A planar asymmetric compound unrolled T-shaped space solar cell array

Abstract

This invention discloses a planar asymmetric composite unfolding T-shaped space solar array, comprising a connecting frame, an inner plate, a middle plate, an outer plate, an unfolding locking mechanism, a pressing and releasing mechanism, and a synchronous linkage device. The connecting frame is connected between the inner plate and the middle plate via the unfolding locking mechanism. The two sides of the middle plate are connected to the inner plate and the outer plate respectively via the unfolding locking mechanism to form the solar array. In the retracted state, the pressing and releasing mechanism provides a certain pressing force to the connecting frame, inner plate, middle plate, and outer plate. After the pressing and releasing mechanism releases the constraint on the solar array, the inner plate, middle plate, and outer plate unfold into a plane under the action of the unfolding locking mechanism and are locked first. The plane is parallel to the spacecraft body. The connecting frame completes the final unfolding and locking, and the connecting frame is perpendicular to the plane. At this time, the solar array has a T-shaped configuration. This invention has a small moment of inertia relative to the entire satellite after unfolding, a close center of mass distance from the entire satellite, and effectively solves the problem of complex unfolding trajectories.

Description

Technical Field

[0001] This invention relates to the field of spacecraft technology, specifically to a planar asymmetric composite unfolding T-shaped space solar cell array. Background Technology

[0002] Space solar arrays are the core power supply equipment for spacecraft. Most spacecraft employ symmetrical solar array designs, which helps maintain satellite attitude stability in space. However, some spacecraft carry special payloads, such as infrared detectors, which are extremely sensitive to temperature changes and require cooling to ensure detection accuracy. When sunlight shines on the solar array, it reflects infrared radiation, affecting satellite calibration accuracy and cooling performance. Therefore, a single solar array design is used in these cases.

[0003] Solar arrays are large, flexible components of spacecraft. Mechanical rotation of various payloads on the spacecraft can cause slight swaying of the solar array. This swaying, transmitted to the spacecraft, can affect attitude control stability, thereby reducing important functional performance such as Earth observation accuracy. The main way to reduce swaying interference is to reduce the moment of inertia of the swaying component relative to the spacecraft's center of mass (control center). Furthermore, during long-term stable operation, the large solar array area causes significant solar radiation pressure interference, which adversely affects attitude control stability. Therefore, efforts should be made to reduce the torque generated by the radiation pressure. This requires shortening the distance between the center of the deployed solar array area and the spacecraft's center of mass (control center), i.e., shortening the lever arm length, thereby reducing the radiation pressure interference torque. On the other hand, to ensure sufficient power generation, the solar array should be designed to have the largest possible area.

[0004] Traditional linear solar arrays have substrates connected in series, and their deployment is typically a one-dimensional linear process, simple and posing no risk of interference to surrounding equipment. However, T-shaped solar arrays are more complex, being two-dimensional structures. Their deployment cannot use the aforementioned one-dimensional linear method, requiring the design of novel deployment methods and path planning. Therefore, a planar asymmetric composite deployment method was designed for T-shaped solar arrays. For planar deployment of solar arrays, the first aspect is the need for rationally planning the deployment trajectory to ensure reliable and stable deployment; and the appropriate ratio of the number, position, and size of the drive springs, which affects the magnitude of the impact on the celestial body during deployment. The second aspect is the need to consider how to simulate the composite deployment process in ground-based deployment tests. To simulate the space environment, ground-based deployment tests need to simulate a microgravity environment. Due to the complexity of the deployment trajectory, the ground compensation method, the installation method on the solar array, and optimized layout are urgent problems to be solved. Summary of the Invention

[0005] In view of this, the present invention provides a planar asymmetric composite unfolding T-shaped space solar cell array, which has a small moment of inertia relative to the whole star after unfolding, a close center of mass distance from the whole star, and reduces the accumulation of solar radiation pressure during long-term stable operation, and effectively solves the problem of complex unfolding trajectory of planar unfolding solar cell arrays.

