Solar cell array for spacecraft and preparation method and application thereof

Abstract

The invention provides a solar cell array for a spacecraft and a preparation method and application of the solar cell array for the spacecraft, the solar cell array for the spacecraft comprises a flexible substrate, a solar cell module and an adhesive tape, the solar cell module and the adhesive tape are arranged on the surfaces of the two sides of the flexible substrate respectively, and the flexible substrate is provided with at least one through hole. The two sides of the adhesive tape comprise an adhesive layer and a non-adhesive layer, and the adhesive layer of the adhesive tape is adhered to the solar cell module through the through hole in the flexible substrate. According to the solar cell array for the spacecraft, a bonding layer does not need to be arranged between the solar cell module and the substrate, so that the working procedures of long-time pressurization and standing curing of a traditional adhesive do not exist, and the production efficiency is greatly improved; as no adhesive is needed, precise and expensive gluing equipment is not needed, and the production cost is greatly reduced.

Description

Technical Field

[0001] The invention belongs to the technical field of aerospace materials and relates to a solar cell array for a spacecraft and a preparation method and application thereof. Background Art

[0002] Solar cell arrays use the photoelectric effect to convert sunlight into electrical energy and are the main energy source for spacecraft. Low-orbit large-scale constellations (such as the Starlink constellation) are a hot spot for current aerospace industry applications, and large-scale production has put forward higher requirements for the efficiency and cost of solar cell array production. Although the production efficiency of spacecraft solar cell arrays, especially large-area flexible solar cell arrays, can be improved by using automated equipment under the current process and preparation methods, it still cannot meet the requirements of large-scale constellations for low cost and high production efficiency of solar cell arrays.

[0003] CN109671789A discloses a gallium arsenide thin film solar cell array for space use and a preparation method thereof, the preparation method comprising the following steps: 1) coating liquid silicone rubber on the surface of a flexible substrate; 2) laying a thin film solar cell string on the corresponding uncured silicone rubber surface, and performing large-area vacuum pressure curing; 3) coating liquid transparent silicone rubber on the surface of the thin film battery array area after vacuum pressure curing; 4) laying a flexible transparent composite film on the corresponding uncured liquid transparent silicone rubber surface, and performing vacuum pressure curing; 5) finally, scrubbing the overflowed glue and performing performance testing after curing to complete the preparation of the flexible thin film solar cell array.

[0004] CN116960229A discloses a method for preparing a large-area spatial fully flexible solar cell array module, which comprises: 1. preparing a flexible polyimide substrate reinforced with pre-embedded cables and a fiber grid; 2. assembling a solar cell array module; 3. coating a film adhesive on the flexible polyimide substrate reinforced with pre-embedded cables and a fiber grid, and placing the solar cell array module on the film adhesive; 4. preparing a mixture of silicone rubber and cerium-containing glass beads; 5. coating a mixture of silicone rubber and cerium-containing glass beads, and laminating it three times; 6. etching or mold pressing.

[0005] The integration method of the solar cell array described in the above scheme is to use an adhesive (cover adhesive) to stick a glass cover on the surface of the solar cell to make a laminated solar cell, and then apply an adhesive (base film adhesive) on the surface of the substrate to stick the laminated solar cell module on the surface of the substrate. The silicone rubber used in these two processes is space-grade low-volatile room temperature vulcanized silicone rubber, which has a long curing time and needs to be pressurized and left to stand during curing. In addition, the silicone rubber stirring and coating process may introduce bubbles, bringing reliability risks. In order to reduce the reliability risk, additional defect monitoring, negative pressure degassing and other process operations are required, which further increases the cost and cycle.

[0006] If the mass production of flexible solar cell arrays still uses the traditional solar cell array integration method, the cost and production cycle of the solar cell array will be difficult to reduce. Therefore, it is particularly important to optimize the integration method of solar cells and substrates to further reduce the preparation cycle and cost of solar cell arrays. Summary of the invention

[0007] The object of the present invention is to provide a solar cell array for spacecraft, and a preparation method and application thereof. In the solar cell array for spacecraft described in the present invention, there is no need to set a bonding layer between the solar cell module and the substrate, so there is no process of long-term pressurization and static curing of traditional adhesives, and the preparation efficiency is greatly improved; since no adhesive is required, there is no need for precise and expensive glue coating equipment, which greatly reduces the production cost.

