Silicon series thermal control film layer and composite solar cell array for commercial spaceflight

Abstract

This invention discloses a silicon-based thermal control film and composite solar cell array for commercial aerospace applications. The thermal control film employs a multi-layer composite structure consisting of a surface modification layer, a buffer layer, a seed layer, a metal reflective layer, and an anti-oxidation layer, bonded to the back of a glass fiber reinforced polyimide flexible substrate. Interface modification technology solves the problems of poor adhesion and easy detachment in traditional flexible thermal control coatings, while a gradient protection design resists low Earth orbit atomic oxygen erosion. When applied to silicon-based flexible solar cell arrays, it achieves a solar absorptivity ≤0.10, a hemispherical emissivity >0.88, and a 20% improvement in heat dissipation efficiency compared to traditional Z-93 white paint. This reduces the on-orbit temperature of the cell array by 24°C, effectively increasing the cell conversion efficiency by 10%. It is suitable for thermal management of large-area HJT cells, perovskite-silicon tandem cells, and other commercial aerospace flexible solar cell arrays.

Description

Technical Field

[0001] This invention belongs to the field of thermal management technology for spacecraft power systems, specifically relating to a silicon-based thermal control film and composite solar cell array for commercial spaceflight, particularly suitable for thermal management of silicon-based flexible solar cell arrays such as large-area heterojunction (HJT) cells and perovskite-silicon tandem cells used in low Earth orbit commercial satellites. Background Technology

[0002] With the explosive growth of the commercial space industry, the construction of low Earth orbit satellite constellations has entered an accelerated phase, placing stringent demands on the lightweight, low-cost, and long-life requirements of spacecraft power systems. Flexible solar arrays, with their advantages of high expansion-to-contraction ratio, lightweight design, and rollable transport, have gradually replaced traditional rigid solar arrays, becoming the mainstream power supply solution for commercial satellites. Among them, HJT cells, due to their high conversion efficiency, good adaptability to low-temperature processes, and strong radiation resistance, and perovskite-silicon tandem cells, with theoretical efficiencies exceeding 35%, have become the core development direction for next-generation space photovoltaic technology.

[0003] However, the inherently high temperature coefficient of silicon-based solar cells (approximately -0.45% / ℃) is far higher than that of triple-junction gallium arsenide cells (approximately -0.25% / ℃), becoming a key bottleneck restricting their on-orbit performance. In the sunlit areas of space, the combined effect of direct sunlight and the cells' own heat generation can cause the surface temperature of flexible solar arrays to rise above 80℃. For every 10℃ increase, the cell conversion efficiency decreases by 4% to 5%, severely impacting the satellite's power supply capacity and mission endurance.

[0004] Existing thermal control technologies for flexible solar arrays suffer from two main defects: First, traditional satellite white paint (such as Z-93 white paint), while possessing high infrared emissivity, has a high solar absorptivity (αs=0.15) and poor thermal conductivity, failing to effectively diffuse local hot spots on the array surface, leading to uneven temperature distribution and further exacerbating efficiency degradation. Second, flexible thermal control films such as silver-plated F-46, although reducing solar absorptivity through a metal reflective layer, suffer from weak interfacial bonding between the polyimide substrate and the silver layer, making them prone to cracking and peeling under the cold and hot cycling environment of space. Simultaneously, the abundant atomic oxygen in the low Earth orbit environment reacts with the silver layer to form black silver oxide, causing a sharp increase in solar absorptivity within 1-2 years, significantly reducing thermal control performance and failing to meet the 5-15 year on-orbit lifespan requirements of commercial satellites.