[0006] The technical solution adopted in this invention is as follows:

[0007] A planar asymmetric composite unfolding T-shaped spatial solar cell array includes a connecting frame, an inner solar panel, a middle solar panel, an outer solar panel, an unfolding locking mechanism, a pressing and releasing mechanism, and a synchronous linkage device.

[0008] The connecting frame is connected between the inner solar panel and the middle solar panel via an unfolding and locking mechanism. The two sides of the middle solar panel are connected to the inner solar panel and the outer solar panel via unfolding and locking mechanisms to form a solar cell array. In the compressed state, the solar cell array is folded up against the side wall of the spacecraft body. From the inside out, the components are the connecting frame, the inner solar panel, the middle solar panel, and the outer solar panel. In the folded state, the compression release mechanism provides a certain compression force to the connecting frame, the inner solar panel, the middle solar panel, and the outer solar panel. After the compression release mechanism releases the constraint on the solar cell array, the inner solar panel, the middle solar panel, and the outer solar panel unfold into a plane under the action of the unfolding and locking mechanism and lock it first. The plane is parallel to the spacecraft body. The connecting frame completes the unfolding and locks it last. The connecting frame is perpendicular to the plane. At this time, the solar cell array has a T-shaped configuration.

[0009] Furthermore, the deployment locking mechanism includes inter-plate hinges and offset inter-plate hinges, both of which are used in pairs;

[0010] The inter-plate hinges are respectively installed between the inner solar panel and the middle solar panel, between the middle solar panel and the outer solar panel, and between the connecting frame and the spacecraft body;

[0011] The offset plate hinge is located between the inner plate and the connecting frame.

[0012] Furthermore, the connecting frame uses two connecting rods, which are respectively connected to both sides of the inner plate, and the connecting frame and the inner plate form a triangle.

[0013] Furthermore, the offset plate hinge includes a locking hinge, a hook hinge, and a constant torque spring assembly;

[0014] The constant torque spring assembly is the driving source, and the locking hinge and hook hinge are both connected to the adapter, so that the connecting frame and the inner plate can rotate 90°.

[0015] Furthermore, the inner plate corner is provided with an integrated inner plate base plate reinforcement.

[0016] Furthermore, the synchronous linkage device includes a connecting frame linkage device and a middle plate linkage device;

[0017] The connecting frame linkage device is installed on the connecting frame to enable the connecting frame and the inner plate to unfold synchronously; the middle plate linkage device is installed on the side of the middle plate to enable the inner plate, middle plate and outer plate to unfold synchronously.

[0018] Furthermore, during ground-based deployment testing, the weight of the solar array is supported by a gravity compensation device installed on the array.

[0019] Furthermore, the gravity compensation device includes an air-floating platform and an air-floating support assembly;

[0020] One air-bearing support assembly is arranged on each of the connecting frame, the inner plate of the solar panel, and the middle plate of the solar panel, and two air-bearing support assemblies are arranged on the outer plate of the solar panel.

[0021] The air-bearing support assembly on the middle plate coincides with the center line of gravity of the solar panel middle plate; the air-bearing support assembly on the inner plate of the solar panel is spaced apart from the center line of gravity of the inner plate of the solar panel to avoid interference with the air-bearing support assembly on the middle plate of the solar panel during the unfolding process; the line connecting the support points of the two air-bearing support assemblies on the outer plate of the solar panel coincides with the center line of gravity of the outer plate of the solar panel; the air-bearing support assembly on the connecting frame is located on the end of the connecting frame away from the hinge of the plate.

[0022] The connecting frame, inner solar panel, middle solar panel, and outer solar panel are suspended on the air flotation platform under the action of the air flotation support assembly.

