A main passive coupled temperature control structure for a Mars aircraft

Abstract

A main-passive coupled temperature control structure for a Mars aircraft, which relates to a temperature control structure for a Mars aircraft. The present invention aims to solve the problems that existing Mars aircraft cannot meet the temperature limits of different temperature control modes in variable temperature control scenarios and cannot minimize temperature control energy to the greatest extent. The present invention includes a fuselage box, an electrical core component, an interaction execution component, and a photovoltaic power generation component; the electrical core component is installed inside the fuselage box, the interaction execution component is installed on the outer surface of the fuselage box, the photovoltaic power generation component is installed on the top of the fuselage box, and the interaction execution component and the photovoltaic power generation component are electrically connected to the electrical core component. The present invention belongs to the technical field of thermal management of aerospace aircraft.

Description

Technical Field

[0001] The present invention relates to a temperature control structure of a Mars spacecraft, belonging to the technical field of thermal management of aerospace vehicles. Background Art

[0002] Mars is Earth's neighbor in the solar system, sharing a similar physical size and topography. Mars exploration is crucial for expanding human living space and exploring the origins of life. With its unique high flight speed, wide exploration range, localized exploration capabilities, and pinpoint landing capabilities, Mars spacecraft are positioned as aerial exploration platforms to assist Mars rovers in efficiently completing their missions, offering a wide range of applications.

[0003] Temperature control systems are fundamental to the survival and mission execution of Martian equipment. As small mechanical devices, the Martian spacecraft's low-power mode leads to uncertainty in the establishment of thermal models, and it is a typical low-temperature system. The Martian spacecraft uses secondary batteries to store and release the limited production capacity of solar cells in the Martian environment. Its temperature control structure design must be lightweight and energy-efficient to achieve self-sustaining balance. The Martian spacecraft is required to survive, communicate, image, fly, move, and sample in the extreme Martian environment. The inherent ratio of its thermal inertia to its heat-exposed area makes achieving these operations more complex and challenging than in the Earth environment.

[0004] Compared to other terrestrial Martian equipment, the design of temperature control structures for Martian spacecraft, which are extremely sensitive to their own weight, has not been extensively studied or experimentally verified. Therefore, developing a coupled active and passive temperature control structure for Martian spacecraft is crucial to ensure that the spacecraft meets the temperature limits of different temperature control modes in variable temperature control scenarios while minimizing temperature control energy. Summary of the Invention

[0005] The present invention aims to solve the problem that existing Mars spacecraft cannot meet the temperature limits of different temperature control modes in variable temperature control scenarios and cannot minimize temperature control energy, and thus proposes an active-passive coupled temperature control structure for Mars spacecraft.

[0006] The technical solution adopted by the present invention to solve the above problems is: the present invention includes a fuselage box, an electrical core component, an interactive execution component and a photovoltaic power generation assembly; the electrical core component is installed in the fuselage box, the interactive execution component is installed on the outer surface of the fuselage box, the photovoltaic power generation assembly is installed on the top of the fuselage box, and the interactive execution component and the photovoltaic power generation assembly are electrically connected to the electrical core component.

[0007] Further, the fuselage box body includes a central load-bearing pipe, two structural interface plates, an electrical core support, a fuselage skin support, and a fuselage skin; the central load-bearing pipe is vertically arranged inside the fuselage skin support, one structural interface plate is fixedly sleeved on the upper end of the central load-bearing pipe, the other structural interface plate is fixedly sleeved on the lower end of the central load-bearing pipe, the electrical core support is fixedly sleeved in the middle of the central load-bearing pipe, the fuselage skin is sleeved on the fuselage skin support, and a ventilation window is opened on the fuselage skin.

[0008] Further, the electrical core components include a battery assembly, a core circuit assembly, and an external sensor assembly; both the battery assembly and the external sensor assembly are electrically connected to the core circuit assembly.

[0009] Further, the battery assembly includes a plurality of battery cells, cell straps, a first temperature detector, a first heater, and a cell skin; the plurality of battery cells are arranged around the central load-bearing pipe in the circumferential direction, the plurality of battery cells are fixedly connected together through the cell straps, the cell straps are connected to the electrical core support, and the first temperature detector, the first heater, and the cell skin are provided on the battery cells.

