Spacecraft based on spherical solar cell array as shell and thermal control design method thereof

Abstract

The invention provides a spacecraft based on a spherical solar cell array as a shell, the spacecraft comprises a spacecraft spherical shell, the spacecraft spherical shell comprises a metal spherical shell and solar cells arranged on the metal spherical shell, and heat pipes with heat transfer capability are uniformly distributed and attached on the inner side surface of the metal spherical shell. A polyimide film is attached to the surface of the idle outer side, without the solar cell, of the heat absorption area of the spherical shell of the spacecraft. The invention further provides a thermal control design method of the spacecraft based on the spherical solar cell array as the shell. Compared with the prior art, the spacecraft with the spherical solar cell array as the shell and the thermal control design method of the spacecraft have the advantages that the spacecraft with the spherical solar cell array as the shell serves as the basis, temperature equalization is conducted on the spherical solar cell array of the spacecraft through the high-heat-conduction heat pipes, and therefore the spherical solar cell array keeps a small temperature gradient.

Description

Technical Field

[0001] This invention relates to spacecraft, and more particularly to a spacecraft based on a spherical solar cell array as its outer shell and its thermal control design method. Background Technology

[0002] Existing spacecraft solar arrays are generally deployable and do not employ thermal control measures, relying solely on their surface to radiate heat into space. For bulk solar arrays, the radiation channel on the back is usually blocked, resulting in higher temperatures. Therefore, increasing the number of solar cells is necessary to ensure energy output. Bulk-mounted spherical solar arrays are a special case, where the curvature of their surface affects the power generation. Examples include the hemispherical solar array used in the orbital atmospheric density detection experimental satellite developed by the Chinese Academy of Sciences and the spherical solar array assembled by Tsinghua University's gravity and atmospheric science satellite.

[0003] The inner surface of a spherical solar array cannot directly radiate heat to the deep space cold background. It only radiates heat to the deep space cold background through the solar cells on the outer surface. Its heat dissipation efficiency is low, which causes the temperature of the area of ​​the spherical solar array directly exposed to the sun to remain at a high temperature level, generally above 120°C. Under high temperature conditions, the number of solar cells needs to be increased to ensure a certain output voltage requirement.

[0004] For spherical solar arrays with modular assembly, the overall temperature uniformity is relatively low, and the temperature of individual modules exposed to direct sunlight is high. Heat pipes are also inconvenient to install, and the only way to ensure energy output is to increase the number of solar cells. For hemispherical solar arrays, with good thermal conductivity, using the outer surface of the other hemisphere as a heat dissipation surface can ensure that the array temperature is reduced.

[0005] Assembled spherical solar arrays prioritize the ease of fabricating small-sized spherical shells, neglecting the temperature uniformity between modules. This can easily lead to excessively high temperatures in individual sun-exposed modules. This conventional thermal design approach increases the number of solar panels in a spherical array, resulting in relatively lower effective power output for the same surface area. Therefore, a one-piece molded metal structure is needed as the solar array substrate. Leveraging its inherent temperature uniformity, combined with high thermal conductivity heat pipes, the temperature at the sun-exposed locations of the spherical solar array can be reduced to below 70°C, achieving high energy output within the effective area.

[0006] The existing hemispherical solar array, with its separate design on the upper and lower hemispheres, cannot utilize heat pipes to transfer heat from the upper hemisphere to the lower hemisphere for dissipation. This prevents full utilization of the lower hemisphere's heat dissipation capacity, preventing the solar array temperature from dropping below 60°C and resulting in relatively low energy output per unit area. Furthermore, this design is unsuitable for satellite applications with higher power consumption requirements and limits the flexibility of heat dissipation paths for internal equipment. The only external heat dissipation window for internal equipment is the lower hemisphere, increasing the complexity of the configuration layout and the challenge of achieving internal temperature uniformity. Summary of the Invention

[0007] To address the problems in the prior art, this invention provides a spacecraft based on a spherical solar cell array as its outer shell and a thermal control design method thereof.

[0008] The present invention provides a spacecraft based on a spherical solar cell array as the outer shell, including a spherical outer shell of the spacecraft, the spherical outer shell of the spacecraft including a metal spherical shell and solar cells disposed on the metal spherical shell, heat pipes with heat transfer capability are uniformly distributed on the inner surface of the metal spherical shell, and a polyimide film is attached to the heat absorption area of ​​the spacecraft spherical outer shell where there are no solar cells.

[0009] As a further improvement of the present invention, the number of heat pipes installed is calculated according to their thermal conductivity and length.

[0010] As a further improvement of the present invention, the polyimide film is an F46 polyimide film.