[0023] Beneficial effects:

[0024] 1. The solar array of this invention, after being deployed, takes on a T-shaped configuration, shifting the flat surface of the solar array outward by a certain distance. This ensures that the deployed configuration of the solar array does not obstruct other payloads or their field of view on the satellite. At the same time, while ensuring that the power generation and weight of the solar array meet the requirements, the solar array has a small moment of inertia relative to the entire satellite after deployment, and a large stiffness after deployment. This will not have an unacceptable impact on the attitude control of the entire satellite. The center of mass is close to the satellite body, and the accumulation of solar radiation pressure during long-term stable operation is reduced, ensuring the stable flight of the spacecraft in orbit. Furthermore, the deployment and locking of the solar array are achieved through the cooperation of the deployment locking mechanism and the synchronous linkage device.

[0025] 2. During the deployment of the solar cell array, the present invention utilizes a combination of "offset plate hinge + adapter" to lock the connecting frame and the inner solar panel at a 90° angle. Furthermore, the use of an adapter avoids interference between the offset plate hinge and the edges of the inner and middle solar panels during deployment, ensuring the smooth deployment of the solar cell array.

[0026] 3. The inner corner of the solar panel of the present invention is provided with an inner panel base plate reinforcement. Through the integrated design of the base plate reinforcement embedded part, the corner of the base plate has a certain rigidity and strength, and can bear large and mismatched bending loads.

[0027] 4. During the deployment of the solar cell array, the present invention utilizes the "connecting frame linkage device + middle plate linkage device" to drive the solar cell array to deploy in segments synchronously, which is beneficial for locking the hinge at the root of the solar cell array and reducing the impact of locking during the deployment of the solar cell array.

[0028] 5. This invention utilizes a gravity compensation device to simulate a zero-gravity environment, enabling planar deployment during ground deployment tests.

[0029] 6. In the deployment of the solar cell array, the present invention utilizes an air-floating platform and air-floating support components to achieve zero-gravity unloading of the solar cell array from the ground. Through the optimized layout of the air-floating support components, the problem of complex deployment trajectory of planar deployable solar cell arrays is effectively solved. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the on-orbit configuration of the present invention.

[0031] Figure 2 This is a schematic diagram of the compressed state of the T-shaped spatial solar cell array of the present invention.

[0032] Figure 3 The three-view diagram shows the deployed state of the T-shaped spatial solar cell array of the present invention.

[0033] Figure 4 This is a schematic diagram of the deployment process of the T-shaped spatial solar cell array of the present invention.

[0034] Figures 5(a) and 5(b) are schematic diagrams of the unfolded and retracted states of the offset plate hinge of the present invention, respectively.

[0035] Figure 6 This is a schematic diagram of the inner plate substrate reinforcement.

[0036] Figure 7 This is a schematic diagram of the layout of the air-bearing support components.

[0037] Among them, 1-spacecraft, 2-flight trajectory, 3-pointing towards the Earth's center, 4-connecting frame, 5-inner solar panel, 6-middle solar panel, 7-outer solar panel, 8-sunlight, 9-spacecraft body, 10-pressure release device I, 11-pressure release device II, 12-inter-panel hinge I, 12′-inter-panel hinge II, 13-middle plate linkage device, 14-offset inter-panel hinge I, 14′-offset inter-panel hinge II, 15-connecting frame Linkage device, 16-Connecting frame movement trajectory, 17-Inner plate movement trajectory, 18-Middle plate movement trajectory, 19-Outer plate movement trajectory, 20-Adapter, 21-Locking hinge, 22-Constant torque spring assembly, 23-Hook hinge, 24-Inner plate base plate reinforcement, 25-Connecting frame air bearing support assembly, 26-Inner plate air bearing support assembly, 27-Middle plate air bearing support assembly, 28-Outer plate air bearing support assembly I, 29-Outer plate air bearing support assembly II. Detailed Implementation

[0038] The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0039] This invention provides a planar asymmetric composite unfolded T-shaped space solar array, which is mounted on spacecraft 1, with a flight trajectory 2 and a direction 3 pointing towards the Earth's center as shown in the figure. Figure 1 As shown, when unfolded, it has a T-shaped configuration, which ensures that the solar array has a small moment of inertia around the star and that its center of mass is close to the star, thus ensuring the spacecraft's stable flight in orbit.