[0010] Further, the core circuit assembly includes at least three core PCBs, PCB connection blocks, a second temperature detector, and a second heater; the core PCBs are evenly arranged around the battery assembly, the surface of the core PCBs is coated with an electrical core skin, the core PCBs are connected to the electrical core support through the PCB connection blocks, and the second temperature detector and the second heater are provided on the core PCBs.

[0011] Further, the external sensor assembly includes an external sensor, a third temperature detector, a third heater, and an external sensor support; the external sensor support is fixedly installed on the fuselage box body, and the external sensor, the third temperature detector, and the third heater are all fixedly installed on the external sensor support.

[0012] Further, the interaction execution components include a plurality of structural arms, an actuator, a high-power unit, a low-power unit, a fourth temperature detector, and a fourth heater; the high-power unit is fixedly connected to the structural interface plate through the structural arms, the actuator is fixedly connected to the structural interface plate through the structural arms, the fourth temperature detector and the fourth heater are provided on the actuator, the low-power unit is installed on the actuator, the high-power unit is composed of a high-power motor and a high-power PCB, and the low-power unit is composed of a low-power motor and a low-power PCB.

[0013] Further, the photovoltaic power generation component includes a photovoltaic circuit and a photovoltaic substrate; the photovoltaic circuit is fixed on the photovoltaic substrate, and a shunt electric heater and a shunt radiator are provided on the lower surface of the photovoltaic substrate.

[0014] The beneficial effects of the present invention are:

[0015] 1. The present invention establishes a closed semi-monocoque fuselage box component through a central load-bearing pipe, a structural interface plate, a fuselage skin bracket, and a fuselage skin, isolating the Mars atmospheric environment from the electrical core components. At the same time, ventilation windows are provided on the fuselage skin to introduce gases with low thermal conductivity in the Mars atmosphere, and an insulating air gap is established in the gap at a certain distance between the fuselage skin and the cell skin and the electrical core skin. The thermal isolation solution implemented by this temperature control structure reduces the heat leakage generated by the electrical core components through solid heat conduction, gas heat conduction, gas heat convection, and blackbody thermal radiation at an extremely low mass cost, reduces the influence of space radiation, static electricity dust, and ground particles on the electrical core components, and improves the low-temperature survival ability and working life of the Mars aircraft;

[0016] 2. The present invention fully tensions the fuselage skin by the fuselage skin bracket or performs co-curing treatment with the fuselage skin bracket to reduce the gas volume change within the range covered by the skin during the operation of the aircraft, reduce the gas flow level in the insulating air gap, increase the thermal resistance of the insulating air gap, and reduce the heat loss caused by the pumping out of warm gas in the insulating air gap;

[0017] 3. The present invention arranges the electrical core components necessary to maintain flight survival inside the fuselage box component, and arranges the photovoltaic power generation components and motion operation components that can survive in the low-temperature environment of Mars in the non-working state outside the fuselage box component, reducing the temperature control objects of the Mars aircraft at night and reducing the required survival energy consumption;

[0018] 4. The present invention maintains the minimum temperature of the battery pack at a healthy level during the silent, dormant, standby, and working states of the Mars aircraft by setting a first temperature detector and a first heater in the cell, and can selectively increase or decrease the cell temperature at a specified working stage to obtain the required battery pack working ability;

[0019] 5. The present invention consolidates the cells with cell straps, reduces the temperature gradient between the cells, makes the temperatures of the cells close to equal, obtains similar discharge capabilities, improves the low-temperature stability of the battery pack, and stabilizes the discharge performance of the battery pack at different temperature states;

[0020] 6. The present invention consolidates the cells with cell straps, reduces the temperature inside the cells, makes the internal temperature distribution of the cells uniform, and makes the aging of each part of the cells uniform, improving the service life of the cells;

[0021] 7. The present invention connects the central load-bearing pipe and the electrical core module through the electrical core bracket, reduces the heat leakage caused by the electrical core module through the central load-bearing pipe, and improves the low-temperature survival ability of the Mars aircraft;