[0011] As a further improvement of the present invention, a flexible high thermal conductivity carbon nanotube copper foil is installed on the heat dissipation area and inner surface of the spherical shell of the spacecraft, and the flexible high thermal conductivity carbon nanotube copper foil is in direct contact with the surface of the equipment structure.

[0012] As a further improvement of the present invention, a multi-layer heat insulation assembly is installed in the heat-absorbing area and on the inner surface of the spherical shell of the spacecraft. The multi-layer heat insulation assembly includes alternating layers of polyester mesh and double-sided aluminized polyester film in the middle layer, a double-sided aluminized polyimide film in the upper layer, and a double-sided aluminized polyester film in the lower layer.

[0013] As a further improvement of the present invention, when there are heat pipes in the area covered by the multi-layer heat insulation component, the heat pipes in the area are covered together with the metal spherical shell, and the heat pipes in this area are not exposed.

[0014] As a further improvement of the present invention, the forward direction of the spacecraft is defined as the +X axis, the direction from the center of the spacecraft to the center of the Earth is defined as the +Z axis, and the +Y axis is determined according to the right-hand rule; the spherical shell of the spacecraft is cut into two metal hemispheres along the plane containing the yaw axis (Z axis) and the pitch axis (Y axis), and heat pipes are installed along the inner arc surface in the ±Y axis direction and evenly arranged in the ±Z axis direction. The spherical shell of the spacecraft is internally connected to an internal main structure, and a heat insulation pad is provided between the internal main structure and the inner surface of the heat absorption area of ​​the metal spherical shell.

[0015] As a further improvement of the present invention, the internal main structure adopts an aluminum honeycomb structure with embedded metal parts.

[0016] As a further improvement of the present invention, the internal main structure includes an XY plane supported main structure plate and a YZ plane supported main structure plate.

[0017] This invention also provides a thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell. The thermal control design for this spacecraft includes temperature homogenization design, heat absorption reduction design, and heat dissipation design.

[0018] Temperature homogenization design includes: using heat pipes to homogenize the temperature of the spacecraft's spherical outer shell;

[0019] The design to reduce heat absorption includes attaching an F46 polyimide film to the heat-absorbing areas of the spacecraft's spherical outer shell, on the unused outer surface where there are no solar cells, to reduce the heat absorbed by the spacecraft's spherical outer shell.

[0020] The heat dissipation design includes: installing flexible, highly thermally conductive carbon nanotube copper foil in the heat dissipation area of ​​the spacecraft's spherical shell and on the inner surface; conducting heat from the inner wall to the heat dissipation area of ​​the spacecraft's spherical shell through the flexible, highly thermally conductive carbon nanotube copper foil; and then radiating heat to the deep space cool background through the solar cells on the surface of the spacecraft's spherical shell.

[0021] As a further improvement of the present invention, the heat dissipation capacity of the flexible high thermal conductivity carbon nanotube copper foil is calculated as follows:

[0022]

[0023] Where: Q represents the thermal conductivity and heat dissipation capacity of the carbon nanotube copper foil, in W;

[0024] K represents the thermal conductivity of carbon nanotube copper foil, in W / m² / ℃.

[0025] L represents the heat conduction path distance, in meters (m).

[0026] S represents the cross-sectional area of ​​the heat conduction path of the carbon nanotube copper foil, in square meters;

[0027] ΔT represents the temperature difference between the hot and cold ends of the carbon nanotube copper foil, in °C.

[0028] As a further improvement of the present invention, the absorption and heat dissipation capabilities of the combined surface formed by the solar cell and the F46 polyimide film are calculated as follows:

[0029] q1=a s ×A×S1 (1)

[0030]

[0031] Wherein: Formula (1) is the formula for calculating absorbed heat, and Formula (2) is the formula for dissipating heat.

[0032] a s The equivalent solar absorptivity is calculated based on the area ratio of the solar cells, the exposed surface of the metal sphere, and the F46 polyimide film.

[0033] A represents the total external heat flow from direct solar radiation and Earth's reflected heat, in W / m².

[0034] q represents the heat dissipated by radiation;

[0035] In formula (1), q1 represents the absorbed heat of direct solar radiation, in W.

[0036] In formula (2), q2 represents the radiative heat dissipation capacity, with the unit W / m².

[0037] ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film.

[0038] σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant;

[0039] S1 is the radiative surface area of ​​the spherical shell;

[0040] T is the absolute temperature of the spherical shell surface, in K;

[0041] The regional energy balance formula for the combined surface of the solar cell array and the F46 polyimide film is as follows:

[0042] q1 = q + q0

[0043] Where q0 represents the heat conducted away from the combined surface region through the spherical metal structure and heat pipes, and the heat conducted away through the spherical shell and heat pipes is:

[0044]

[0045] Where: q 01 and q 02 These represent the thermal conductivity and heat dissipation capacity of the spherical shell and the heat pipe, respectively, in W.