[0008] In order to achieve the purpose of the invention, the present invention adopts the following technical solutions:

[0009] In a first aspect, the present invention provides a solar cell array for a spacecraft, the solar cell array for a spacecraft comprising a flexible substrate and solar cell components and tape respectively arranged on both side surfaces of the flexible substrate, at least one through hole being arranged on the flexible substrate, the two sides of the tape comprising an adhesive layer and a non-adhesive layer, and the adhesive layer of the tape is adhered to the solar cell components through the through hole on the flexible substrate.

[0010] In the solar cell array for spacecraft of the present invention, there is no adhesive layer between the solar cell module and the flexible substrate, which avoids the process of long-term pressurization and static curing of traditional adhesives, greatly shortens production time and production cost, and improves production efficiency. The adhesive tape is set on the flexible substrate, on the other side of the solar cell module, which is more convenient when replacement and maintenance are required. In traditional solar cell arrays, expensive space-grade low-volatile adhesives need to be set between the solar cell module and the substrate. The adhesive tape of the present invention is low-cost and does not require the use of precise and expensive glue coating equipment, which greatly reduces the cost of materials and equipment.

[0011] Preferably, the flexible substrate includes any one of a polyimide flexible substrate, a polyimide-glass fiber reinforced flexible substrate or an FPC flexible substrate, or a combination of at least two thereof. Typical but non-limiting combinations include a combination of a polyimide flexible substrate and an FPC flexible substrate, a combination of a polyimide flexible substrate and a polyimide-glass fiber reinforced flexible substrate, or a combination of a polyimide-glass fiber reinforced flexible substrate and an FPC flexible substrate, etc.

[0012] Preferably, the thickness of the flexible substrate is 0.1 mm to 0.25 mm, for example, 0.1 mm, 0.15 mm, 0.18 mm, 0.2 mm, 0.22 mm or 0.25 mm, etc., and is not limited to the listed values, and other values ​​not listed within the numerical range are also applicable.

[0013] Preferably, the through hole comprises an open through hole.

[0014] Preferably, the shape of the open through hole includes any one of an X-shape, a cross shape, a M-shape, a spot shape, a W-shape or a T-shape, or a combination of at least two of them. Typical but non-limiting combinations include a combination of an X-shape and a T-shape, a combination of a W-shape and a T-shape, or a combination of an X-shape and a W-shape, etc.

[0015] The open through holes of the present invention can avoid the risk of forming a closed cavity between the through holes after the tape is attached to the closed through holes, thereby preventing the occurrence of on-track reliability problems.

[0016] Preferably, the solar cell assembly comprises at least two monolithic solar cells connected in series and / or in parallel.

[0017] Preferably, the monolithic solar cell includes a rigid solar cell and / or a flexible solar cell.

[0018] Preferably, the rigid solar cell comprises any one of a rigid triple-junction gallium arsenide solar cell, a rigid silicon solar cell or a rigid perovskite solar cell, or a combination of at least two thereof. Typical but non-limiting combinations include a combination of a rigid triple-junction gallium arsenide solar cell and a rigid perovskite solar cell, a combination of a rigid silicon solar cell and a rigid perovskite solar cell, or a combination of a rigid triple-junction gallium arsenide solar cell and a rigid silicon solar cell, etc.

[0019] Preferably, the flexible solar cell comprises any one of a flexible triple-junction gallium arsenide solar cell, a flexible silicon solar cell, a flexible perovskite solar cell or a flexible CIGS solar cell, or a combination of at least two thereof. Typical but non-limiting combinations include a combination of a flexible triple-junction gallium arsenide solar cell and a flexible CIGS solar cell, a combination of a flexible silicon solar cell and a flexible CIGS solar cell, or a combination of a flexible triple-junction gallium arsenide solar cell and a flexible silicon solar cell, etc.

[0020] Preferably, at least one open through hole is provided on the substrate below the monolithic solar cell.