[0022] 8. By deploying the core PCB around the battery pack, the present invention absorbs the heat leaked from the battery pack to increase the temperature of the core PCB, improves the utilization rate of temperature control energy, reduces the survival energy required by the Mars aircraft at night, and reduces the preheating energy required for the Mars aircraft to work during the day;

[0023] 9. By respectively arranging the second, third, fourth, fifth, and sixth temperature sensors and the second, third, fourth, fifth, and sixth heaters on the core PCB, the external sensor assembly, the actuator, the high-power PCB, and the low-power PCB, the present invention actively heats the specified module to be in an available state when the external environment cannot naturally heat the aircraft, broadens the range of operable environmental states of the aircraft, and expands the Mars longitude, latitude, and altitude areas where the aircraft can be deployed; improves the flexibility of the aircraft to perform tasks;

[0024] 10. By arranging a winding heat storage member and a housing heat dissipation member on the high-power motor, the present invention reduces the heating rate of the high-power motor during operation, extends the maximum single working duration of the Mars aircraft, reduces the overheating risk of the high-power motor, and improves the high-temperature safety margin of the aircraft;

[0025] 11. By arranging a housing heat absorption member on the low-power motor and an insulating spacer between the low-power PCB and the low-power motor stator, the present invention reduces the risk of low-temperature flameout of the low-power motor caused by gusts or shadows during operation, and improves the low-temperature safety margin of the aircraft;

[0026] 12. By arranging a shunt electrothermal unit on the photovoltaic substrate, the present invention converts the excess electric energy of the photovoltaic circuit into heat energy through a shunt electrothermal heater, and then dissipates the heat outside the aircraft through a shunt radiator, improves the reliability and safety of the aircraft's photovoltaic power generation components, and reduces the overheating risk of the electrical core components of the Mars aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] [[ID=ID=17]] Figure 1 is the overall structural schematic diagram of the present invention;

[0028] Figure 2 is the structural schematic diagram of the fuselage box component;

[0029] Figure 3 is the structural schematic diagram of the operation execution component;

[0030] Figure 4 is the structural schematic diagram of the battery assembly and the core circuit assembly;

[0031] Figure 5 is the structural schematic diagram of the external sensor assembly;

[0032] Figure 6 is the structural schematic diagram of the high-power unit;

[0033] Figure 7 It is a schematic diagram of the low-power component structure;

[0034] Figure 8 It is a schematic diagram of the actuator structure;

[0035] Figure 9 It is a schematic diagram of the photovoltaic power generation component structure. Specific implementation manners

[0036] Specific implementation manner one: Combining Figures 1 to 9 to illustrate this implementation manner, a main passive coupled temperature control structure of a Mars aircraft described in this implementation manner includes a fuselage box 1100, an electrical core component 1200, an interaction execution component 1300, and a photovoltaic power generation component 1400; the electrical core component 1200 is installed in the fuselage box 1100, the interaction execution component 1300 is installed on the outer surface of the fuselage box 1100, the photovoltaic power generation component 1400 is installed on the top of the fuselage box 1100, and the interaction execution component 1300 and the photovoltaic power generation component 1400 are electrically connected to the electrical core component 1200.

[0037] Specific implementation manner two: Combining Figure 1 and Figure 2 to illustrate this implementation manner, the fuselage box 1100 of a main passive coupled temperature control structure of a Mars aircraft described in this implementation manner includes a central load-bearing pipe 1110, two structural interface plates 1120, an electrical core support 1130, a fuselage skin support 1140, and a fuselage skin 1150; the central load-bearing pipe 1110 is vertically arranged in the fuselage skin support 1140, one structural interface plate 1120 is fixedly sleeved on the upper end of the central load-bearing pipe 1110, the other structural interface plate 1120 is fixedly sleeved on the lower end of the central load-bearing pipe 1110, the electrical core support 1130 is fixedly sleeved in the middle of the central load-bearing pipe 1110, the fuselage skin 1150 is sleeved on the fuselage skin support 1140, and an air vent 1160 is opened on the fuselage skin 1150.