[0046] K01 and K 02 The thermal conductivity of the spherical shell and the heat pipe are respectively, in W / m² / ℃;

[0047] L 01 and L 02 These are the heat conduction path distances for the spherical shell and the heat pipe, respectively, in meters (m).

[0048] S 01 and S 02 These are the cross-sectional areas of the heat conduction paths of the spherical shell and the heat pipe, respectively, in square meters (m²).

[0049] ΔT 01 and ΔT 02 These represent the temperature difference between the hot and cold ends of the spherical shell and the heat pipe, respectively, in °C.

[0050] As a further improvement to the present invention, the evaluation calculation of the heat dissipation area is as follows:

[0051] To determine the position where the spacecraft's spherical outer shell surface reaches energy equilibrium, considering the total absorption of external heat flow, the heat loss due to the voltage divider resistance on the back of the shell, and the radiative heat to the cold background of deep space, the energy balance formula is as follows:

[0052] Q 外 +Q 内 =Q 散 (3)

[0053] in,

[0054] Q 外 The external heat flow absorbed by the spacecraft;

[0055] Q 内 The heat generated by the spherical shell itself;

[0056] Q 散 The radiative heat dissipation of the spherical shell to the cold background of deep space is expressed in W.

[0057] Q 外 This includes energy absorbed from direct sunlight, energy reflected from sunlight by the planet, and infrared radiation emitted by the planet itself to the spacecraft.

[0058] Q 外 =q 直 ×α s ×S 直 +q 反 ×α s ×S 反 +q 红 ×ε×S 红 (4)

[0059] in,

[0060] q 直 The heat flux density radiated by the sun to the surface of the spacecraft's spherical shell;

[0061] q 反 The heat flux density of the planet reflecting solar radiation energy to the spacecraft's spherical shell;

[0062] q 红 The heat flux density of infrared energy emitted by the planet itself reaching the spacecraft's spherical shell is expressed in W / m².

[0063] a s The equivalent solar absorptivity is calculated based on the area ratio of the solar cells, the exposed surface of the metal sphere, and the F46 polyimide film.

[0064] ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film.

[0065] σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant;

[0066] S 直 Area directly exposed to sunlight;

[0067] S 反 The area of ​​the Earth's spherical shell where sunlight reflects off it;

[0068] S 红 The visible area of ​​the Earth's own infrared radiation reaching the spherical shell;

[0069] Q 散 The radiative heat from the spherical shell surface to the cold background of deep space;

[0070]

[0071] in,

[0072] ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film.

[0073] σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant;

[0074] S0 is the area of ​​the smallest unit used to calculate the equilibrium locality;

[0075] T is the absolute temperature of the spherical shell surface, measured in K.

[0076] 2) Calculate the temperature equilibrium point according to formula (3), form the boundary line between the heat dissipation area and the heat absorption area based on multiple temperature equilibrium points, and combine the equipment layout and heat conduction path design to restrict the constraints and realize the utilization of the spherical arc area.

[0077] The beneficial effects of the present invention are as follows: Through the above scheme, a spacecraft based on a spherical solar cell array as the outer shell and its thermal control design method are provided. Based on the spacecraft with the spherical solar cell array as the outer shell, a high thermal conductivity heat pipe is used to homogenize the temperature of the spherical solar cell array of the spacecraft, so that the spherical solar cell array itself maintains a small temperature gradient. Attached Figure Description

[0078] To more clearly illustrate the technical solutions of the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other solutions can be obtained based on these drawings without creative effort.

[0079] Figure 1 This is a schematic diagram of a spacecraft based on a spherical solar cell array as its outer shell, according to the present invention. Detailed Implementation

[0080] It should be noted that, unless otherwise specified, the embodiments and features described in the present invention can be combined with each other.

[0081] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this invention. Furthermore, the terms "first," "second," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined with "first," "second," etc., may explicitly or implicitly include one or more of that feature. In the description of this invention, unless otherwise stated, "a plurality of" means two or more.

[0082] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal communication between two components. Those skilled in the art will understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0083] The present invention will be further described below with reference to the accompanying drawings and specific embodiments.

[0084] like Figure 1 As shown, a spacecraft based on a spherical solar cell array as its outer shell includes a spherical outer shell of the spacecraft. The spherical outer shell of the spacecraft includes a metal spherical shell 1 and solar cells disposed on the metal spherical shell 1. Heat pipes 4 with heat transfer capabilities are uniformly distributed on the inner surface of the metal spherical shell 1. A polyimide film is attached to the heat-absorbing area of ​​the spacecraft spherical outer shell where there are no solar cells.

