Self-adaptive thermal control system based on pump-driven two-phase flow and thermochromism cooperative regulation and control

Abstract

The invention relates to the technical field of spacecraft thermal control, and discloses a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromism cooperative regulation, and the system is dynamically coupled with a thermal switch and a thermochromism coating through a pump-driven two-phase flow loop to realize autonomous bidirectional regulation of a thermal environment in a spacecraft. When heat dissipation is needed in the cabin, the pump drives the two-phase flow to start, the temperature of the condenser rises to enable the bottom thermal switch to expand, the condensation plate is pushed to be in contact with the radiation heat dissipation plate to transfer heat, meanwhile, the thermochromic coating is heated to be in a high infrared emissivity state, and an active heat dissipation channel is constructed. And when heat dissipation is not needed, the temperature of the condenser is reduced, the thermal switch shrinks and resets, the condensation plate is separated from the radiation heat dissipation plate, the thermochromism coating recovers the low-emissivity state, and double blocking of physical isolation and radiation suppression of the heat dissipation channel is achieved. The contradiction between passive heat leakage and active power consumption of a traditional thermal control system can be solved, and the environmental adaptability and the energy utilization efficiency of a spacecraft thermal management system are improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of spacecraft thermal control technology, in particular to a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation. BACKGROUND

[0002] During the on-orbit operation of a spacecraft, the internal thermal load presents significant dynamic fluctuation characteristics: under high-power consumption conditions, the heat generated by the equipment needs to be quickly discharged, and under low-power consumption or hibernation conditions, the cold intrusion of the external deep cold environment needs to be strictly blocked. The traditional thermal control system faces a core contradiction in such scenarios: the dynamic change of heat dissipation demand and the static characteristics of the regulation means. When the equipment is in a low thermal load or no heat dissipation demand, the fixed heat conduction path and the constant-emissivity radiation plate still continuously dissipate heat to space, resulting in a sharp drop in cabin temperature, which needs to rely on active heating compensation (such as electric heaters) to maintain the temperature control target. This double burden of passive heat leakage and active energy consumption seriously restricts the energy efficiency and reliability of long-term missions of spacecraft.

[0003] In the field of spacecraft thermal control, existing technical solutions generally have the technical bottleneck of single regulation dimension for the cold intrusion problem of the radiation heat dissipation plate in the deep cold environment. The mechanical unfolding type radiation plate realizes the adjustment of the heat dissipation area through foldable structures, typical such as louver or umbrella unfolding mechanisms, which can realize dynamic regulation of the heat dissipation area, but the complex mechanical drive system faces serious reliability challenges in the space environment, such as mechanism jamming and lubrication failure in microgravity environment, and the integrated design with the heat dissipation pipeline is also relatively complex. The electrochromic coating technology realizes dynamic adjustment of the emissivity based on the characteristics of intelligent materials such as tungsten oxide, which can effectively regulate the radiation heat dissipation efficiency, but the blocking effect of the heat conduction path is limited, and it needs continuous power supply to maintain the working state, which significantly increases the energy consumption burden of the system. The phase change energy storage device technology uses the latent heat characteristics of phase change materials such as paraffin to realize heat buffering, which can delay the cold intrusion process, but has obvious technical limitations: first, a larger volume of phase change material needs to be configured to achieve effective heat buffering effect, which significantly increases the system weight and space occupation; second, the phase change material forms a solid phase heat conduction barrier during the heat dissipation stage, which hinders the effective discharge of heat. This two-way restriction makes the technology face major challenges in actual engineering applications. In addition, other auxiliary technologies such as variable heat conduction structure, thermal switch, etc. also have their own applicability limitations. These technical solutions generally can only realize single-dimensional regulation, lack of systematic and comprehensive solutions, resulting in the need for passive compensation measures such as electric heating for the thermal control system, not only increasing energy consumption, but also failing to fundamentally solve the cold intrusion problem, which seriously restricts the thermal control performance of spacecraft in long-term on-orbit missions.