[0038] The central load-bearing pipe 1110, the structural interface plate 1120, and the electrical core support 1130 are made of materials with low thermal conductivity and high specific strength. Approximate thermal expansion rate treatment is performed at the mechanical connection between the structural interface plate 1120 and the electrical core support 1130 and the central load-bearing pipe 1110. The central load-bearing pipe 1110, the structural interface plate 1120, and the electrical core support 1130 adopt a hollow or hollowed-out structure;

[0039] At the place where the central load-bearing pipe 1110 and the structural interface plate 1120 pass through the fuselage skin 1150, the hollow structure is filled with materials with low thermal conductivity and low density;

[0040] The outer surfaces of the central load-bearing tube 1110, the structural interface plate 1120, and the electrical core support 1130 within the coverage of the fuselage skin 1150 are coated with a coating having a high infrared reflectivity and a low visible light absorptivity, and the inner surfaces are coated with a coating having a low infrared reflectivity and a high visible light absorptivity;

[0041] The central load-bearing tube 1110 and the structural interface plate 1120 are made of hollow structural members of carbon fiber / cyanate resin composite material with a thermal conductivity lower than 4 W / mK. The electrical core support 1130 is a hollowed-out support made of titanium alloy material. The mechanical connection method between the structural interface plate 1120 and the electrical core support 1130 and the central load-bearing tube 1110 is adhesive bonding;

[0042] At the locations where the central load-bearing tube 1110 and the structural interface plate 1120 pass through the fuselage skin 1150, the hollow structure is filled with rigid cyanate foam;

[0043] The outer surfaces of the central load-bearing tube 1110, the structural interface plate 1120, and the electrical core support 1130 within the coverage of the fuselage skin 1150 are coated with an S-13-GP / LO-1 coating, and the inner surfaces are coated with an Al / Blk-Cr coating;

[0044] The fuselage skin supports 1140 are distributed around the electrical core support 1130 and maintain a certain distance air gap. The outer side of the fuselage skin supports 1140 is covered with a fuselage skin 1150 provided with ventilation windows 1160. The fuselage skin 1150 is fully tensioned by the fuselage skin supports 1140 or undergoes co-curing treatment with the fuselage skin supports 1140;

[0045] The fuselage skin supports 1140 are cables or struts made of materials with low thermal conductivity and high specific strength. The gap distance is large enough but not large enough to introduce gas convection. The fuselage skin 1150 is a lightweight film made of materials with low thermal conductivity and high specific strength;

[0046] The outer surface of the fuselage skin 1150 is coated with a coating having a low infrared reflectivity and a high visible light absorptivity, and the inner surface is coated with a coating having a high infrared reflectivity and a low visible light absorptivity;

[0047] The fuselage skin supports 1140 are struts made of carbon fiber / cyanate resin composite material. The air gap distance is 300 mm. The fuselage skin 1150 is a lightweight PI film made of polyimide. The fuselage skin 1150 undergoes co-curing treatment with the fuselage skin supports 1140;

[0048] The outer surface of the fuselage skin 1150 is coated with an Al / Blk-Cr coating, and the inner surface is coated with an Au coating.

[0049] The other components and connection relationships are the same as those in the first specific embodiment.

[0050] Specific embodiment three: Figure 4 To illustrate this embodiment, the electrical core component 1200 of the active-passive coupled temperature control structure of a Mars aircraft in this embodiment includes a battery assembly 1210, a core circuit assembly 1220, and an external sensor assembly 1230; both the battery assembly 1210 and the external sensor assembly 1230 are electrically connected to the core circuit assembly 1220. The other components and connection relationships are the same as those in the first specific embodiment.

[0051] Specific embodiment four: Figure 4 To illustrate this embodiment, the battery assembly 1210 of the active-passive coupled temperature control structure of a Mars aircraft in this embodiment includes a plurality of battery cells 1211, a battery cell strap 1212, a first temperature detector 1213, a first heater 1214, and a battery cell skin 1215; the plurality of battery cells 1211 are arranged around the central load-bearing tube 1110 in the circumferential direction, and the plurality of battery cells 1211 are fixedly connected together through the battery cell strap 1212. The battery cell strap 1212 is connected to the electrical core support 1130. The first temperature detector 1213, the first heater 1214, and the battery cell skin 1215 are provided on the battery cell 1211.