[0085] The number of heat pipes to be installed is calculated based on their thermal conductivity and length.

[0086] The polyimide film is an F46 polyimide film.

[0087] Flexible, highly thermally conductive carbon nanotube copper foil 8 is installed on the heat dissipation area of ​​the spacecraft's spherical shell and on its inner surface.

[0088] The flexible, highly thermally conductive carbon nanotube copper foil 8 is in direct contact with the surface of the internal equipment 7.

[0089] A multi-layer heat insulation component 6 is installed in the heat-absorbing area of ​​the spherical shell of the spacecraft and on the inner surface. The multi-layer heat insulation component 6 includes alternating layers of polyester mesh and double-sided aluminized polyester film in the middle layer, double-sided aluminized polyimide film in the upper layer, and double-sided aluminized polyester film in the lower layer.

[0090] When there are heat pipes in the area covered by the multi-layer insulation assembly 6, the heat pipes 4 and the metal spherical shell 1 in the area are covered together, and the heat pipes 4 are not exposed in this area.

[0091] Define the spacecraft's forward flight direction as the +X axis, the spacecraft's center pointing towards the Earth's center as the +Z axis, and determine the +Y axis according to the right-hand rule. The spacecraft's spherical outer shell is divided into two metal hemispheres along the plane containing the yaw axis (Z axis) and the pitch axis (Y axis). The two metal hemispheres are combined to form a metal spherical shell 1. Heat pipes 4 are installed along the inner arc surface in the ±Y axis direction and evenly arranged in the ±Z axis direction. The spacecraft's spherical outer shell is internally connected to an internal main structure. A heat insulation pad 5 is provided between the internal main structure and the inner surface of the heat-absorbing area of ​​the metal spherical shell.

[0092] The internal main structure adopts an aluminum honeycomb structure with embedded metal components.

[0093] The internal main structure includes an XY plane supported main structure plate 2 and a YZ plane supported main structure plate 3.

[0094] This invention also provides a thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell. Based on the spacecraft with the spherical solar cell array as its outer shell, a high thermal conductivity heat pipe is used to homogenize the temperature of the spherical solar cell array, so that the spherical solar cell array itself maintains a small temperature gradient. At the same time, the contact thermal resistance between the internal main structure and each part of the spherical solar cell array is designed so that the spherical solar cell array and the internal equipment are all within a good temperature range.

[0095] Define the spacecraft's forward flight direction as the +X axis, the spacecraft's center pointing towards the Earth's center as the +Z axis, and determine the +Y axis according to the right-hand rule. Divide the spacecraft's spherical outer shell into two hemispheres along the plane containing the yaw axis (Z-axis) and pitch axis (Y-axis). Based on the maximum variation range (±90°) of the Beta angle (orbital solar angle, the angle between the solar direction and the satellite's orbital plane), sunlight can directly illuminate the celestial body within a 180° region from the +Y direction to the -Y direction (passing through the -Z direction), ensuring that the two hemispheres equally distribute the overall direct solar energy. Simultaneously, the Earth's infrared radiation to the celestial body and the energy emitted by the Earth to the Sun can also be evenly distributed between the two metal hemispheres. Based on the premise of equal heat flow distribution between the two metal hemispheres, to ensure the temperature range of the spherical solar array, only temperature homogenization of the solar array needs to be ensured. This requires uniformly distributing heat pipes 4 with specific heat transfer capabilities on the inner sides of the solar array in both hemispheres. Additionally, within a 100cm radius of the illuminated area without solar cells... 2F46 polyimide film is applied to the unused area to reduce heat absorption. The thermal conductivity of the spun metal spherical shell 1 is then utilized to achieve uniform temperature within the spherical solar array, reducing the temperature difference between the highest temperature point in the irradiated external heat flow area and the lowest temperature point facing the cryogenic background. For spacecraft internal equipment with the spherical solar array as its outer shell, a good operating temperature design is crucial. First, heat transfer from the spherical solar array must be prevented. Within the heat-absorbing area of ​​the spherical solar array, a 15-unit multi-layer thermal insulation component 6 (alternating layers of polyester mesh and 6μm thick double-sided aluminized polyester film, stacked 15 times, with a 10μm thick double-sided aluminized polyester film at the bottom, and a 20μm thick double-sided aluminized polyimide film on top) is applied for radiative heat exchange between the solar array and internal equipment, covering the multi-layer thermal insulation component. In region 6, where heat pipe 4 is present, heat pipe 4 and the surface of the spherical shell need to be covered together, and heat pipe 4 should not be exposed in this region. At the same time, a 10mm thick heat insulation pad 5 is used to thermally connect the heat absorption area of ​​the spherical solar cell array with the internal main structure to reduce heat transfer. The area of ​​the spherical solar cell array with heat dissipation capacity is used as a heat dissipation window for the internal equipment to the deep space cold background. By designing the contact heat transfer coefficient between the main structure and the spherical solar cell array, the heat conduction path between the equipment and the spherical solar cell array is shortened, so that the internal equipment has a good heat transfer and heat dissipation channel.