[0021] Preferably, taking the projected area of ​​the monolithic solar cell on the flexible substrate as 100%, the area of ​​the monolithic solar cell in the vertical direction of the open through hole is 10% to 70%, for example, 10%, 15%, 30%, 45% or 50%, etc., not limited to the listed values, and other unlisted values ​​within the numerical range are also applicable.

[0022] The area of ​​the single solar cell in the vertical direction of the open through hole of the present invention is the area covered by the single solar cell on the open through hole.

[0023] In the solar cell array for spacecraft of the present invention, each single solar cell may correspond to one through hole, or one single solar cell may correspond to multiple through holes, and each single solar cell only needs to be adhered to the tape through the through hole.

[0024] Preferably, the series gap of the monolithic solar cell is 1 mm to 1.5 mm, for example, 1 mm, 1.1 mm, 1.2 mm, 1.3 mm, 1.4 mm or 1.5 mm, etc., and is not limited to the listed values, and other values ​​not listed within the numerical range are also applicable.

[0025] Preferably, the parallel gap of the monolithic solar cell is 0.75 mm to 2 mm, for example, 0.75 mm, 0.8 mm, 1 mm, 1.5 mm or 2 mm, etc., and is not limited to the listed values, and other values ​​not listed within the numerical range are also applicable.

[0026] Preferably, the non-adhesive layer of the tape comprises a polyimide substrate.

[0027] Preferably, the adhesive layer of the tape comprises a silicone adhesive.

[0028] Preferably, the silicone adhesive comprises a pressure-sensitive solid adhesive.

[0029] The pressure-sensitive solid adhesive of the present invention is prepared by condensing silicone rubber and MQ resin, wherein the MQ resin is composed of tetrafunctional siloxane segments (SiO 4 / 2 , Q) and a monofunctional siloxane chain (R 3 SiO 1 / 2 , M) through co-hydrolysis-condensation reaction of organosilicon compounds to generate a three-dimensional spherical structure silicone grease. Compared with solid adhesives such as acrylic acid, silicone adhesives have excellent high-temperature stability and can ensure normal operation in the harsh thermal environment of space.

[0030] Preferably, the thickness of the polyimide substrate is 0.02 mm to 0.05 mm, for example, 0.02 mm, 0.025 mm, 0.03 mm, 0.04 mm or 0.05 mm, etc., and is not limited to the listed values, and other values ​​not listed within the numerical range are also applicable.

[0031] Preferably, the total thickness of the tape is 0.05 mm to 0.1 mm, for example, 0.05 mm, 0.06 mm, 0.07 mm, 0.08 mm, 0.09 mm or 0.1 mm, etc., and is not limited to the listed values, and other unlisted values ​​within the numerical range are also applicable.

[0032] Preferably, the initial tack of the adhesive tape should be no less than 28# (GB / T 4852).

[0033] Preferably, the tape completely covers the through hole.

[0034] Preferably, the area of ​​the tape is 30% to 70% larger than the area of ​​the open through hole, for example, 30%, 40%, 50%, 60% or 70%, etc., and is not limited to the listed values, and other unlisted values ​​within the numerical range are also applicable.

[0035] In a second aspect, the present invention provides a method for preparing a solar cell array for a spacecraft as described in the first aspect, the preparation method comprising the following steps:

[0036] (1) performing a hole-opening process on the flexible substrate to form a flexible substrate provided with through holes, cleaning the surface of the substrate, and pasting a release film on one side surface of the flexible substrate;

[0037] (2) Paste the tape on the other side of the release film of the flexible substrate so that the adhesive layer of the tape is in full contact with the release film through the through hole;

[0038] (3) removing the release film, pasting the solar cell assembly to the position of the through hole on one side of the original release film, and applying pressure with an air bag to ensure that the adhesive layer of the tape is fully adhered to the solar cell assembly, thereby obtaining the solar cell array for the spacecraft.

[0039] The pressurization treatment by airbags of the present invention is to pressurize both the tape side and the solar cell module side at the same time, wherein the airbag on the tape side is a fixed airbag that plays a role of buffering and filling, and the airbag on the solar cell module side is a pressure-applying side.

[0040] The release film of the present invention can prevent the contamination or damage of the solar cell adhesive surface during subsequent operations, and can also prevent the back tape from making unnecessary contact with other surfaces to affect the adhesive force.