[0004] Therefore, overcoming the limitations of existing thermal control technologies that rely on single-dimensional regulation, and developing a novel thermal control system capable of simultaneously achieving autonomous opening and closing of heat dissipation channels and dual-path blocking of reverse cooling, is key to resolving the contradiction between passive heat leakage and active energy consumption during spacecraft operation in orbit. This technological breakthrough will significantly improve the environmental adaptability and energy efficiency of spacecraft thermal management systems, providing reliable thermal assurance for long-term on-orbit missions. Summary of the Invention

[0005] The purpose of this invention is to provide an adaptive thermal control system based on the coordinated regulation of pump-driven two-phase flow and thermochromic properties, which aims to solve the problems in the prior art where single-dimensional regulation cannot effectively reduce heat leakage when there is no heat dissipation requirement, and lacks adaptive adjustment capability.

[0006] To achieve the above objectives, the following technical solution is adopted:

[0007] This invention provides an adaptive thermal control system based on the coordinated regulation of pump-driven two-phase flow and thermochromic properties, comprising an adaptive adjustment system for the pump-driven two-phase flow loop and the heat transport path:

[0008] The pump-driven two-phase flow circuit includes a closed-loop circuit consisting of a circulating pump, an evaporator, a condenser, a liquid receiver, and connecting pipelines; the evaporator is attached to the heat source inside the spacecraft and is used to absorb heat to evaporate the working fluid into a gaseous state.

[0009] The adaptive heat transport path adjustment system includes a thermal switch, a condenser, and a radiant heat sink. The thermal switch is installed at the bottom of the condenser and includes a temperature-sensitive deformable material. When the temperature of the condenser rises to a threshold, the temperature-sensitive deformable material expands and deforms, pushing the condenser into contact with the radiant heat sink for heat conduction. When the temperature falls below the threshold, it contracts and resets, causing the radiant heat sink to disengage from the condenser. The radiant heat sink is installed at the top of the condenser and its surface is coated with a thermochromic coating. Its infrared emissivity changes with temperature, switching to a high emissivity state at high temperatures and reverting to a low emissivity state at low temperatures.

[0010] Preferably, the system includes the following operating states:

[0011] Heat dissipation demand activation state: When there is a heat dissipation demand from the heat source inside the spacecraft, the circulation pump starts, the working fluid absorbs heat and vaporizes in the evaporator and is transported to the condenser. The temperature of the condenser rises to the thermal switch trigger threshold, driving the condenser to contact the radiant heat dissipation plate; at the same time, the temperature of the radiant heat dissipation plate rises to the phase change threshold of the thermochromic coating, the thermochromic coating switches to a high infrared emissivity state, and heat is discharged simultaneously through contact heat conduction and high radiation heat dissipation.

[0012] Heat dissipation demand off state: when there is no heat dissipation demand from the internal heat source, the circulating pump stops running, the condenser temperature drops below the heat switch trigger threshold, the heat switch retracts to disengage the condenser from the radiation heat sink, and the radiation heat sink temperature drops below the thermochromic coating phase transition threshold, the thermochromic coating returns to a low infrared emissivity state, blocking the reverse heat conduction path.

[0013] Preferably, the temperature-sensitive deformation material of the heat switch includes any one or a combination of the following:

[0014] Paraffin-based composite: phase transition temperature 10-30℃, volume expansion after heating and melting, pushing the piston or flexible cavity to produce ≥3mm displacement, driving the condenser to contact the radiation heat sink, and retracting to disengage when the temperature decreases;

[0015] Shape memory alloy: phase transition temperature 10-30℃, vertical deformation ≥3mm after austenitic phase transition, driving the condenser to contact the radiation heat sink, and retracting to disengage when the temperature decreases.