[0005] In addition, existing thermal control solutions are mostly general-purpose designs that are not optimized for the temperature characteristics of silicon-based solar cells, nor do they solve the consistency and cost issues in mass production, making them difficult to adapt to the large-scale deployment needs of commercial space constellations. Summary of the Invention

[0006] To address the aforementioned deficiencies in existing technologies, the present invention aims to provide a silicon-based thermal control film layer with strong interfacial bonding, resistance to atomic oxygen erosion, and good heat dissipation performance, as well as a composite solar cell array using this thermal control film layer. Through a multi-layer gradient structure design, the invention solves the problems of easy peeling, easy oxidation, and low heat dissipation efficiency of traditional thermal control coatings, significantly reducing the on-orbit temperature of silicon-based flexible solar cell arrays and improving their conversion efficiency and service life.

[0007] In this invention, the "silicon-based thermal control film composite structure" (hereinafter referred to as "thermal control film") refers to an integral structure composed of a glass fiber reinforced polyimide substrate and multiple functional film layers through physical or chemical methods. The glass fiber reinforced polyimide substrate serves as the substrate, and the surface modification layer, buffer layer, seed layer, metal reflective layer, and anti-oxidation layer deposited sequentially on it constitute the thermal control functional layers. One side of the glass fiber reinforced polyimide substrate of this structure is the composite surface with the solar cell array, and the other side of the anti-oxidation layer is the surface exposed to the space environment.

[0008] To achieve the above objectives, the present invention adopts the following technical solution: A silicon-based thermal control film composite structure for commercial aerospace uses, comprising, from the composite surface with the battery array to the exposed surface, a glass fiber reinforced polyimide (GF / PI) substrate, a surface modification layer, a buffer layer, a seed layer, a metal reflective layer, and an anti-oxidation layer.

[0009] Furthermore, the glass fiber reinforced polyimide substrate serves as the supporting substrate for the thermal control film layer and also as the base material for the flexible solar cell array, with a thickness of 0.1~0.5mm, preferably 0.2mm. When the substrate thickness is less than 0.1mm, its mechanical strength decreases significantly, making it unable to withstand the thermal stress generated by the space thermal cycles, and it is prone to warping and deformation leading to film layer cracking. When the substrate thickness is greater than 0.5mm, it significantly increases the overall weight of the solar cell array, while also increasing thermal resistance and reducing heat dissipation efficiency. Within the 0.1~0.5mm range, the substrate can meet the mechanical strength requirements of the space environment while ensuring lightweight and heat dissipation performance. This substrate, through glass fiber reinforcement modification, has a thermal expansion coefficient reduced by more than 60% and a mechanical strength increased by more than 50% compared to pure polyimide, effectively resisting the thermal stress generated by the space thermal cycles and preventing film layer cracking.

[0010] Furthermore, the surface modification layer is a polyorganosiloxane with a thickness of 0.1~20μm, preferably 5μm. The polyorganosiloxane molecule contains a large number of silanol groups, which can form covalent bonds with the carboxyl groups generated on the surface of the GF / PI substrate through plasma activation, thereby improving the interfacial adhesion from traditional physical adsorption to chemical bonding, significantly enhancing the interfacial adhesion and solving the problem of film peeling.

[0011] Furthermore, the buffer layer is a nitride (Si3N4, AlN) or oxide (TiO2, ZnO) with a thickness of 10~100nm, preferably 20nm Si3N4. This layer is used to match the difference in thermal expansion coefficients between the GF / PI substrate and the metal reflective layer, alleviate thermal mismatch stress during thermal cycling, and at the same time act as a barrier layer to prevent metal atoms from diffusing into the substrate.

[0012] Furthermore, the seed layer is a NiCr alloy (Ni:Cr=80:20) or Al2 / 3O. 1-x The thickness of this layer is 2~10nm, preferably 5nm NiCr. The lattice constant of this layer is highly matched with that of metal reflective layers such as aluminum and silver, which can induce the metal grains to grow along a preferred orientation, significantly improving the flatness and crystallinity of the metal reflective layer, thereby enhancing its reflectivity and thermal conductivity.