[0041] Preferably, the release film in step (1) comprises a transparent polyethylene terephthalate substrate light release film.

[0042] Preferably, the release force of the release film in step (1) is ≤20g / 25mm.

[0043] The release force test process of the present invention uses a 1-inch tape to perform the release force test, so the unit is g / 25mm (specifically refer to GB / T 25256, 180° peel force).

[0044] Preferably, the pressure of the pressurization treatment in step (3) is 1 kPa to 10 kPa, for example, 1 kPa, 2 kPa, 5 kPa, 8 kPa or 10 kPa, etc., and is not limited to the listed values. Other values ​​not listed within the numerical range are also applicable.

[0045] Preferably, the time of the pressurization treatment in step (3) is 8s to 15s, for example, 8s, 9s, 10s, 11s, 12s, 13s, 14s or 15s, etc., and is not limited to the listed values. Other values ​​not listed within the numerical range are also applicable.

[0046] In a third aspect, the present invention provides a spacecraft, wherein the spacecraft comprises the solar cell array for spacecraft as described in the first aspect.

[0047] Compared with the prior art, the present invention has the following beneficial effects:

[0048] (1) In the solar cell array for spacecraft described in the present invention, there is no need to provide an adhesive layer between the solar cell module and the substrate, and therefore there is no need for the traditional adhesive to be subjected to long-term pressurization and static curing, thereby greatly improving the preparation efficiency. Since no adhesive is required, there is no need for sophisticated and expensive glue coating equipment, thereby greatly reducing the production cost.

[0049] (2) The structure of the solar cell array for spacecraft described in the present invention is simple. Due to the special structure, the tape is arranged on the flexible substrate, on the opposite side of the solar cell assembly, which is more convenient when the sheet needs to be replaced for maintenance.

[0050] (3) The integration method described in the present invention can shorten the integration time of the solar cell array (the entire process from series welding of components to completion of component pasting) to about 1 hour, which is more than 20 hours shorter than the conventional integration method. The cost in large-scale production can be reduced by 15% to 25%, and the production efficiency can be increased by more than 10 times.

[0051] (4) The solar cell array for spacecraft of the present invention can significantly shorten the preparation time while ensuring the power, and when rigid solar cells are used, the mass surface density can reach 2.05 kg / m2 By adjusting the amount of various materials and the preparation process, the mass surface density of the solar cell array for spacecraft can reach 1.79kg / m 2 The thickness is within 0.7mm and the finished product is in normal condition, with no debonding or cracking after the mechanical vibration test. BRIEF DESCRIPTION OF THE DRAWINGS

[0052] Figure 1 It is a schematic diagram of the structure of a solar cell array for a spacecraft according to Example 1 of the present invention, wherein 1 is a polyimide flexible substrate, 2 is an X-shaped through hole, 3 is a single-chip solar cell, and 4 is an adhesive tape.

[0053] Figure 2 It is a schematic diagram of the structure of a solar cell array for a spacecraft according to Example 2 of the present invention, wherein 5 is a polyimide flexible substrate, 6 is a triangular through hole, 7 is a single-chip solar cell, and 8 is an adhesive tape.

[0054] Figure 3 It is a schematic diagram of the structure of a solar cell array for a spacecraft according to Example 3 of the present invention, wherein 9 is a polyimide flexible substrate, 10 is a W-shaped through hole, 11 is a single-chip solar cell, and 12 is an adhesive tape. DETAILED DESCRIPTION

[0055] The technical solution of the present invention is further described below by specific implementation methods. It should be understood by those skilled in the art that the embodiments are only used to help understand the present invention and should not be regarded as specific limitations of the present invention.

[0056] Example 1

[0057] The present embodiment provides a solar cell array for a spacecraft, wherein the solar cell array for a spacecraft uses a polyimide flexible substrate with a thickness of 0.2 mm, an X-shaped through hole is provided on the polyimide flexible substrate, a solar cell assembly is arranged on one side of the polyimide flexible substrate, a single solar cell in the solar cell assembly is a rigid triple-junction gallium arsenide solar cell, a series gap of a single solar cell in the solar cell assembly is 1 mm, and a parallel gap is 1 mm, the solar cell assembly is bonded to an adhesive layer of an adhesive tape through the X-shaped through hole, the adhesive layer of the adhesive tape is a pressure-sensitive solid adhesive, a base material of the adhesive tape is polyimide, a thickness of the base material is 0.03 mm, a total thickness of the adhesive tape is 0.06 mm, a total area of ​​the through hole is 40% of an area of ​​the solar cell assembly, the adhesive tape completely covers the through hole, and an area of ​​the adhesive tape is 50% larger than an area of ​​the through hole.