[0040] The space solar array includes structural components (connecting frame 4, inner solar panel 5, middle solar panel 6, outer solar panel 7), deployment and locking mechanism, pressing and releasing mechanism, and synchronous linkage device.

[0041] The deployment locking mechanism includes inter-plate hinges and offset inter-plate hinges, both of which are used in pairs. Three pairs of inter-plate hinges are respectively located between the inner solar panel 5 and the middle solar panel 6, between the middle solar panel 6 and the outer solar panel 7, and between the connecting frame 4 and the spacecraft body 9. The offset inter-plate hinge is located between the inner solar panel 5 and the connecting frame 4. The paired inter-plate hinges include inter-plate hinge I12 and inter-plate hinge II12′, and the paired offset inter-plate hinges include offset inter-plate hinge I14 and offset inter-plate hinge II14′.

[0042] The offset plate hinge includes a locking hinge 21, a hook hinge 23, and a constant torque spring assembly 22. The constant torque spring assembly 22 is the drive source. Both the locking hinge 21 and the hook hinge 23 are connected to the adapter 20, causing the connecting frame 4 and the inner solar panel 5 to rotate 90°, as shown in Figures 5(a) and 5(b). The constant torque spring assembly 22 can easily adjust the hinge driving torque by increasing or decreasing the number of constant torque spring plates. After the satellite enters orbit, the constant torque spring assembly 22 drives the solar array to deploy and lock it in place, thereby maintaining the on-orbit deployment stiffness of the solar array.

[0043] The synchronous linkage device includes a connecting frame linkage device 15 and a middle plate linkage device 13; the connecting frame linkage device 15 is installed on the connecting frame 4, so that the connecting frame 4 and the inner solar panel 5 can be deployed synchronously; the middle plate linkage device 13 is installed on the side of the middle solar panel 6, so that the inner solar panel 5, the middle solar panel 6, and the outer solar panel 7 can be deployed synchronously.

[0044] The solar array folds up and presses firmly against the spacecraft body 9, as... Figure 2 As shown, at this time, the solar array is in a folded state. The root connector on the spacecraft body 9 is connected to the connecting frame 4 via inter-plate hinge I 12 and inter-plate hinge II 12'. The connecting frame 4 is connected between the inner solar panel 5 and the middle solar panel 6 via offset inter-plate hinge I 14 and offset inter-plate hinge II 14'. The inner solar panel 5B side and the middle solar panel 6A side are connected via inter-plate hinge I 12 and inter-plate hinge II 12'. The middle solar panel 6B side and the outer solar panel 7A side are connected via inter-plate hinge I 12 and inter-plate hinge II 12'. Figure 3 As shown in the front view, the right side of the inner solar panel 5 is side A, and the left side is side B. The middle solar panel 6 and the outer solar panel 7 are aligned with the A and B sides of the inner solar panel 5.

[0045] The inner five corners of the solar panel must connect to both the inter-panel hinge and the offset inter-panel hinge, such as... Figure 6 As shown. By designing an integrated inner plate substrate reinforcement 24, the corner stiffness and strength of the inner plate 5 substrate of the solar panel are made capable of withstanding large and mismatched bending loads at two load-bearing points.

[0046] When the satellite is launched, such as Figure 2 As shown, the connecting frame 4, the inner solar panel 5, the middle solar panel 6, and the outer solar panel 7 are pressed and fixed to the side wall of the spacecraft body 9 from the inside out by several sets of clamping and releasing devices. The clamping and releasing devices provide a certain preload to ensure that the solar array can withstand the mechanical load of the launch phase. In this embodiment, five sets of clamping and releasing devices are used, including clamping and releasing device I 10 and clamping and releasing device II 11.