[0016] Preferably, the thermochromic coating is a doped modified vanadium dioxide film, with tungsten, molybdenum, niobium or fluorine elements doped to control its phase transition temperature to match the heat switch trigger temperature; the thermochromic coating is uniformly deposited on the surface of the radiation heat sink by magnetron sputtering process.

[0017] Preferably, in the pump-driven two-phase flow circuit, a pressure sensor and a temperature sensor are installed at the outlet of the circulating pump, and temperature sensors are installed at the outlets of the evaporator, condenser and liquid reservoir, respectively.

[0018] Preferably, the condenser adopts a water-cooled plate structure made of titanium alloy, with a heat exchange channel inside and a mechanical connection structure at the bottom for fixed connection with the heat switch, and the top is precisely mechanically polished to reduce contact thermal resistance.

[0019] Preferably, the radiation heat sink adopts a titanium alloy substrate, and the back of the substrate is precisely processed to ensure efficient thermal coupling when in contact with the condenser.

[0020] Preferably, the heat switch is fixedly installed between the bottom of the condenser and the spacecraft structure by bolts, with the deformation direction consistent with the contact direction of the condenser and the radiation heat sink.

[0021] Preferably, the thermochromic coating phase transition threshold is 10-30℃.

[0022] Preferably, the liquid reservoir is used to compensate for the volume change of the working medium and maintain system pressure balance, and its outlet is directly connected with the inlet of the circulating pump.

[0023] The present application has the following remarkable technical effects:

[0024] The present application adopts the dynamic coupling of pump-driven two-phase flow loop, thermal switch and thermochromic coating, realizes the precise regulation and efficient management of the internal thermal environment of the spacecraft. The system adopts an adaptive regulation mechanism constructed by temperature-responsive smart materials, which can quickly establish an efficient heat transfer channel when heat dissipation is required: the thermal switch expands to produce precise displacement, pushing the condenser and the radiation heat sink into close contact; at the same time, the thermochromic coating changes phase to a high-emissivity state, and the two work together to form a complete active heat dissipation path. Under low temperature conditions, the system realizes physical isolation of the heat transfer path through the automatic reset of the thermal switch, and cooperates with the low-emissivity characteristics of the thermochromic coating to form a double barrier against reverse heat transfer in the deep cooling environment, effectively maintaining the thermal balance inside the cabin.

[0025] The regulation mechanism of the system is completely based on the thermal response characteristics of the materials to realize autonomous work, without additional energy input to complete the automatic matching of the heat dissipation power. This self-driven feature not only significantly reduces the dependence on the energy system of the spacecraft, avoiding the energy consumption problem of traditional active thermal control systems, but also eliminates the electromagnetic interference and reliability risks that may be caused by the electric control system, especially suitable for long-term on-orbit mission requirements.

[0026] The system adopts modular design, consisting only of core functional components such as thermal switch, condenser and radiation heat sink, with compact structure and high reliability. This simple and efficient design concept avoids the potential failure risks of complex mechanical transmission mechanisms, and all components are manufactured using space-grade materials and processes, which can adapt to the harsh conditions of temperature alternation and vacuum radiation in space environment, ensuring the long-term stable operation of the system during the mission period. At the same time, the passive regulation characteristics of the system greatly reduce the maintenance requirements, providing a more reliable and durable thermal management solution for spacecraft. BRIEF DESCRIPTION OF DRAWINGS

[0027] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following will briefly introduce the drawings needed to be used in the description of the embodiments. Obviously, the drawings in the following description are only some embodiments of the present application, and other drawings can also be obtained by those skilled in the art without creating any inventive labor. In the drawings:

[0028] Figure 1 Fig. 1 shows a structure diagram of an adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application;

[0029] Figure 2 Fig. 2 shows a heat dissipation mode diagram of an adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application;

[0030] Figure 3 Fig. 1 shows a schematic diagram of a non-radiation mode of a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application;

[0031] Figure 4 Fig. 2 shows a structural diagram of a thermal switch in a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application; wherein (a) is a perspective view of the thermal switch; (b) is a cross-sectional view of the thermal switch;

[0032] Figure 5 Fig. 3 shows a structural diagram of a condenser in a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application;

[0033] Figure 6 Fig. 4 shows a structural diagram of a radiation heat sink in a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation according to an embodiment of the present application.