[0058] The schematic diagram of the structure of the solar cell array for the spacecraft is as follows Figure 1 As shown:

[0059] The polyimide flexible substrate 1 is provided with an X-shaped through hole 2, and a single solar cell 3 covers the X-shaped through hole 2 and is bonded to a tape 4 on the back (the tape is on the back, and the tape and the single solar cell are separated to show the structure).

[0060] The solar cell array for spacecraft is manufactured by the following method:

[0061] (1) performing a hole-opening process on a polyimide flexible substrate to form a polyimide flexible substrate provided with an X-shaped through hole, cleaning the surface of the substrate, and pasting a transparent polyethylene terephthalate substrate light release film on the front side of the polyimide flexible substrate;

[0062] (2) Clean the back of the substrate and stick a tape on the back of the polyimide flexible substrate release film so that the adhesive layer of the tape is in full contact with the transparent polyethylene terephthalate substrate light release film through the through hole;

[0063] (3) The transparent polyethylene terephthalate substrate light release film is removed, and the solar cell assembly is adhered to the position of the through hole on the front side. A pressure of 5 kPa is applied on one side of the tape by an airbag for 10 seconds to ensure that the adhesive layer of the tape is fully adhered to the solar cell assembly, thereby obtaining the solar cell array for the spacecraft.

[0064] Example 2

[0065] The present embodiment provides a solar cell array for spacecraft, the solar cell array for spacecraft uses a polyimide-glass fiber reinforced flexible substrate with a thickness of 0.15 mm, the polyimide-glass fiber reinforced flexible substrate is provided with a three-shaped through hole, a solar cell module is arranged on one side of the polyimide-glass fiber reinforced flexible substrate, the single solar cell in the solar cell module is a rigid triple-junction gallium arsenide solar cell, the series gap of the single solar cell in the solar cell module is 1 mm, and the parallel gap is 0.75 mm, the solar cell module is bonded to the adhesive layer of the tape through the three-shaped through hole, the adhesive layer of the tape is a pressure-sensitive solid adhesive, the base material of the tape is polyimide, the thickness of the base material is 0.02 mm, the total thickness of the tape is 0.05 mm, the total area of ​​the through hole is 30% of the area of ​​the solar cell module, the tape completely covers the through hole, and the area of ​​the tape is 60% larger than the area of ​​the through hole.

[0066] The schematic diagram of the structure of the solar cell array for the spacecraft is as follows Figure 2 As shown:

[0067] The polyimide flexible substrate 5 is provided with a triangular through hole 6, and a single solar cell 7 covers the triangular through hole 6 and is bonded to a tape 8 on the back (the tape is on the back, and the tape and the single solar cell are separated to show the structure).

[0068] The solar cell array for spacecraft is manufactured by the following method:

[0069] (1) Performing a hole-opening process on a polyimide-glass fiber reinforced flexible substrate to form a polyimide flexible substrate with three-shaped through holes, cleaning the surface of the substrate, and pasting a transparent polyethylene terephthalate substrate light release film on the front side of the polyimide-glass fiber reinforced flexible substrate;

[0070] (2) Clean the back of the substrate and stick a tape on the back of the polyimide-glass fiber reinforced flexible substrate release film so that the adhesive layer of the tape is in full contact with the transparent polyethylene terephthalate substrate light release film through the through hole;

[0071] (3) The transparent polyethylene terephthalate substrate light release film is removed, and the solar cell assembly is adhered to the position of the front through hole. A pressure of 1 kPa is applied on one side of the tape by an airbag for 15 seconds to ensure that the adhesive layer of the tape is fully adhered to the solar cell assembly, thereby obtaining the solar cell array for the spacecraft.