[0052] The battery cell 1211 uses a low-temperature battery cell with excellent battery-to-battery reproducibility, high specific energy, and radiation resistance. The battery cell strap 1212 is one or more isothermal straps made of a material with high specific strength and high thermal conductivity and connected to the surface area of the battery cell 1211 where high temperatures are characterized during operation.

[0053] The first temperature detector 1213 is composed of a plurality of temperature sensors in parallel that are radiation-resistant, have a measurement range including -135°C to 50°C, have a redundant detection circuit, and have a certain accuracy. The first heater 1214 is composed of a plurality of electric heaters in parallel that have sufficient power, can achieve large-area heating, have a redundant power supply circuit, and can adjust the heating power by controlling the duty cycle. The battery cell skin 1215 is a lightweight film made of a material with low thermal conductivity and low density.

[0054] Coatings with low infrared reflectivity and high visible light absorption rate are provided on the inner and outer surfaces of the battery cell skin 1215.

[0055] The battery cell 1211 uses 18650 battery cells of Sony-VTC6. The battery cell strap 1212 is composed of two isothermal straps made of 7075 aluminum alloy. The mechanical connection method between the battery cell 1211 and the battery cell strap 1212 is adhesive bonding. The mechanical connection method between the electrical core support 1130 and the battery cell strap 1212 is threaded connection.

[0056] The first temperature detector 1213 is composed of several PRTs with two sets of detection circuits connected in parallel. The first heater 1214 is composed of several Kapton film heaters with two sets of power supply circuits connected in parallel. The first temperature detector 1213 and the first heater 1214 are coated on the side of the battery cell 1211. The battery cell skin 1215 is a lightweight PI film made of polyimide.

[0057] Au coatings are provided on the inner and outer surfaces of the battery cell skin 1215.

[0058] Other components and connection relationships are the same as those in the second or third specific embodiment.

[0059] Specific embodiment five: With reference to Figure 4 To illustrate this embodiment, the core circuit component 1220 of the active and passive coupled temperature control structure of a Mars aircraft includes at least three core PCBs 1221, a PCB connection block 1222, a second temperature detector 1223, and a second heater 1224. The core PCBs 1221 are evenly arranged around the battery component 1210. The surface of the core PCB 1221 is coated with an electrical core skin 1225. The core PCB 1221 is connected to the electrical core support 1130 through the PCB connection block 1222. The second temperature detector 1223 and the second heater 1224 are provided on the core PCB 1221.

[0060] The core PCBs 1221 are distributed around the battery pack and maintain a certain distance air gap.

[0061] The core PCB 1221 is made of a material with high specific strength and suitable for vacuum. The PCB connection block 1222 is made of a material with high specific strength and low thermal conductivity.

[0062] The second temperature detector 1223 is composed of multiple temperature sensors with radiation resistance, a measurement range including -135°C to 50°C, redundant detection circuits, and a certain accuracy, connected in parallel. The second heater 1224 is composed of multiple electric heaters with sufficient power, capable of realizing local area heating, integrable to the PCB, and adjustable heating power by controlling the duty cycle, connected in parallel. The air gap distance is large enough but not large enough to introduce gas convection. The electrical core skin 1225 is a lightweight film made of a material with low thermal conductivity and low density.

[0063] The mechanical connection method of the PCB connection block 1222 with the core PCB 1221 and the electrical core support 1130 is threaded connection.

[0064] The core PCB 1221 is made of FR4, and the PCB connection block 1222 is made of PEEK.

[0065] The second temperature detector 1223 is composed of several PRTs with two sets of detection circuits connected in parallel. The second heater 1224 is composed of several winding heaters connected in parallel. The second temperature detector 1223 and the second heater 1224 are arranged around the components that need to be heated additionally when the core PCB 1221 is working. The electrical core skin 1225 is a lightweight PI film made of polyimide.

[0066] Other compositions and connection relationships are the same as those in the second or third specific implementation manner.

[0067] Specific implementation manner six: Figures 4 to 8 This specific implementation manner will be described in combination with the following content. The external sensor assembly 1230 of the active-passive coupled temperature control structure of a Mars aircraft in this specific implementation manner includes an external sensor 1231, a third temperature detector 1232, a third heater 1233, and an external sensor bracket 1234. The external sensor bracket 1234 is fixedly installed on the fuselage box 1100, and the external sensor 1231, the third temperature detector 1232, and the third heater 1233 are all fixedly installed on the external sensor bracket 1234.