[0034] Reference signs:

[0035] 1-circulating pump, 2-evaporator, 3-condenser, 301-cold plate, 302-heat exchange pipeline, 4-radiation heat sink, 401-substrate, 402-thermochromic coating, 5-thermal switch, 501-sleeve, 502-phase change material, 503-piston, 6-liquid storage tank. DETAILED DESCRIPTION

[0036] The present application is further explained by the following specific examples, which are not intended to limit the scope of the application. Other objects, aspects and embodiments of the application will occur to those skilled in the art upon consideration of the specification and will be understood to fall within the spirit of the application. The following examples and features of the examples can be combined with each other, if not contradictory, without departing from the spirit of the application.

[0037] The specific embodiments of the present application will be further described in the following with reference to the drawings and examples.

[0038] The present application provides a self-adaptive thermal control system based on pump-driven two-phase flow and thermochromic cooperative regulation, as shown in Figure 1 The pump-driven two-phase flow circuit comprises a circulating pump 1, an evaporator 2, a condenser 3, and a liquid storage tank 6. The outlet of the circulating pump 1 is connected to the inlet of the evaporator 2, the outlet of the evaporator 2 is connected to the inlet of the condenser 3, the outlet of the condenser 3 is connected to the inlet of the liquid storage tank 6, and the outlet of the liquid storage tank 6 is connected to the inlet of the circulating pump 1.

[0039] The heat transport path adaptive adjustment system comprises a thermal switch 5, a condenser 3, and a radiation heat sink 4; the bottom of the thermal switch 5 is connected to the internal structure of the spacecraft, the top of the thermal switch 5 is connected to the bottom of the condenser 3, the top of the condenser 3 is spaced apart from the bottom of the radiation heat sink 4, and there are two working modes; in the heat dissipation mode, the top of the condenser 3 is in contact with the bottom of the radiation heat sink 4, and in the mode without heat dissipation requirement, the top of the condenser 3 is not in contact with the bottom of the radiation heat sink 4; the radiation heat sink 4 is installed on the external structure of the spacecraft.

[0040] The working principle of the system is as follows:

[0041] As shown in Figure 2 , it is a schematic diagram of the heat dissipation mode of the system. In the heat dissipation mode, when the internal heat source of the spacecraft generates heat, the heat is first conducted to the evaporator 2, so that the temperature of the working medium in the evaporator is raised and phase change vaporization occurs. When the evaporator outlet temperature sensor detects that the temperature of the working medium reaches the preset threshold value, the control unit automatically starts the circulating pump 1. The circulating pump 1 drives the liquid working medium from the liquid reservoir 6 to the evaporator 2, and the working medium is converted into a gaseous state after absorbing heat in the evaporator, and then enters the condenser 3. As the temperature of the condenser 3 continues to rise, heat is conducted to the thermal switch 5, and when the temperature reaches the phase change threshold of the thermal switch, the thermal switch 5 produces a deformation displacement, pushing the condenser 3 into close thermal contact with the radiation heat sink 4. At the same time, the temperature of the radiation heat sink 4 rises to the phase change temperature of the thermochromic coating, and the coating switches from a low infrared emissivity state to a high infrared emissivity state. The gaseous working medium releases heat in the condenser 3 and condenses into a liquid state, returns to the liquid reservoir 6 through the pipeline, and completes a complete heat transport cycle. In this process, the system realizes efficient heat dissipation from the spacecraft to space through the synergistic effect of thermal contact conduction and high radiation heat dissipation.