[0013] Furthermore, the metal reflective layer is made of aluminum or silver, with a thickness of 100-500 nm, preferably 200 nm aluminum. When the aluminum reflective layer thickness is less than 180 nm, the aluminum layer is discontinuous, with numerous pinhole defects, leading to a significant increase in solar absorptivity. When the thickness is greater than 220 nm, the reflectivity increase is less than 0.5%, but material costs and deposition time increase significantly. Within the 180-220 nm range, the aluminum layer is continuous and dense, ensuring the thermal control requirement of a solar absorptivity ≤0.10 while also considering production efficiency and cost. This layer is the core functional layer of the thermal control film. On the one hand, it reflects more than 92% of sunlight through high reflectivity, reducing the solar absorptivity; on the other hand, it rapidly diffuses the heat generated by the battery array laterally due to its high thermal conductivity, eliminating local hot spots. Choosing aluminum instead of silver as the preferred material can reduce material costs by more than 70%, making it suitable for commercial aerospace mass applications.

[0014] Further, the anti-oxidation layer is a NiCr alloy with a thickness of 500 nm to 1 μm, preferably 500 nm. This invention uses a NiCr alloy as the anti-oxidation layer material, with a thickness controlled between 500 nm and 1 μm, and employs magnetron sputtering to deposit the layer, ensuring a film density >95%. This structural design allows a dense and stable Cr2O3 passivation film to rapidly form on the surface of the anti-oxidation layer when exposed to the space environment, effectively blocking the erosion of atomic oxygen in low Earth orbit. By using the structure and preparation process described in claims 1-6 of this invention, an atomic oxygen flux of 1×10¹ can be achieved. 6 After irradiation with atoms / cm², the change in solar absorptivity is less than 0.02, ensuring the thermal control film layer maintains stable performance for more than 15 years in orbit.

[0015] Furthermore, the thermal control film composite structure of the present invention achieves excellent thermal control performance and environmental adaptability through the synergistic effect of the above-mentioned layers: solar absorptivity αs≤0.10, hemispherical emissivity ε>0.88; after 1000 cycles of -180℃ to +120℃, the film adhesion retention rate is >95%; after an atomic oxygen flux of 1×10¹ 6 After irradiation with atoms / cm², the change in solar absorptivity is <0.02.

[0016] This invention also provides a composite solar cell array for commercial aerospace, comprising a silicon-based flexible solar cell array body and the aforementioned thermal control film composite structure. The thermal control film composite structure is bonded to the back of the silicon-based flexible solar cell array body using an epoxy adhesive (such as 3MDP460) with a temperature resistance of not less than 200°C, without altering the original structure of the cell array, and is directly compatible with existing manufacturing processes.

[0017] Furthermore, the silicon-based flexible solar cell array body includes, but is not limited to, large-area HJT flexible solar cell arrays and perovskite-silicon tandem flexible solar cell arrays.

[0018] The beneficial effects of this invention are as follows: Significantly improved heat dissipation performance: Through the synergistic design of a low solar absorptivity metal reflective layer and a high infrared emissivity surface, the heat dissipation efficiency of the thermal control film layer is 20% higher than that of the traditional Z-93 white paint. This can reduce the on-orbit temperature of silicon-based flexible solar cell arrays from 82.4℃ to 58.4℃, a reduction of 24℃, which is equivalent to a 10% increase in cell conversion efficiency.

[0019] Significantly enhanced environmental adaptability: The surface modification layer enhances interface adhesion, the buffer layer alleviates thermal mismatch, and the anti-oxidation layer resists atomic oxygen erosion, solving the problems of easy peeling and oxidation of traditional flexible thermal control coatings. It is expected to have an on-orbit service life of more than 15 years, meeting the long-life requirements of commercial aerospace.

[0020] Low cost and strong adaptability: Low-cost aluminum is preferred as the metal reflective layer, reducing material costs by more than 70%; the thermal control film layer can be mass-produced through roll-to-roll process and integrated with existing solar cell arrays through bonding, without the need to modify the production line, making it suitable for large-scale deployment of commercial aerospace constellations.