[0072] Example 3

[0073] The present embodiment provides a solar cell array for a spacecraft, wherein the solar cell array for a spacecraft uses an FPC flexible substrate with a thickness of 0.25 mm, the FPC flexible substrate is provided with a Chinese character "W"-shaped through hole, a solar cell module is arranged on one side of the FPC flexible substrate, a single solar cell in the solar cell module is a rigid triple-junction gallium arsenide solar cell, a series gap of a single solar cell in the solar cell module is 1.5 mm, and a parallel gap is 2 mm, the solar cell module is bonded to an adhesive layer of an adhesive tape through the Chinese character "W"-shaped through hole, the adhesive layer of the adhesive tape is a pressure-sensitive solid adhesive, a base material of the adhesive tape is polyimide, a thickness of the base material is 0.05 mm, a total thickness of the adhesive tape is 0.1 mm, a total area of ​​the through hole is 50% of an area of ​​the solar cell module, the adhesive tape completely covers the through hole, and an area of ​​the adhesive tape is 70% larger than an area of ​​the through hole.

[0074] The schematic diagram of the structure of the solar cell array for the spacecraft is as follows Figure 3 As shown:

[0075] The polyimide flexible substrate 9 is provided with a king-shaped through hole 10, and a single-piece solar cell 11 is covered on the king-shaped through hole 10 and bonded to the tape 12 on the back (the tape is on the back, and for showing the structure, the tape and the single-piece solar cell are separated).

[0076] The solar cell array for spacecraft is prepared by the following method:

[0077] (1) Perform an opening process on the FPC flexible substrate to form a polyimide flexible substrate provided with king-shaped through holes, clean the surface of the substrate, and paste a transparent polyethylene terephthalate base material light release film on the front surface of the FPC flexible substrate;

[0078] (2) Clean the back surface of the substrate, paste a tape on the back side of the release film of the FPC flexible substrate, and make the adhesive layer of the tape fully contact the transparent polyethylene terephthalate base material light release film through the through holes;

[0079] (3) Remove the transparent polyethylene terephthalate base material light release film, paste the solar cell assembly at the position of the front through holes, apply a pressure of 10 kPa through an airbag on one side of the tape for 8 s for pressure treatment, so that the adhesive layer of the tape is fully adhered to the solar cell assembly, and obtain the solar cell array for spacecraft.

[0080] Example 4

[0081] The difference between this example and Example 1 is only that the thickness of the flexible substrate is 0.05 mm, and other conditions and parameters are exactly the same as those in Example 1.

[0082] Example 5

[0083] The difference between this example and Example 1 is only that the thickness of the flexible substrate is 0.4 mm, and other conditions and parameters are exactly the same as those in Example 1.

[0084] Example 6

[0085] The difference between this example and Example 1 is only that the area of the single-piece solar cell in the vertical direction of the through hole is 5% of the area of the single-piece solar cell, and other conditions and parameters are exactly the same as those in Example 1.

[0086] Example 7

[0087] The difference between this example and Example 1 is only that the area of the single-piece solar cell covering the through hole is 80% of the area of the single-piece solar cell, and other conditions and parameters are exactly the same as those in Example 1.

[0088] Example 8

[0089] The only difference between this embodiment and embodiment 1 is that the pressure of the pressurization treatment is 0.5 kPa, and the other conditions and parameters are exactly the same as those in embodiment 1.

[0090] Example 9

[0091] The only difference between this embodiment and embodiment 1 is that the pressure of the pressurization treatment is 20 kPa, and the other conditions and parameters are exactly the same as those in embodiment 1.

[0092] Example 10

[0093] The only difference between this embodiment and embodiment 1 is that the time of the pressure treatment is 5 seconds, and the other conditions and parameters are exactly the same as those in embodiment 1.

[0094] Embodiment 11

[0095] The only difference between this embodiment and embodiment 1 is that the time of the pressurization treatment is 20 seconds, and the other conditions and parameters are exactly the same as those in embodiment 1.