[0042] As shown in Figure 3 , it is a schematic diagram of the system without heat dissipation mode. In the heat dissipation requirement off state, the system executes the heat dissipation mode, and when the internal heat source of the spacecraft stops heating, the temperature of the working medium in the evaporator 2 gradually decreases. When the evaporator outlet temperature sensor detects that the evaporator outlet temperature is lower than the preset threshold value, the circulating pump 1 is automatically turned off. With the interruption of the heat source, the temperature of the condenser 3 continues to decrease, and when the temperature is lower than the phase change threshold of the thermal switch 5, the thermal switch 5 is retracted under the action of the reset mechanism, so that the condenser 3 and the radiation heat sink 4 are completely separated, and the heat conduction path is physically blocked. At the same time, the radiation heat sink 4 loses the heat source input, and the temperature gradually decreases to below the phase change temperature of the thermochromic coating, and the coating returns to a low emissivity state from a high emissivity state. At this time, the system enters a dormant heat preservation state, and through the physical isolation of the thermal switch 5 and the low emissivity characteristics of the thermochromic coating, the reverse heat conduction of the external deep cold environment is effectively blocked, and the internal thermal balance of the spacecraft is maintained.

[0043] In some embodiments, the thermal switch 5 is made of a shape memory alloy material with austenitic and martensitic phase transformation or paraffin wax with solid-liquid phase transformation properties, and its phase transformation temperature point is precisely designed. The thermal switch 5 is fixedly connected between the bottom of the condenser 3 and the spacecraft structure by bolts. When the system temperature exceeds a set threshold (10-30°C), the thermal switch will deform and displace in a predetermined direction, pushing the condenser into contact with the radiant heat sink; when the temperature decreases, the thermal switch automatically resets, separating the two.

[0044] In some embodiments, such as Figure 4 The diagram shown is a structural diagram of a thermal switch 5. The thermal switch 5 includes a sleeve 501, a phase change material 502, and a piston 503, wherein the phase change material 502 is disposed at the bottom of the sleeve 501, and the piston 503 is disposed at the upper end of the phase change material 502.

[0045] In this embodiment, the sleeve 501 is made of stainless steel and has a cylindrical hollow structure. Its inner wall is precision-machined to form a smooth guide surface (roughness Ra≤0.4μm), providing a low-friction channel for the axial movement of the piston 503. The bottom of the sleeve 501 is closed and fixed to the spacecraft structure with bolts. An annular limiting structure is provided at the top opening to prevent the piston 503 from detaching from the sleeve due to excessive displacement. The phase change material 502 fills the closed cavity between the bottom of the sleeve 501 and the piston 503. It is made of paraffin-based composite material or NiTi-based shape memory alloy (phase change temperature 10~30℃), and its volume expands or contracts significantly with temperature changes. The piston 503 is a titanium alloy cylinder. Its lower end is in direct contact with the phase change material 502, and its upper end is rigidly connected to the bottom of the condenser 3 with bolts. It can move axially along the sleeve 501 under the drive of the phase change material 502 (displacement ≥3mm).

[0046] In some embodiments, the condenser 3 adopts a water-cooled plate structure made of titanium alloy and has an internal heat exchange channel; the bottom is provided with a specially designed mechanical connection structure for reliable fixation with the thermal switch; the top of the condenser is precision mechanically polished to obtain extremely high surface flatness, ensuring minimal contact thermal resistance when it is in close contact with the radiant heat sink.

[0047] In some embodiments, such as Figure 5 The diagram shown is a structural diagram of the condenser 3. The condenser 3 includes a cold plate 301 and a heat exchange pipe 302 mounted on the cold plate 301.