[0096] The specific measures are as follows: The spherical solar array of the spacecraft is divided into two hemispheres along the plane containing the yaw axis (Z-axis) and pitch axis (Y-axis). Heat pipes 4 with a certain heat transfer capacity are evenly distributed and attached to the inner side of the two spun metal hemisphere-shaped solar arrays. The heat pipes 4 are installed along the inner arc surface in the ±Y-axis direction and evenly arranged in the ±Z-axis direction. The number of heat pipes 4 is determined according to the heat absorbed by the spherical solar array and the heat transfer capacity of the heat pipes themselves (for example, if a single hemisphere solar array absorbs 1000W of heat under illumination, and the energy radiated from its illuminated area is 500W, then the heat that needs to be conducted to other non-illuminated areas of the solar array through the heat pipes 4 is 500W. With the selected heat pipe 4 having a heat transfer capacity of 200W*m and a length of 2m, five heat pipes 4 need to be evenly distributed without considering the derating of the heat pipes 4). At the same time, F46 polyimide film is attached to the surface of the illuminated area where no solar cells are attached to reduce heat absorption and further reduce the temperature value of the illuminated area. The internal equipment of the spacecraft is temperature-neutralized through heat pipes 4, flexible high thermal conductivity carbon nanotube copper foil 8, and embedded metal components within the aluminum honeycomb structure. The good thermal conductivity between the internal structural plates and the heat-dissipating spherical shell allows heat from the internal equipment to radiate to the deep-space cold background through the surface of the solar array. The overall thermal resistance of the heat conduction path is related to the amount of heat to be dissipated and the heat dissipation capacity of the heat-dissipating area of ​​the solar array, requiring detailed design and simulation (e.g., if the internal equipment has a heat dissipation of 100W and its heat conduction path thermal resistance is 10℃ / W, then the temperature difference between the equipment and the heat dissipation area of ​​the spherical shell should not exceed 10℃; if the equipment is maintained below 40℃, the temperature of the heat dissipation area of ​​the spherical shell should not exceed 30℃). The spacecraft thermal design principle based on a spherical solar array as the outer shell is as follows... Figure 1 As shown.

[0097] The implementation principle: The thermal control design of a spacecraft system with a spherical solar array as its shell mainly consists of three parts. Firstly, the temperature of the spherical solar array is ensured to meet cooling design requirements by controlling the heat pipe layout and insulation design measures. Secondly, F46 polyimide film is applied to the areas of the solar array where no solar cells are located in the irradiated area to reduce the heat absorbed by the solar array. This part innovatively utilizes a flexible film application method to change the solar absorptivity and infrared emissivity of the solar array surface. Thirdly, the heat from the inner wall is conducted to the heat-dissipating areas of the spherical solar array through the thermal conductivity of multilayer flexible high thermal conductivity carbon nanotube copper foil, and then radiated to the deep-space cool background through the solar cells on the surface of the solar array. This part utilizes the unique enhanced heat dissipation method of flexible high thermal conductivity carbon nanotube copper foil, whose heat conduction capacity is:

[0098]

[0099] Where: Q represents the thermal conductivity and heat dissipation capacity of the carbon nanotube copper foil, in W;

[0100] K represents the thermal conductivity of carbon nanotube copper foil, in W / m² / ℃.

[0101] L represents the heat conduction path distance, in meters (m).

[0102] S represents the cross-sectional area of ​​the heat conduction path of the carbon nanotube copper foil, in square meters;

[0103] ΔT represents the temperature difference between the hot and cold ends of the carbon nanotube copper foil, in °C.

[0104] Q is the heat conducted through copper foil. With the heat being constant, what is ultimately controlled is Delta T (the temperature difference between the hot end temperature T1 and the cold end temperature T2). When the hot end temperature T1 is controlled at a specific value, the equipment installed near its structural plate can be guaranteed to be within the required temperature range.