[0096] Comparative Example 1

[0097] This comparative example adopts the traditional spacecraft solar cell array integration method. First, the substrate surface is cleaned, and liquid silicone rubber of a specific shape is sprayed or printed on the surface of the flexible substrate to ensure that the silicone rubber shape under each solar cell is complete and the thickness (0.1mm) is uniform. Then the solar cell module is pasted on top of the silicone rubber shape, and the solar cell module is pressurized with a sandbag, a soft heavy object or an air bag for 1h~2h. The pressurizing object is removed and the module is left to stand for 12h~24h. It is then placed in a low-pressure, high-temperature device for curing for 6h~12h.

[0098] Performance Test:

[0099] For a block with an area of ​​1m 2 The solar panels of different embodiments are compared. 360 pieces of Ge substrate triple-junction GaAs solar cells with 30% efficiency, 0.35mm thickness and 40mm×60mm size are attached to the substrate surface. The test items include the time consumption of the solar cell array integration process, the weight and thickness of the finished product, the power and the mechanical vibration test. The test results are shown in Table 1:

[0100] Table 1

[0101]

[0102] It can be seen from Table 1 that from Examples 1-11, the solar cell array for spacecraft of the present invention can greatly shorten the preparation time while ensuring power, and the mass can be within 2.05 kg and the thickness can be within 0.81 mm. By adjusting the amount of various materials and the preparation process, the mass of the solar cell array for spacecraft can be within 1.79 kg and the thickness can be within 0.7 mm, and the finished product is in normal condition, with no debonding or cracking after the mechanical vibration test.

[0103] By comparing Example 1 with Examples 4-5, it can be seen that in the solar cell array for spacecraft of the present invention, the thickness of the flexible substrate will affect its effect. When the thickness of the flexible substrate is controlled within the range of 0.1 mm to 0.25 mm, the effect of the solar cell array for spacecraft is better. If the thickness of the flexible substrate is too small, the substrate bearing capacity is limited and there is a risk of solar cell fragmentation. If the thickness of the flexible substrate is too large, the solar panel is heavier and the envelope size is larger, thereby increasing the launch cost of the spacecraft.

[0104] By comparing Example 1 with Examples 6-7, it can be seen that in the solar cell array for spacecraft described in the present invention, the total area of ​​the solar cell components covering the through holes accounts for an area ratio in the solar cell components that affect its effect. When the area ratio of the solar cell components covering the through holes is controlled to be 30%-50%, the effect of the solar cell array for spacecraft is better. If the area ratio of the solar cell components covering the through holes is too small, insufficient adhesion may cause the solar cell panels to fall off. If the area ratio of the solar cell components covering the through holes is too large, the back tape will overlap and the bearing capacity of the substrate structure will be reduced.

[0105] By comparing Example 1 with Examples 8-9, it can be seen that during the preparation process of the solar cell array for spacecraft according to the present invention, the pressure of the pressurization treatment will affect its effect. When the pressure of the pressurization treatment is controlled at 1 kPa~10 kPa, the effect of the solar cell array for spacecraft is better. If the pressure of the pressurization treatment is too small, the tape is not fully pressed against the substrate and the solar cell sheet, and the bonding force is insufficient. If the pressure of the pressurization treatment is too large, the risk of the solar cell sheet breaking may increase.

[0106] By comparing Example 1 with Examples 10-11, it can be seen that in the preparation process of the solar cell array for spacecraft of the present invention, the time of pressurization treatment will affect its effect. When the time of pressurization treatment is controlled within 8s~15s, the effect of the solar cell array for spacecraft is better. If the time of pressurization treatment is too long, after the tape has been bonded, the extra pressurization time will only reduce the production efficiency without any practical effect, and increase the risk of breakage of the solar cell. If the time of pressurization treatment is too short, the tape is not fully pressed against the substrate and the solar cell, and the bonding force is insufficient.

[0107] By comparing Example 1 and Comparative Example 1, it can be seen that in the solar cell array for spacecraft described in the present invention, there is no need to set a bonding layer between the solar cell module and the substrate, so there is no process of long-term pressurization and static curing of traditional adhesives, and the preparation efficiency is greatly improved. Since no adhesive is required, the quality and thickness are significantly reduced, and there is no need for precise and expensive glue coating equipment. Without affecting the photoelectric conversion efficiency and reliability, the production cost is greatly reduced.