[0048] In this embodiment, the cold plate 301 can be manufactured by titanium alloy integrated forming process, and an embedded groove matched with the heat exchange pipeline 302 is designed inside the cold plate 301, so that the heat exchange pipeline 302 and the cold plate 301 are tightly attached in surface contact. The heat exchange pipeline 302 is made of thin-walled stainless steel, and is distributed in a serpentine shape in the embedded groove of the cold plate 301, and the two ends are respectively sealed connected with the inlet and outlet pipelines of the pump-driven two-phase flow circuit through flange structures. When the gaseous working medium is transported from the evaporator 2 to the condenser 3, the working medium first enters the heat exchange pipeline 302, and then the heat exchange pipeline 302 and the cold plate 301 perform forced convection heat exchange through the pipeline wall surface, so that the heat is transferred to the cold plate 301; the cold plate 301 as an intermediate heat sink realizes secondary heat transfer through the polished surface at the top (when in contact with the radiation heat dissipation plate 4) or the thermal switch 5 at the bottom (when connected with the spacecraft structure).

[0049] In some embodiments, the radiation heat dissipation plate 4 is made of a titanium alloy substrate, and a VO2 thermochromic film doped with tungsten / molybdenum is plated on the surface, and the phase transition temperature is matched with the 10-30℃ trigger range of the thermal switch 5. The back surface of the substrate is precisely processed to ensure ultra-high flatness, and high-efficiency heat coupling with the condenser 3 is realized. This design makes the emissivity of the heat dissipation plate automatically adjustable with temperature, and the heat dissipation plate is in a high-infrared-emissivity state at high temperature (> 30℃), and is in a low-infrared-emissivity state at low temperature (< 10℃).

[0050] In some embodiments, as shown in FIG. 4, the radiation heat dissipation plate 4 is a structure diagram. The radiation heat dissipation plate 4 includes a substrate 401 and a thermochromic coating 402 covering a surface of the substrate 401. Figure 6

[0051] In this embodiment, the substrate 401 can be made of titanium alloy material, and is precisely machined to form a flat rectangular plate structure, and the thickness is designed to be 3-5 mm, which not only meets the lightweight requirement of the spacecraft structure, but also ensures good structural stability and heat conduction performance. One surface of the substrate 401 (the side facing the space) is used as the bearing surface of the thermochromic coating 402, and the surface is polished to control the roughness to be Ra≤0.8μm, which provides a smooth substrate for uniform coating of the coating; the other surface (the side facing the condenser 3) is also precisely processed to ensure that high-efficiency heat coupling can be realized when in contact with the condenser 3, and the contact thermal resistance is reduced.

[0052] The thermochromic coating 402 is a doped modified vanadium dioxide (VO2) film, which is uniformly deposited on the bearing surface of the substrate 401 by a magnetron sputtering process, and the coating thickness is controlled to be 500-800nm. By doping tungsten, molybdenum, niobium or fluorine elements, the phase transition temperature can be accurately controlled to 10-30℃, which is matched with the trigger temperature of the thermal switch 5, so that the two work cooperatively.

[0053] ​The above embodiments are only used for illustrating the present application, and are not intended to limit the present application. Those skilled in the art can make various changes and modifications without departing from the spirit and scope of the present application. Therefore, all equivalent technical solutions belong to the scope of the present application, and the patent protection scope of the present application should be defined by the claims.

Claims

1. A self-adaptive thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic materials, characterized in that, The system comprises a pump-driven two-phase flow loop and a heat transport path adaptive adjustment system. The pump-driven two-phase flow loop comprises a closed loop formed by a circulating pump (1), an evaporator (2), a condenser (3), a liquid reservoir (6) and connecting pipelines; the evaporator (2) is attached to an internal heat source of a spacecraft and is used to absorb heat to vaporize the working medium into a gaseous state. The heat transport path adaptive adjustment system comprises a thermal switch (5), the condenser (3) and a radiation heat sink (4); the thermal switch (5) is installed at the bottom of the condenser (3); the thermal switch (5) comprises a temperature-sensitive deformation material; when the temperature of the condenser (3) rises to a threshold value, the temperature-sensitive deformation material expands and deforms, pushes the condenser (3) into contact with the radiation heat sink (4) to conduct heat; when the temperature is lower than the threshold value, the temperature-sensitive deformation material shrinks to reset the condenser (3) to be separated from the radiation heat sink (4); the radiation heat sink (4) is installed at the top of the condenser (3); the surface of the radiation heat sink (4) is coated with a thermochromic coating, the infrared emissivity of which changes with temperature; when the temperature is high, the thermochromic coating switches to a high emissivity state; when the temperature is low, the thermochromic coating returns to a low emissivity state.

2. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The system comprises the following working states: Heat dissipation demand activation state: when there is a heat dissipation demand from an internal heat source of a spacecraft, the circulating pump (1) is started, the working medium is vaporized in the evaporator (2) by absorbing heat and is transported to the condenser (3), the temperature of the condenser (3) rises to the trigger threshold value of the thermal switch (5), which drives the condenser (3) into contact with the radiation heat sink (4); at the same time, the temperature of the radiation heat sink (4) rises to the phase transition threshold value of the thermochromic coating, the thermochromic coating switches to a high infrared emissivity state, and heat is simultaneously discharged through contact heat conduction and high radiation heat dissipation; Heat dissipation demand shutdown state: when there is no heat dissipation demand from an internal heat source, the circulating pump (1) stops running, the temperature of the condenser (3) drops below the trigger threshold value of the thermal switch (5), the thermal switch (5) shrinks to reset the condenser (3) to be separated from the radiation heat sink (4); the temperature of the radiation heat sink (4) drops below the phase transition threshold value of the thermochromic coating, the thermochromic coating returns to a low infrared emissivity state, and the reverse heat conduction path is blocked.

3. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The temperature-sensitive deformation material of the thermal switch (5) comprises any one or a combination of the following: Paraffin-based composite: phase transition temperature 10-30℃, volume expansion after melting, pushing the piston or flexible cavity to produce ≥3mm displacement, driving the condenser (3) into contact with the radiation heat sink (4), and shrinking to reset to separate when the temperature decreases; Shape memory alloy: phase transition temperature 10-30℃, vertical deformation amount ≥3mm after austenite phase transition, driving the condenser (3) into contact with the radiation heat sink (4), and external force compression reset to separate when the temperature decreases.

4. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The thermochromic coating is a doped modified vanadium dioxide film, the phase transition temperature of which is adjusted by doping tungsten, molybdenum, niobium or fluorine elements to match the trigger temperature of the thermal switch (5); the thermochromic coating is uniformly deposited on the surface of the radiation heat sink (4) by a magnetron sputtering process.

5. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The pump drives two-phase flow circuit, the circulating pump (1) outlet is provided with pressure sensor and temperature sensor, the evaporator (2) outlet, the condenser (3) outlet and the liquid reservoir (6) outlet are respectively provided with temperature sensor.

6. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The condenser (3) adopts water-cooled plate structure of titanium alloy material, is provided with heat exchange flow channel inside, is provided with mechanical connection structure for fixed connection with the thermal switch (5) at the bottom, and is subjected to precise mechanical polishing treatment at the top to reduce contact thermal resistance.

7. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 6, characterized in that, The radiation heat sink (4) adopts titanium alloy substrate, and the back surface of the substrate is subjected to precise machining to ensure high efficient thermal coupling when contacting with the condenser (3).

8. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The thermal switch (5) is fixedly installed between the bottom of the condenser (3) and the spacecraft structure by bolts, and the deformation direction is consistent with the contact direction of the condenser (3) and the radiation heat sink (4).

9. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 2, characterized in that, The thermochromic coating phase transition threshold is 10-30 DEG C.

10. The self-adapting thermal control system based on synergistic regulation of pump-driven two-phase flow and thermochromic according to claim 1, characterized in that, The liquid reservoir (6) is used for compensating the volume change of working medium and maintaining system pressure balance, and the outlet thereof is directly connected with the inlet of the circulating pump (1).

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