[0021] Good temperature uniformity: The high thermal conductivity metal reflective layer can quickly diffuse heat laterally, keeping the temperature difference on the surface of the battery array within ±3℃, avoiding battery performance degradation and shortened lifespan caused by local overheating. Attached Figure Description

[0022] Figure 1 This is a schematic diagram of the overall structure of a composite solar cell array, where 1 represents the silicon-based flexible solar cell array body, 2 represents the epoxy adhesive layer, and 3 represents the thermal control film composite structure.

[0023] Figure 2 This is a schematic diagram of the layered structure of a silicon-based thermal control film composite, where 4 represents the GF / PI substrate, 5 represents the surface modification layer, 6 represents the Si3N4 buffer layer, 7 represents the NiCr seed layer, 8 represents the aluminum reflective layer, and 9 represents the NiCr anti-oxidation layer. Figure 2 The left arrow points to the composite surface with the battery array, and the right arrow points to the exposed surface. Detailed Implementation

[0024] The specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings. This embodiment is implemented based on the technical solution of the present invention, and provides detailed implementation steps and operating parameters; however, the scope of protection of the present invention is not limited to the following embodiment.

[0025] The overall structure of this embodiment is as follows: Figure 1 As shown, the composite solar cell array is composed of a silicon-based flexible solar cell array body 1, an epoxy adhesive layer 2, and a thermal control film layer composite structure 3, which are sequentially stacked; the layered composition of the thermal control film layer composite structure 3 is as follows: Figure 2 As shown, from the composite surface with the battery array to the exposed surface, the layers are sequentially arranged as follows: GF / PI substrate 4, surface modification layer 5, Si3N4 buffer layer 6, NiCr seed layer 7, aluminum reflective layer 8, and NiCr anti-oxidation layer 9. The overall process flow of this embodiment is as follows: first, the GF / PI substrate is cleaned and surface activated; then, according to… Figure 2 The layer sequence shown is used to deposit each functional film layer sequentially to obtain a thermal control film layer composite structure; finally, according to... Figure 1 The assembly relationship shown demonstrates how the thermal control film composite structure is bonded to the back of the HJT flexible solar cell array using epoxy adhesive to obtain the final product.

[0026] Step 1: Pretreatment of GF / PI substrate: A 0.2mm thick GF / PI substrate, modified with glass fiber reinforcement, exhibits a coefficient of thermal expansion reduced by over 60% and mechanical strength increased by over 50% compared to pure polyimide. This allows it to withstand the thermal stress generated by extreme thermal cycling in space, meeting the requirements of space environments. The substrate was cut into standard 100mm × 100mm samples and subjected to a multi-step ultrasonic cleaning process: first, ultrasonic cleaning with anhydrous ethanol for 10 minutes, utilizing ethanol's organic solvent properties to dissolve and remove oil, fingerprints, and other organic impurities from the substrate surface; then, ultrasonic cleaning with deionized water for 10 minutes to rinse away residual ethanol and water-soluble impurities; finally, ultrasonic cleaning with isopropanol for 5 minutes, leveraging isopropanol's miscibility with water and its lower boiling point to quickly remove moisture from the substrate surface, preventing water residue from affecting the adhesion of subsequent film layers. After cleaning, the substrate was dried with 99.999% pure nitrogen gas to ensure no moisture or impurities remained on the surface.

[0027] The dried substrate was placed in an RF plasma cleaner, and the vacuum level was evacuated to ≤1×10⁻³Pa. Argon gas was introduced as the working gas, with a flow rate of 50 sccm and a working pressure of 10 Pa. The RF power supply was set to 100W, and the substrate surface was subjected to plasma activation treatment for 5 minutes. This treatment breaks the original chemical bonds and introduces a large number of carboxyl active groups by bombarding the substrate surface with high-energy argon ions, providing reaction sites for the subsequent chemical bonding of surface modification layers and solving the problems of weak binding force and easy detachment of traditional physical adsorption.