[0047] After the spacecraft enters orbit and separates from the launch vehicle, the five sets of clamping and releasing devices release the constraints on the solar array. The solar array then unfolds passively under the passive drive of constant torque spring assemblies 22, consisting of three pairs of inter-panel hinges and one pair of offset inter-panel hinges. The connecting frame 4 and the inner solar panel 5 unfold synchronously under the action of the connecting frame linkage device 15. The inner solar panel 5, the middle solar panel 6, and the outer solar panel 7 unfold synchronously under the action of the middle panel linkage device 13. The connecting frame 4 and the middle solar panel 6 unfold in segments. The unfolding process of the solar array is as follows: Figure 4 As shown in the figure, the motion trajectories of each structural component are displayed, including the motion trajectory 16 of the connecting frame, the motion trajectory 17 of the inner plate, the motion trajectory 18 of the middle plate, and the motion trajectory 19 of the outer plate.

[0048] In the final stage of solar array deployment, the inner solar panel 5, middle solar panel 6, and outer solar panel 7 are deployed and locked in place under the drive of the constant torque spring assembly 22 of the two pairs of inter-panel hinges and the action of the middle panel linkage device 13. The connecting frame 4 and the inner solar panel 5 are deployed and finally locked in place under the drive of the constant torque spring assembly 22 of the pair of inter-panel hinges and the pair of offset inter-panel hinges and the action of the connecting frame linkage device 15. At this time, the inner solar panel 5, middle solar panel 6, and outer solar panel 7 form a plane, which is parallel to the spacecraft body 9 and perpendicular to the connecting frame 4 at 90°. This plane is asymmetrical with respect to the connecting frame 4, thus achieving the so-called "planar asymmetrical composite deployment". After the solar array is deployed, a solar array with a T-shaped connecting frame and three solar panels is formed. The solar array is parallel to the spacecraft body 9 and provides energy to the satellite through sunlight 8.

[0049] In this embodiment, the connecting frame 4 uses two connecting rods, which are respectively connected to both sides of the inner plate of the solar panel 5. The connecting frame 4 and the inner plate of the solar panel 5 form a triangle.

[0050] During ground deployment testing, a gravity compensation device mounted on the solar array supports its weight, simulating a zero-gravity deployment of the solar array. The gravity compensation device includes an air-floating platform and air-floating support components. Compressed air within pipelines ensures that the air-floating support components support the weight of each component of the solar array, such as… Figure 7 As shown. The layout of the air-bearing support assembly follows the principles and requirements of ensuring stable support as much as possible, coinciding with the centroid of the solar panel, and not interfering with each other during the deployment process.

[0051] One air-bearing support assembly is arranged on each of the connecting frame 4, the inner solar panel 5, and the middle solar panel 6, while two air-bearing support assemblies are arranged on the outer solar panel 7. The middle plate air-bearing support assembly 27 on the middle solar panel 6 coincides with the center line of gravity of the middle solar panel 6. The inner plate air-bearing support assembly 26 on the inner solar panel 5 is spaced away from the center line of gravity of the inner solar panel 5 to avoid interference with the air-bearing support assembly on the middle solar panel 6 during the unfolding process. The line connecting the support points of the outer plate air-bearing support assemblies I 28 and II 29 on the outer solar panel 7 coincides with the center line of gravity of the outer solar panel 7. The connecting frame air-bearing support assembly 25 on the connecting frame 4 is located on the end of the connecting frame 4 away from the panel hinge to provide support. The connecting frame 4, the inner solar panel 5, the middle solar panel 6, and the outer solar panel 7 are suspended on the air-bearing platform under the action of the air-bearing support assemblies. Through the cooperation of the various air-bearing support assemblies, the planar unfolding of the solar cell array is realized.