[0105] The use of F46 polyimide film leverages its low solar absorptivity and high infrared emissivity. Since solar cells themselves have high solar absorptivity and infrared emissivity, attaching F46 film reduces the equivalent solar absorptivity of the solar array surface, thereby reducing heat loss while maintaining a high capacity for infrared radiation against the cold cosmic background. The combined surface absorption and heat dissipation capabilities of the solar array surface and the F46 polyimide film are as follows:

[0106] q1=a s ×A×S1 (1)

[0107] Wherein: Formula (1) is the formula for calculating heat absorption, and Formula (2) is the formula for heat dissipation. The meanings of each symbol are as follows:

[0108] a s The equivalent solar absorptivity is calculated based on the area ratio of the solar cells, the exposed surface of the metal spherical shell, and the F46 polyimide film.

[0109] A represents the total external heat flow from direct solar radiation and Earth's reflected heat, in W / m².

[0110] q represents the heat dissipated by radiation;

[0111] In Formula 1, q1 represents the absorbed heat from direct solar radiation, in W; and in Formula 2, q2 represents the radiative heat dissipation capacity, in W / m².

[0112] ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the battery cell, the exposed surface of the metal spherical shell, and the F46 polyimide film.

[0113] σ is the radiation constant of a blackbody (Stephen-Boltzmann constant);

[0114] S1 is the radiative surface area of ​​the spherical shell;

[0115] T is the absolute temperature of the spherical shell surface, in K;

[0116] The regional energy balance formula for the combined surface of the solar cell array and the F46 polyimide film is as follows:

[0117] q1 = q + q0

[0118] Where q0 represents the heat conducted away from the combined surface region through the spherical metal structure and heat pipes, and the heat conducted away through the spherical shell and heat pipes is:

[0119]

[0120] Where: q 01 and q 02 These represent the thermal conductivity and heat dissipation capacity of the spherical shell and the heat pipe, respectively, in W.

[0121] K 01 and K 02 The thermal conductivity of the spherical shell and the heat pipe are respectively, in W / m² / ℃;

[0122] L 01 and L 02 These are the heat conduction path distances for the spherical shell and the heat pipe, respectively, in meters (m).

[0123] S 01 and S 02 These are the cross-sectional areas of the heat conduction paths of the spherical shell and the heat pipe, respectively, in square meters (m²).

[0124] ΔT 01 and ΔT 02 These represent the temperature differences at the hot and cold ends of the spherical shell and heat pipe, respectively, in °C. Detailed calculation and simulation are required to determine the location of internal equipment conducting heat to the spherical solar cell array. The following calculation method can be used to evaluate the heat dissipation area:

[0125] 1) Determining the location where the spherical shell surface can achieve energy balance mainly considers the total amount of heat flow absorbed from the orbit, the heat loss due to the voltage divider resistance on the back of the shell, the radiant heat to the cold background of deep space, and the local temperature level. The energy balance formula is as follows:

[0126] Q 外 +Q 内 =Q 散 (3)

[0127] Q 外 The extraorbital heat flux absorbed by the spacecraft, Q 内 Q is the heat generated by the spherical shell itself. 散 The values ​​represent the radiative heat dissipation of the spherical shell to the deep space cold background, all in W.

[0128] Q 外This includes energy absorbed from direct sunlight, energy reflected from sunlight by the planet, and infrared radiation emitted by the planet itself to the spacecraft.

[0129] Q 外 =q 直 ×α s ×S 直 +q 反 ×α s ×S 反 +q 红 ×ε×S 红 (4)

[0130] Where: q 直 Let q be the heat flux density radiated from the sun to the surface of the spacecraft's spherical shell. 反 q is the heat flux density of the planet reflecting solar radiation energy to the spacecraft's spherical shell. 红 The heat flux density of infrared energy emitted by the planet itself reaching the spacecraft's spherical shell is expressed in W / m².

[0131] α s Consistent with formula (1);

[0132] ε is consistent with that in formula (2);

[0133] σ is consistent with that in formula (2);

[0134] S 直 Area directly exposed to sunlight, S 反 The area of ​​the Earth's spherical shell where sunlight reflects off it, S 红 The visible area of ​​the Earth's spherical shell reached by its own infrared radiation.

[0135] Q 散 This represents the radiative heat from the spherical shell surface to the deep-space cold background.

[0136]

[0137] Where: ε is consistent with formula (2);

[0138] σ is consistent with formula (2);

[0139] S0 is the smallest unit area for calculating the equilibrium local area. It is recommended to use the surface area occupied by a single solar cell as the unit, and to consider the surrounding F46 polyimide surface and the exposed metal surface being equally divided proportionally.

[0140] T is the absolute temperature of the spherical shell surface, measured in K.

[0141] 2) Calculate the temperature equilibrium point according to Formula 3, and form the boundary line between the heat dissipation area and the heat absorption area based on multiple equilibrium points. Then, combine the equipment layout and heat conduction path design constraints to make reasonable use of the better spherical arc surface area.