[0028] Step 2: Surface modification layer deposition: Mix the polyorganosiloxane DC-184 stock solution with ethyl acetate at a volume ratio of 1:3. Stir with a magnetic stirrer for 10 minutes until the solution is homogeneous. Let stand for 30 minutes to remove air bubbles and prevent pinhole defects caused by bubble breakage during spin coating. Drop the prepared solution onto the surface of the activated substrate obtained in step 1. Start the spin coater, first rotating at a low speed of 500 rpm for 5 seconds to evenly spread the solution across the entire substrate surface, then rotating at a high speed of 3000 rpm for 30 seconds to precisely control the film thickness. The higher the speed and the longer the time, the thinner the final film thickness. Place the spin-coated sample in a forced-air oven and cure using a programmed temperature rise method: first, hold at 80℃ for 30 minutes to completely evaporate the ethyl acetate solvent, then raise the temperature to 120℃ and hold for 2 hours to allow the polyorganosiloxane molecules to undergo a cross-linking and curing reaction, forming a continuous and dense surface modification layer. Figure 2 The component marked 5 in the middle. This layer forms covalent bonds with the carboxyl groups on the substrate surface through silanol groups, which increases the interfacial adhesion by more than 3 times.

[0029] Step 3: Buffer layer deposition: The sample obtained in step 2 was placed in the vacuum chamber of the magnetron sputtering apparatus, and the mechanical pump and molecular pump were turned on in sequence to evacuate to a background vacuum of ≤5×10⁻. 4 The higher the base vacuum, the lower the impurity content and the better the density of the deposited film. A silicon nitride (Si3N4) ceramic target is installed, the target-substrate distance is adjusted to 80 mm, argon gas is introduced as the sputtering gas, the flow rate is controlled at 30 sccm, the working pressure is adjusted to 1.5 × 10⁻³ Torr, and the DC sputtering power is set to 50 W. Pre-sputtering for 5 min removes the oxide layer and impurities on the target surface to prevent impurities from entering the film and affecting performance. Then, the baffle is opened and deposition is performed for 40 s to form a Si3N4 buffer layer with a thickness of approximately 20 nm on the surface modification layer. Figure 2 The component marked 6. This layer is used to match the difference in thermal expansion coefficients between the GF / PI substrate and the subsequent metal reflective layer, alleviate thermal mismatch stress during space thermal cycling, and also act as an atomic barrier layer to prevent metal atoms from diffusing into the substrate.

[0030] Step 4: Seed layer deposition: Maintaining the vacuum state of the vacuum chamber, the target material was changed to a NiCr alloy target (nickel to chromium mass ratio of 80:20). The argon flow rate was kept constant at 30 sccm, and the working pressure at 1.5 × 10⁻³ Torr. The DC sputtering power was adjusted to 30 W. After pre-sputtering for 3 minutes to remove impurities from the target surface, the baffle was opened, and deposition was carried out for 15 seconds. During deposition, the substrate was rotated at a uniform speed of 10 rpm to ensure that the film layer was uniformly distributed across the entire substrate surface, free of pinholes and defects. Finally, a NiCr seed layer with a thickness of approximately 5 nm was formed. Figure 2 The component marked with 7. The lattice constant of NiCr alloy is highly matched with that of metal reflective layers such as aluminum and silver, which can induce the metal grains to grow along a preferred orientation, significantly improving the smoothness and crystallinity of the metal reflective layer.

[0031] Step 5: Deposition of the metal reflective layer: The sample obtained in step 4 was transferred to an electron beam evaporation stage and evacuated to a background vacuum of ≤5×10⁻. 5 A higher vacuum level (Pa) reduces gas inclusions in the film, increasing its density. High-purity aluminum particles (99.999%) are loaded into a molybdenum boat. The boat is first heated with a small current to pre-melt the aluminum particles and held at that temperature for 5 minutes to remove adsorbed gaseous impurities and prevent splashing during the evaporation process. The electron beam power is adjusted to control the aluminum evaporation rate at 0.5 nm / s. Too fast an evaporation rate results in a porous film, reducing reflectivity and thermal conductivity; too slow a rate reduces production efficiency. The total evaporation time is 400 s, depositing an aluminum reflective layer approximately 200 nm thick at room temperature. Figure 2 The component marked with 8 is also the core functional layer of the thermal control film. This layer reflects more than 92% of sunlight, reducing the solar absorptivity, and rapidly diffuses the heat generated by the battery array laterally, eliminating localized hot spots.