[0052] In summary, the above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A planar asymmetric composite unfolded T-shaped spatial solar cell array, characterized in that, Includes a connecting frame, inner solar panel, middle solar panel, outer solar panel, unfolding and locking mechanism, pressing and releasing mechanism, and synchronous linkage device; The connecting frame is connected between the inner solar panel and the middle solar panel via an unfolding and locking mechanism. The two sides of the middle solar panel are connected to the inner solar panel and the outer solar panel via unfolding and locking mechanisms to form a solar cell array. In the compressed state, the solar cell array is folded up against the side wall of the spacecraft body. From the inside out, the components are the connecting frame, the inner solar panel, the middle solar panel, and the outer solar panel. In the folded state, the compression release mechanism provides a certain compression force to the connecting frame, the inner solar panel, the middle solar panel, and the outer solar panel. After the compression release mechanism releases the constraint on the solar cell array, the inner solar panel, the middle solar panel, and the outer solar panel unfold into a plane under the action of the unfolding and locking mechanism and lock it first. The plane is parallel to the spacecraft body. The connecting frame completes the unfolding and locks it last. The connecting frame is perpendicular to the plane. At this time, the solar cell array has a T-shaped configuration. During ground-based deployment testing, its weight is supported by a gravity compensation device set on the solar array. The gravity compensation device includes an air-floating platform and an air-floating support assembly; One air-bearing support assembly is arranged on each of the connecting frame, the inner plate of the solar panel, and the middle plate of the solar panel, and two air-bearing support assemblies are arranged on the outer plate of the solar panel. The air-bearing support assembly on the middle plate coincides with the centroid line of the solar panel middle plate. The air-bearing support assembly on the inner plate of the solar panel is positioned away from the center line of the inner plate to avoid interference with the air-bearing support assembly on the middle plate of the solar panel during the unfolding process; the line connecting the support points of the two air-bearing support assemblies on the outer plate of the solar panel coincides with the center line of the outer plate of the solar panel; the air-bearing support assembly on the connecting frame is located on the end of the connecting frame away from the hinge of the plate. The connecting frame, inner solar panel, middle solar panel, and outer solar panel are suspended on the air flotation platform under the action of the air flotation support assembly.

2. The planar asymmetric composite unfolded T-shaped space solar cell array as described in claim 1, characterized in that, The deployment locking mechanism includes inter-plate hinges and offset inter-plate hinges, both of which are used in pairs. The inter-plate hinges are respectively installed between the inner solar panel and the middle solar panel, between the middle solar panel and the outer solar panel, and between the connecting frame and the spacecraft body; The offset plate hinge is located between the inner plate and the connecting frame.

3. The planar asymmetric composite unfolded T-shaped space solar cell array as described in claim 1, characterized in that, The connecting frame uses two connecting rods, which are respectively connected to both sides of the inner plate, and the connecting frame and the inner plate form a triangle.

4. The planar asymmetric composite unfolded T-shaped space solar cell array as described in claim 1, characterized in that, The offset plate hinge includes a locking hinge, a hook hinge, and a constant torque spring assembly; The constant torque spring assembly is the driving source, and the locking hinge and hook hinge are both connected to the adapter, so that the connecting frame and the inner plate can rotate 90°.

5. The planar asymmetric composite unfolded T-shaped space solar cell array as described in claim 1, characterized in that, The inner plate corner is provided with an integrated inner plate base reinforcement.

6. The planar asymmetric composite unfolded T-shaped space solar cell array as described in claim 1, characterized in that, The synchronous linkage device includes a connecting frame linkage device and a middle plate linkage device; The connecting frame linkage device is installed on the connecting frame to enable the connecting frame and the inner plate to unfold synchronously; the middle plate linkage device is installed on the side of the middle plate to enable the inner plate, middle plate and outer plate to unfold synchronously.

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