[0068] The third temperature detector 1232 is composed of several temperature sensors with radiation resistance, a measurement range including -135°C to 50°C, redundant detection circuits, and a certain accuracy, connected in parallel. The third heater 1233 is composed of several electric heaters with sufficient power, capable of realizing local area heating, integrable to the PCB, and adjustable heating power by controlling the duty cycle, connected in parallel.

[0069] The third temperature detector 1232 is composed of several PRTs with two sets of detection circuits connected in parallel. The third heater 1233 is composed of several winding heaters connected in parallel. The third temperature detector 1232 and the third heater 1233 are arranged around the components that need to be heated additionally when the external sensor assembly 1230 is working.

[0070] Other compositions and connection relationships are the same as those in the first or third specific implementation manner.

[0071] Specific implementation manner seven: Figures 3 to 8To describe this embodiment, the interactive execution component 1300 of the active-passive coupled temperature control structure of a Mars aircraft includes a plurality of structural support arms 1310, an actuator 1320, a high-power unit 1330, a low-power unit 1340, a fourth temperature detector 1321, and a fourth heater 1322; the high-power unit 1330 is fixedly connected to the structural interface board 1120 through the structural support arm 1310, the actuator 1320 is fixedly connected to the structural interface board 1120 through the structural support arm 1310, the fourth temperature detector 1321 and the fourth heater 1322 are provided on the actuator 1320, the low-power unit 1340 is installed on the actuator 1320, the high-power unit 1330 is composed of a high-power motor 1331 and a high-power PCB 1332, and the low-power unit 1340 is composed of a low-power motor 1341 and a low-power PCB 1342.

[0072] The structural support arm 1310 is a hollow structural member made of a material with low thermal conductivity and high specific strength. The mechanical connection between the structural interface board 1120 and the structural support arm 1310 is treated with an approximate coefficient of thermal expansion. The hollow structure is filled with a material with low thermal conductivity and low density at the mechanical connections between the structural support arm 1310 and the structural interface board 1120 and the actuator 1320.

[0073] The fourth temperature detector 1321 is composed of a plurality of temperature sensors with radiation resistance, a measurement range including -135°C to 50°C, a redundant detection circuit, and a certain accuracy, connected in parallel. The fourth heater 1322 is composed of a plurality of electric heaters with sufficient power, capable of realizing local area heating, integratable to the PCB, and adjustable in heating power by controlling the duty cycle, connected in parallel.

[0074] The substrates of the high-power PCB 1332 and the low-power PCB 1342 are made of a material with high strength and suitable for vacuum.

[0075] The hollow structure is filled with rigid cyanate ester foam at the mechanical connections between the structural support arm 1310 and the structural interface board 1120 and the actuator 1320.

[0076] The fourth temperature detector 1321 is composed of several PRTs with two sets of detection circuits connected in parallel. The fourth heater 1322 is composed of several winding heaters connected in parallel. The fourth temperature detector 1321 is arranged near the output shaft housing of the actuator 1320, and the fourth heater 1322 is arranged near the input shaft housing of the actuator 1320.

[0077] The substrates of the high-power PCB 1332 and the low-power PCB 1342 are made of FR4.

[0078] The high-power motor 1331 is provided with a winding heat storage member 1331-1 and a housing heat dissipation member 1331-2. The winding heat storage member 1331-1 is mechanically connected to the winding, and the housing heat dissipation member 1331-2 is mechanically connected to the housing. The high-power PCB 1332 is provided with a fifth temperature detector 1332-1 and a fifth heater 1332-2. The low-power motor 1341 is provided with a housing heat absorption member 1341-1, and the housing heat absorption member 1341-1 is mechanically connected to the housing. The low-power PCB 1342 is provided with a sixth temperature detector 1342-1 and a sixth heater 1342-2;

[0079] The winding heat storage member 1331-1 is made of a material with high specific heat, high specific strength, and high thermal conductivity. The housing heat dissipation member 1331-2 is made of a material with high infrared reflectivity and low visible light absorption rate. The housing heat absorption member 1341-1 is made of a material with low infrared reflectivity and high visible light absorption rate;