[0032] Step 6: Deposition of anti-oxidation layer: The sample obtained in step 5 was transferred back to the magnetron sputtering instrument, the NiCr alloy target was reinstalled, argon gas was introduced, the flow rate was controlled at 30 sccm, the working pressure was adjusted to 2.0 × 10⁻³ Torr, and the DC sputtering power was set to 80 W. After pre-sputtering for 3 min, the baffle was opened, and deposition was carried out for 250 s, forming a NiCr anti-oxidation layer with a thickness of approximately 500 nm on the surface of the aluminum reflective layer. Figure 2 The component marked 9 completes the fabrication of the thermal control film composite structure. When exposed to the space environment, this layer will rapidly form a dense Cr2O3 passivation film on its surface, effectively blocking the erosion of the internal aluminum reflective layer by atomic oxygen and ultraviolet radiation from low Earth orbit.

[0033] Step 7: Integration with flexible solar cell array: The thermal control film composite structure obtained in step 6 was cut to a size that perfectly matches the HJT flexible solar cell array. A layer of 3MDP460 epoxy adhesive was uniformly coated onto the GF / PI substrate surface of the thermal control film, controlling the adhesive layer thickness to be between 20 and 30 μm. Too thin an adhesive layer would result in poor adhesion, while too thick a layer would increase thermal resistance and reduce heat dissipation efficiency. Figure 1 The layering sequence shown is as follows: the thermal control film layer is aligned and bonded to the back of the HJT flexible solar cell array, and the interface air bubbles are gently pressed out. Then, it is placed in a vacuum press and kept at 0.5MPa pressure and 60℃ temperature for 4 hours to allow the epoxy adhesive to fully cross-link and cure, thus obtaining the composite structure solar cell array.

[0034] Performance testing: The thermal control film composite structure and the composite structure solar cell array prepared in this embodiment were subjected to comprehensive performance tests. The meanings of each index and the test results are as follows: Optical properties: Solar absorptivity αs refers to the ratio of solar radiation energy absorbed by the material to the incident solar radiation energy; the lower the value, the less sunlight is absorbed. Hemispherical emissivity ε refers to the ratio of infrared radiation energy emitted by the material into the hemispherical space to the energy emitted by a blackbody at the same temperature; the higher the value, the stronger the outward heat dissipation capacity. In this embodiment, the sample has a solar absorptivity αs = 0.08 in the 250~2500nm solar spectrum and a hemispherical emissivity ε = 0.90 in the 8~14μm infrared spectrum.

[0035] Heat dissipation efficiency: Simulates the low-temperature vacuum environment of space in a thermal vacuum test chamber (vacuum degree 1×10⁻). 5 The same thermal power (Pa, heat sink temperature -173℃) was applied to the sample, and the steady-state temperature was tested and compared with that of conventional Z-93 white paint. The results show that the heat dissipation efficiency of the thermal control film layer in this embodiment is 1.2 times that of conventional Z-93 white paint.

[0036] On-orbit temperature: The sample was carried aboard a remote sensing microsatellite at an altitude of 500 km and an inclination of 97°, and the average surface temperature of the battery array in the sunlit area was collected by the on-orbit telemetry system. The results showed that the average temperature of the battery array in this embodiment was 58.4°C, which was 24°C lower than the control group without the composite thermal control film.

[0037] Conversion efficiency: Under standard test conditions (AM1.5G spectrum, 1000W / m² irradiance, 25℃), the conversion efficiency of the solar array was tested using a solar simulator. The results show that after the temperature decreased, the conversion efficiency of the HJT cell increased from 25.0% to 27.5%, an increase of 10%.