[0142] Analysis shows that by utilizing flexible, highly thermally conductive carbon nanotube copper foil for flexible installation and by subdividing the heat dissipation capacity of the spherical shell, the design objectives of optimal performance, shortest cycle, and fewest iterations can be achieved. This effectively reduces the temperature of the solar array, thereby reducing the number of solar panels and the area of ​​the solar array.

[0143] This invention provides a spacecraft based on a spherical solar cell array as its outer shell and its thermal control design method. Without changing the initial design of the main body structure, a low thermal resistance heat conduction path is constructed for the internal equipment using flexible carbon nanotube copper foil, which can meet the heat dissipation requirements of the internal equipment. At the same time, by comprehensively designing the heat insulation and temperature uniformity of each part, as well as changing the equivalent solar absorptivity and infrared emissivity, the temperature index requirements after reducing the number of solar cell array panels can be achieved, and the output voltage of the solar cell array can meet specific requirements.

[0144] The above description, in conjunction with specific preferred embodiments, provides a further detailed explanation of the present invention. It should not be construed that the specific implementation of the present invention is limited to these descriptions. For those skilled in the art, various simple deductions or substitutions can be made without departing from the concept of the present invention, and all such modifications and substitutions should be considered within the scope of protection of the present invention.

Claims

1. A spacecraft based on a spherical solar cell array as its outer shell, characterized in that: The spacecraft includes a spherical outer shell, which comprises a metal spherical shell and solar cells disposed on the metal spherical shell. Heat pipes with heat transfer capabilities are uniformly distributed on the inner surface of the metal spherical shell, and a polyimide film is attached to the heat-absorbing area of ​​the spacecraft spherical shell where there are no solar cells.

2. The spacecraft based on a spherical solar cell array as its outer shell according to claim 1, characterized in that: The polyimide film is an F46 polyimide film.

3. The spacecraft based on a spherical solar cell array as its outer shell according to claim 1, characterized in that: Flexible, highly thermally conductive carbon nanotube copper foil is installed on the heat dissipation area of ​​the spacecraft's spherical outer shell and on its inner surface.

4. The spacecraft based on a spherical solar cell array as its outer shell according to claim 1, characterized in that: A multi-layer heat insulation assembly is installed on the heat-absorbing area and inner surface of the spherical shell of the spacecraft. The multi-layer heat insulation assembly includes alternating layers of polyester mesh and double-sided aluminized polyester film in the middle layer, double-sided aluminized polyimide film in the upper layer, and double-sided aluminized polyester film in the lower layer.

5. The spacecraft based on a spherical solar cell array as its outer shell according to claim 3, characterized in that: When heat pipes are present in areas covered by multi-layer insulation components, the heat pipes in the area are covered together with the metal spherical shell, and the heat pipes in this area are not exposed.

6. The spacecraft based on a spherical solar cell array as its outer shell according to claim 1, characterized in that: Define the spacecraft's forward flight direction as the +X axis, the spacecraft's center pointing towards the Earth's center as the +Z axis, and determine the +Y axis according to the right-hand rule. The spacecraft's spherical outer shell is divided into two metal hemispheres along the plane containing the yaw axis (Z axis) and the pitch axis (Y axis). Heat pipes are installed along the inner arc surface in the ±Y axis direction and evenly arranged in the ±Z axis direction. An internal main structure is connected inside the spacecraft's spherical outer shell, and a heat insulation pad is provided between the internal main structure and the inner surface of the heat-absorbing area of ​​the metal spherical shell.

7. A thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell, characterized in that: The thermal control design of the spacecraft based on any one of claims 1 to 6 includes temperature equalization design, heat absorption reduction design, and heat dissipation design. Temperature homogenization design includes: using heat pipes to homogenize the temperature of the spacecraft's spherical outer shell; The design to reduce heat absorption includes attaching an F46 polyimide film to the heat-absorbing areas of the spacecraft's spherical outer shell, on the unused outer surface where there are no solar cells, to reduce the heat absorbed by the spacecraft's spherical outer shell. The heat dissipation design includes: installing flexible, highly thermally conductive carbon nanotube copper foil in the heat dissipation area of ​​the spacecraft's spherical shell and on the inner surface; conducting heat from the inner wall to the heat dissipation area of ​​the spacecraft's spherical shell through the flexible, highly thermally conductive carbon nanotube copper foil; and then radiating heat to the deep space cool background through the solar cells on the surface of the spacecraft's spherical shell.