[0080] The fifth temperature detector 1332-1 and the sixth temperature detector 1342-1 are composed of multiple temperature sensors in parallel that are radiation-resistant, have a measurement range including -135°C to 50°C, have a redundant detection circuit, and have a certain accuracy. The fifth heater 1332-2 and the sixth heater 1342-2 are composed of multiple electric heaters in parallel that have sufficient power, can achieve local area heating, can be integrated into the PCB, and can adjust the heating power by controlling the duty cycle;

[0081] The connection methods of the winding heat storage member 1331-1 to the winding, the housing heat dissipation member 1331-2 to the housing, and the housing heat absorption member to the housing are adhesive bonding. The winding heat storage member 1331-1 is made of AlBeMet aluminum-beryllium alloy, the housing heat dissipation member 1331-2 is made of silver-plated Teflon, and the housing heat absorption member 1341-1 is made of Al / Blk-Cr;

[0082] The fifth temperature detector 1332-1 and the sixth temperature detector 1342-1 are composed of several PRTs with two sets of detection circuits in parallel. The fifth heater 1332-2 and the sixth heater 1342-2 are composed of several winding heaters in parallel. The fifth temperature detector 1332-1 and the fifth heater 1332-2 are arranged around the components that need additional temperature rise when the high-power PCB 1332 is working. The sixth temperature detector 1342-1 and the sixth heater 1342-2 are arranged around the components that need additional temperature rise when the low-power PCB 1342 is working;

[0083] The high-power PCB 1332 is connected to the stator of the high-power motor 1331 through a heat insulation spacer 1332-3 and is electrically connected to its winding. The low-power PCB 1342 is connected to the stator of the low-power motor 1341 through a cold insulation spacer 1342-3 and is electrically connected to its winding;

[0084] The heat insulation spacer 1332-3 and the cold insulation spacer 1342-3 are cylindrical cylinders with a certain height made of materials with high specific strength and low thermal conductivity;

[0085] The heat insulation spacer 1332-3 and the cold insulation spacer 1342-3 are cylindrical cylinders made of PEEK.

[0086] Other compositions and connection relationships are the same as those in the first or second specific embodiment.

[0087] Specific embodiment eight: In combination with Figure 9 To illustrate this embodiment, the photovoltaic power generation component 1400 of the active and passive coupled temperature control structure of a Mars aircraft described in this embodiment includes a photovoltaic circuit 1410 and a photovoltaic substrate 1420; the photovoltaic circuit 1410 is fixed on the photovoltaic substrate 1420, and a shunt electric heater 1421 and a shunt radiator 1422 are provided on the lower surface of the photovoltaic substrate 1420.

[0088] The photovoltaic substrate 1420 is made of materials with high specific strength, high specific stiffness, and low thermal conductivity. The shunt electric heater 1421 is composed of a plurality of electric heaters connected in parallel with sufficient power, capable of realizing large-area heating, having a redundant power supply circuit, and adjustable heating power by controlling the duty cycle. The shunt radiator 1422 is made of materials with high infrared reflectivity and low visible light absorption rate;

[0089] The shunt electric heater 1421 is composed of several Kapton film heaters with two sets of power supply circuits connected in parallel. The shunt radiator 1422 is made of silver-plated Teflon.

[0090] Other compositions and connection relationships are the same as those in the first specific embodiment.

[0091] The above are only the preferred embodiments of the present invention, and do not impose any form of limitation on the present invention. Although the present invention has been disclosed above with the preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make some changes or modifications to the above-disclosed technical content to make equivalent embodiments with equivalent changes. However, as long as the technical content of the present invention is not departed from, and based on the technical essence of the present invention, any simple modification, equivalent replacement, and improvement of the above embodiments within the spirit and principles of the present invention still fall within the protection scope of the technical solution of the present invention.

