path. Energy release mode (heat release): When the spacecraft enters the shadow zone or the power consumption of the equipment decreases, the solidified composite phase change material releases latent heat, which is transferred through the heat-conducting core 2 and then heats the spacecraft through the liquid metal circulation.

[0022] Preferably, the liquid metal in the liquid metal flow channel is a gallium-indium-tin alloy; such as a eutectic or near-eutectic alloy composed of gallium, indium, and tin, with a melting point below 0°C, a boiling point above 500°C, and a thermal conductivity greater than 15 W / (m²). K); This working fluid has a low vapor pressure, making it suitable for microgravity environments in space.

[0023] Preferably, the inlet and outlet of the liquid metal flow channel are connected and equipped with a space drive unit, such as an electromagnetic pump, for driving the liquid metal to circulate in a microgravity environment, forming an actively controllable closed-loop circulation circuit; that is, an electromagnetic pump can also be used to forcefully transport the liquid metal from the condensation end back to the evaporation end along the high thermal conductivity extension cavity 5.

[0024] To better illustrate the technical effects of the present invention, the following experiments are conducted: A microchannel aluminum heat pipe based on solid-liquid dual working fluid for heat storage and heat dissipation (CN209279746U) disclosed on the patent website was used as a comparative example.

[0025] The embodiments in this application are as follows: Taking a certain type of low-orbit satellite integrated electronic box as an example; the dimensions of the sealed load-bearing shell are 250mm×150mm×40mm, and the materials of the sealed load-bearing shell and the heat-conducting core are both aluminum alloy; The liquid metal is a gallium-indium-tin alloy; the composite phase change material is a composite material of paraffin and expanded graphite, with the following performance indicators: phase change point 28℃, latent heat 200 J / g, thermal conductivity 8 W / (m·K); it is prepared by vacuum melt impregnation method, and the preparation method is as follows: Porous expanded graphite blocks are placed in a vacuum environment and molten paraffin is injected. Capillary force is used to make the paraffin fully fill the graphite pores, forming a shaped composite phase change material. This solves the leakage problem of pure paraffin and greatly improves the thermal conductivity, such as increasing the thermal conductivity from 0.2 W / (m·K) to 8 W / (m·K). The sealed load-bearing outer shell and the heat-conducting core are integrally formed from aluminum alloy using metal 3D printing technology. The main channel cavity is a plate-type main channel layer, and the high thermal conductivity extension cavity contains a total of 120 cylindrical needle ribs with a diameter of 2mm and a height of 30mm. Each needle rib has a through microchannel with a diameter of 0.8mm pre-fabricated along the axial direction at the center of its center.

[0026] The embodiments and comparative examples of the present invention were compared and analyzed in seven aspects: heat load handling capacity, high temperature suppression effect, low temperature maintenance capacity, temperature fluctuation range, power consumption of alternative active heating, thermal response time, and space environment adaptability. The experimental results are recorded in Table 1 below.

[0027] Table 1 Performance Analysis Table

[0028] Therefore, it can be concluded that the embodiments of this application have achieved a leapfrog improvement in all key performance indicators such as thermal buffering capacity, temperature stability, energy saving, and response speed; especially importantly, the present invention solves the fundamental problem of poor adaptability of the comparative examples in the microgravity environment of space.

[0029] The above description is merely a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any equivalent substitutions or modifications made by those skilled in the art within the scope of the technology disclosed in the present invention, based on the technical solution and inventive concept of the present invention, should be covered within the scope of protection of the present invention.

Claims

1. A commercial aerospace energy storage and heat dissipation device, comprising a sealed load-bearing outer shell (1) located on the outer layer, characterized in that, The inner side of the sealed load-bearing shell (1) is provided with a heat-conducting core (2). The inner side of the heat-conducting core (2) is provided with a number of phase change material filling cavities (3). The heat-conducting core (2) is provided with a number of liquid metal flow channels that are wrapped around the outside of the phase change material filling cavities (3). Each liquid metal flow channel includes a main channel cavity (4) located at both ends and a high thermal conductivity extension cavity (5) located between the two main channel cavities (4). The main channel cavity (4) and the high thermal conductivity extension cavity (5) are both provided with closed microchannels for liquid metal flow. The phase change material filling cavity (3) is filled with a composite phase change material. The composite phase change material is a composite material formed with organic phase change material or molten salt phase change material as the matrix and porous expanded graphite or metal foam as the thermally conductive reinforcing skeleton. The liquid metal flow channels are distributed in an orthogonal array or staggered array on the outside of the phase change material filling cavity (3).

2. The commercial aerospace energy storage and heat dissipation device according to claim 1, characterized in that, The high thermal conductivity extension cavity (5) is one of a needle rib array, a fin array, or a column array, wherein the cross-section of the needle rib is circular, elliptical, or polygonal; the cross-section of the fin is plate-shaped, wavy, or perforated plate-shaped structure; and the cross-section of the column is circular, square, or hexagonal.

3. The commercial aerospace energy storage and heat dissipation device according to claim 1, characterized in that, The sealed load-bearing outer shell (1) and the heat-conducting core (2) are integrally formed by 3D printing, precision casting or diffusion welding processes.

4. The commercial aerospace energy storage and heat dissipation device according to claim 1, characterized in that, In each of the liquid metal channels, one main channel cavity (4) is the evaporation end, which is used to absorb external heat, and the liquid metal in the evaporation end is in a gaseous state; the other main channel cavity (4) is the condensation end, which is used to release the heat absorbed in the high thermal conductivity extension cavity (5), and the liquid metal in the condensation end is in a liquid state.

5. A commercial aerospace energy storage and heat dissipation device according to claim 1, characterized in that, The liquid metal in the liquid metal channel is a gallium indium tin alloy.

6. A commercial aerospace energy storage and heat dissipation device according to claim 1, characterized in that, The inlet and outlet of the liquid metal flow channel are connected and equipped with a space drive unit for driving the liquid metal to circulate in a microgravity environment.

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