8. The thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell according to claim 8, characterized in that: The heat dissipation capacity of the flexible, highly thermally conductive carbon nanotube copper foil is calculated as follows: Where: Q represents the thermal conductivity and heat dissipation capacity of the carbon nanotube copper foil, in W; K represents the thermal conductivity of carbon nanotube copper foil, in W / m² / ℃. L represents the heat conduction path distance, in meters (m). S represents the cross-sectional area of ​​the heat conduction path of the carbon nanotube copper foil, in square meters; ΔT represents the temperature difference between the hot and cold ends of the carbon nanotube copper foil, in °C.

9. The thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell according to claim 8, characterized in that: The absorption and heat dissipation capabilities of the combined surface formed by the solar cell and the F46 polyimide film are calculated as follows: q1=a s ×A×S1 (1) Wherein: Formula (1) is the formula for calculating absorbed heat, and Formula (2) is the formula for dissipating heat. a s The equivalent solar absorptivity is calculated based on the area ratio of the solar cells, the exposed surface of the metal sphere, and the F46 polyimide film. A represents the total external heat flow from direct solar radiation and Earth's reflected heat, in W / m². q represents the heat dissipated by radiation; In formula (1), q1 represents the absorbed heat of direct solar radiation, in W. In formula (2), q2 represents the radiative heat dissipation capacity, with the unit W / m². ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film. σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant; S1 is the radiative surface area of ​​the spherical shell; T is the absolute temperature of the spherical shell surface, in K; The regional energy balance formula for the combined surface of the solar cell array and the F46 polyimide film is as follows: q1 = q + q0 Where q0 represents the heat conducted away from the combined surface region through the spherical metal structure and heat pipes, and the heat conducted away through the spherical shell and heat pipes is: Where: q 01 and q 02 These represent the thermal conductivity and heat dissipation capacity of the spherical shell and the heat pipe, respectively, in W. K 01 and K 02 The thermal conductivity of the spherical shell and the heat pipe are respectively, in W / m² / ℃; L 01 and L 02 These are the heat conduction path distances for the spherical shell and the heat pipe, respectively, in meters (m). S 01 and S 02 These are the cross-sectional areas of the heat conduction paths of the spherical shell and the heat pipe, respectively, in square meters (m²). ΔT 01 and ΔT 02 These represent the temperature difference between the hot and cold ends of the spherical shell and the heat pipe, respectively, in °C.

10. The thermal control design method for a spacecraft based on a spherical solar cell array as its outer shell according to claim 8, characterized in that: The evaluation calculations for the heat dissipation area are as follows: To determine the position where the spacecraft's spherical outer shell surface reaches energy equilibrium, considering the total absorption of external heat flow, the heat loss due to the voltage divider resistance on the back of the shell, and the radiative heat to the cold background of deep space, the energy balance formula is as follows: Q 外 +Q 内 =Q 散 (3) in, Q 外 The external heat flow absorbed by the spacecraft; Q 内 The heat generated by the spherical shell itself; Q 散 The radiative heat dissipation of the spherical shell to the cold background of deep space is expressed in W. Q 外 This includes energy absorbed from direct sunlight, energy reflected from sunlight by the planet, and infrared radiation emitted by the planet itself to the spacecraft. Q 外 =q 直 ×α s ×S 直 +q 反 ×α s ×S 反 +q 红 ×ε×S 红 (4) in, q 直 The heat flux density radiated by the sun to the surface of the spacecraft's spherical shell; q 反 The heat flux density of the planet reflecting solar radiation energy to the spacecraft's spherical shell; q 红 The heat flux density of infrared energy emitted by the planet itself reaching the spacecraft's spherical shell is expressed in W / m². a s The equivalent solar absorptivity is calculated based on the area ratio of the solar cells, the exposed surface of the metal sphere, and the F46 polyimide film. ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film. σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant; S 直 Area directly exposed to sunlight; S 反 The area of ​​the Earth's spherical shell where sunlight reflects off it; S 红 The visible area of ​​the Earth's own infrared radiation reaching the spherical shell; Q 散 The radiative heat from the spherical shell surface to the cold background of deep space; in, ε is the equivalent infrared emissivity of the outer surface of the spherical shell, calculated based on the area ratio of the solar cell, the exposed surface of the metal spherical shell, and the F46 polyimide film. σ is the radiation constant of a blackbody, also known as the Stefan-Boltzmann constant; S0 is the area of ​​the smallest unit used to calculate the equilibrium locality; T is the absolute temperature of the spherical shell surface, measured in K. 2) Calculate the temperature equilibrium point according to formula (3), form the boundary line between the heat dissipation area and the heat absorption area based on multiple temperature equilibrium points, and combine the equipment layout and heat conduction path design to restrict the constraints and realize the utilization of the spherical arc area.

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