[0108] The applicant declares that the above is only a specific implementation mode of the present invention, but the protection scope of the present invention is not limited thereto. Those skilled in the art should understand that any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present invention shall fall within the protection scope and disclosure scope of the present invention.

Claims

1. A solar cell array for a spacecraft, characterized in that: The solar cell array for spacecraft includes a flexible substrate and solar cell components and tapes respectively arranged on the two side surfaces of the flexible substrate, at least one through hole is arranged on the flexible substrate, and the two sides of the tape include a sticky layer and a non-sticky layer, and the sticky layer of the tape is adhered to the solar cell components through the through holes on the flexible substrate.

2. The solar cell array for a spacecraft as claimed in claim 1, characterized in that: The flexible substrate includes any one of a polyimide flexible substrate, a polyimide-glass fiber reinforced flexible substrate or an FPC flexible substrate, or a combination of at least two thereof; And / or, the thickness of the flexible substrate is 0.1 mm to 0.25 mm.

3. The solar cell array for a spacecraft as claimed in claim 1, characterized in that: The through hole comprises an open through hole; And / or, the shape of the open through hole includes any one of an X-shape, a cross-shape, a 'M'-shape, a spot-shape, a 'K'-shape or a triangle-shape, or a combination of at least two of them.

4. The solar cell array for a spacecraft as claimed in claim 1, characterized in that: The solar cell assembly comprises at least two single-chip solar cells connected in series and / or in parallel; And / or, the monolithic solar cell comprises a rigid solar cell and / or a flexible solar cell; And / or, the rigid solar cell comprises any one or a combination of at least two of a rigid triple-junction gallium arsenide solar cell, a rigid silicon solar cell or a rigid perovskite solar cell; And / or, the flexible solar cell comprises any one or a combination of at least two of a flexible triple-junction gallium arsenide solar cell, a flexible silicon solar cell, a flexible perovskite solar cell or a flexible CIGS solar cell; And / or, at least one open through hole is provided at the substrate below the monolithic solar cell; And / or, taking the projection area of ​​the single solar cell on the flexible substrate as 100%, the area of ​​the single solar cell in the vertical direction of the open through hole is 10% to 70%.

5. The solar cell array for a spacecraft as claimed in claim 4, characterized in that: The series gap of the monolithic solar cell is 1 mm to 1.5 mm; And / or, the parallel gap of the single solar cell is 0.75 mm to 2 mm.

6. The solar cell array for a spacecraft as claimed in claim 1, characterized in that: The non-adhesive layer of the tape includes a polyimide substrate; and / or, the adhesive layer of the tape comprises a silicone adhesive; and / or, the silicone adhesive comprises a pressure-sensitive solid adhesive; And / or, the thickness of the polyimide substrate is 0.02 mm to 0.05 mm; And / or, the total thickness of the adhesive tape is 0.05 mm to 0.1 mm; and / or, the tape completely covers the through hole; And / or, the area of ​​the tape is 30% to 70% larger than the area of ​​the open through hole.

7. A method for preparing a solar cell array for a spacecraft as claimed in any one of claims 1 to 6, characterized in that: The preparation method comprises the following steps: (1) performing a hole-opening process on the flexible substrate to form a flexible substrate provided with through holes, cleaning the surface of the substrate, and pasting a release film on one side surface of the flexible substrate; (2) Paste the tape on the other side of the release film of the flexible substrate so that the adhesive layer of the tape is in full contact with the release film through the through hole; (3) removing the release film, pasting the solar cell assembly to the position of the through hole on one side of the original release film, and applying pressure with an air bag to ensure that the adhesive layer of the tape is fully adhered to the solar cell assembly, thereby obtaining the solar cell array for the spacecraft.

8. The preparation method according to claim 7, characterized in that: The release film in step (1) comprises a transparent polyethylene terephthalate substrate light release film; And / or, the release force of the release film in step (1) is ≤20g / 25mm.

9. The preparation method according to claim 7, characterized in that: The pressure of the pressurization treatment in step (3) is 1 kPa to 10 kPa; And / or, the time of the pressurization treatment in step (3) is 8s~15s.

10. A spacecraft, characterized in that: The spacecraft comprises the solar cell array for spacecraft according to any one of claims 1-6.

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