[0038] Environmental adaptability: The thermal cycling test was conducted according to aerospace standards, involving 1000 cycles from -180℃ to +120℃. After the test, the film adhesion retention rate was >98%, with no cracking or peeling. The atomic oxygen irradiation test simulated the atomic oxygen environment in low Earth orbit, with an irradiation flux of 1×10⁻⁶. 16 The solar absorptivity changed by 0.01 after testing, and the thermal control performance remained stable.

[0039] To verify the universality of the parameter range of this invention, comparative samples with substrate thicknesses of 0.1 mm and 0.5 mm, aluminum reflective layer thicknesses of 180 nm and 220 nm, and anti-oxidation layer thicknesses of 500 nm, 750 nm, and 1 μm were prepared and their performance tested according to the same method described above. The results showed that the solar absorptivity of all comparative samples was ≤0.10, and the hemispherical emissivity was >0.88; after 1000 cycles of thermal cycling from -180℃ to +120℃, the film adhesion retention rate was >95%; and after an atomic oxygen flux of 1×10¹ 6 After irradiation with atoms / cm², the change in solar absorptivity was less than 0.02, proving that the parameter range defined in this invention can stably achieve the technical effect described above.

[0040] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A silicon-based thermal control film composite structure for commercial aerospace, characterized in that, From the composite surface with the battery array to the exposed surface, it sequentially includes a glass fiber reinforced polyimide substrate, a surface modification layer, a buffer layer, a seed layer, a metal reflective layer, and an anti-oxidation layer; The surface modification layer is a polyorganosiloxane with a thickness of 0.1~20μm; The buffer layer is a nitride or oxide with a thickness of 10~100nm; The seed layer is a NiCr alloy or Al2 / 3O. 1-x The thickness is 2~10nm; The metal reflective layer is made of aluminum or silver and has a thickness of 100~500 nm. The anti-oxidation layer is a NiCr alloy with a thickness of 500nm~1μm.

2. The thermal control film composite structure according to claim 1, characterized in that, The glass fiber reinforced polyimide substrate has a thickness of 0.1~0.5mm, a glass transition temperature >250℃, and a coefficient of thermal expansion in the Z-axis direction <68.46×10⁻⁻⁻⁶. 6 / ℃, tensile strength >150MPa.

3. The thermal control film composite structure according to claim 1, characterized in that, The buffer layer is any one of silicon nitride, aluminum nitride, titanium dioxide, or zinc oxide.

4. The thermal control film composite structure according to claim 1, characterized in that, The mass ratio of Ni:Cr alloy in the seed layer is Ni:Cr = 80:

20.

5. The thermal control film composite structure according to claim 1, characterized in that, The metal reflective layer is made of aluminum and has a thickness of 180~220nm.

6. The thermal control film composite structure according to claim 1, characterized in that, The thermal control film composite structure has a solar absorptivity αs≤0.10 in the 250~2500nm solar spectrum and a hemispherical emissivity ε>0.88 in the 8~14μm infrared spectrum.

7. The thermal control film composite structure according to claim 1, characterized in that, After undergoing 1000 cycles of thermal cycling from -180℃ to +120℃, the adhesion retention rate of the thermal control film composite structure is >95%. Atomic oxygen flux 1×10¹ 6 After irradiation with atoms / cm², the change in solar absorptivity is <0.

02.

8. A composite solar cell array for commercial spaceflight, characterized in that, The invention includes a silicon-based flexible solar cell array body and a thermal control film composite structure as described in any one of claims 1 to 7; one side of the glass fiber reinforced polyimide substrate of the thermal control film composite structure is bonded to the back side of the silicon-based flexible solar cell array body by an epoxy adhesive with a temperature resistance of not less than 200°C, and the anti-oxidation layer side of the thermal control film composite structure is a space-exposed surface.

9. The composite structure solar cell array according to claim 8, characterized in that, The silicon-based flexible solar cell array body is either a large-area heterojunction (HJT) flexible solar cell array or a perovskite-silicon tandem flexible solar cell array.

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