Claims

1. A main passive coupled temperature control structure for a Mars aircraft, characterized in that: The described active-passive coupled temperature control structure of a Mars aircraft includes a fuselage box (1100), an electrical core component (1200), an interaction execution component (1300), and a photovoltaic power generation component (1400); the electrical core component (1200) is installed inside the fuselage box (1100), the interaction execution component (1300) is installed on the outer surface of the fuselage box (1100), the photovoltaic power generation component (1400) is installed on the top of the fuselage box (1100), and the interaction execution component (1300) and the photovoltaic power generation component (1400) are electrically connected to the electrical core component (1200). The fuselage box (1100) includes a central load-bearing tube (1110), two structural interface plates (1120), an electrical core support (1130), a fuselage skin support (1140), and a fuselage skin (1150); the central load-bearing tube (1110) is vertically arranged inside the fuselage skin support (1140), one structural interface plate (1120) is fixedly sleeved on the upper end of the central load-bearing tube (1110), the other structural interface plate (1120) is fixedly sleeved on the lower end of the central load-bearing tube (1110), the electrical core support (1130) is fixedly sleeved in the middle of the central load-bearing tube (1110), the fuselage skin (1150) is sleeved on the fuselage skin support (1140), and a ventilation window (1160) is opened on the fuselage skin (1150).

2. The main and passive coupled temperature control structure of a Mars aircraft according to claim 1, characterized in that: The electrical core component (1200) includes a battery component (1210), a core circuit component (1220), and an external sensor component (1230); the battery component (1210) and the external sensor component (1230) are both electrically connected to the core circuit component (1220).

3. The main passive coupling type temperature control structure of a Mars aircraft according to claim 1 or 2, characterized in that: The battery component (1210) includes a plurality of battery cells (1211), a battery cell band (1212), a first temperature sensor (1213), a first heater (1214), and a battery cell skin (1215); the plurality of battery cells (1211) are arranged around the central load-bearing tube (1110) in the circumferential direction, the plurality of battery cells (1211) are fixedly connected together through the battery cell band (1212), the battery cell band (1212) is connected to the electrical core support (1130), and the battery cell (1211) is provided with the first temperature sensor (1213), the first heater (1214), and the battery cell skin (1215).

4. The main and passive coupled temperature control structure of a Mars aircraft according to claim 1 or 2, characterized in that: The core circuit component (1220) includes at least three core PCBs (1221), a PCB connection block (1222), a second temperature sensor (1223), and a second heater (1224); the core PCBs (1221) are evenly arranged around the battery component (1210), the surface of the core PCB (1221) is covered with an electrical core skin (1225), the core PCBs (1221) are connected to the electrical core support (1130) through the PCB connection block (1222), and the core PCB (1221) is provided with the second temperature sensor (1223) and the second heater (1224).

5. A main and passive coupled temperature control structure for a Mars aircraft according to claim 1 or 2, characterized in that: The external sensor assembly (1230) includes an external sensor (1231), a third temperature detector (1232), a third heater (1233), and an external sensor bracket (1234); the external sensor bracket (1234) is fixedly installed on the fuselage box (1100), and the external sensor (1231), the third temperature detector (1232), and the third heater (1233) are all fixedly installed on the external sensor bracket (1234).

6. The main and passive coupling type temperature control structure of a Mars aircraft according to claim 1, characterized in that: The interaction execution component (1300) includes a plurality of structural arms (1310), an actuator (1320), a high-power unit (1330), a low-power unit (1340), a fourth temperature detector (1321), and a fourth heater (1322); the high-power unit (1330) is fixedly connected to the structural interface board (1120) through the structural arm (1310), the actuator (1320) is fixedly connected to the structural interface board (1120) through the structural arm (1310), the fourth temperature detector (1321) and the fourth heater (1322) are provided on the actuator (1320), the low-power unit (1340) is installed on the actuator (1320), the high-power unit (1330) is composed of a high-power motor (1331) and a high-power PCB (1332), and the low-power unit (1340) is composed of a low-power motor (1341) and a low-power PCB (1342).

7. A main and passive coupling type temperature control structure for a Mars aircraft according to claim 1, characterized in that: The photovoltaic power generation component (1400) includes a photovoltaic circuit (1410) and a photovoltaic substrate (1420); the photovoltaic circuit (1410) is fixed on the photovoltaic substrate (1420), and a shunt electric heater (1421) and a shunt radiator (1422) are provided on the lower surface of the photovoltaic substrate (1420).

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