Folding paper derived spacecraft deployable mechanism and unfolding method

Abstract

The invention provides a folding-paper derived spacecraft deployable mechanism and a deploying method. The folding-paper derived spacecraft deployable mechanism comprises a solar wing deploying driving mechanism connected with a spacecraft platform; the solar wing unfolding driving mechanism comprises a solar wing plate assembly, a first driving mechanism and a second driving mechanism, the solar wing plate assembly at least comprises a first solar wing plate and a second solar wing plate, and the first solar wing plate is connected with the spacecraft platform through the first driving mechanism; the same ends of the opposite sides of the first solar wing plate and the second solar wing plate are connected with a first hinge, and the other ends are connected with a second driving mechanism; the first driving mechanism drives the solar wing plate assembly to rotate by 90 degrees to be unfolded, the second driving mechanism drives the second solar wing plate to be unfolded relative to the first solar wing plate, and therefore the other solar wing plates are driven to be unfolded at the same time based on the Miura folding principle. By applying the Mipu origami to the deployable structure of the spacecraft, the constructed deployable structure has the characteristic of single degree of freedom, and the solar wing assembly can be converted from the folded state to the unfolded state only through one drive.

Description

Technical Field

[0001] The present invention belongs to the technical field of aerospace vehicles, and in particular relates to an origami-derived spacecraft deployable mechanism and a deployment method. Background Art

[0002] Due to limitations in rocket carrying capacity and fairing volume, the size of spacecraft solar panels is significantly constrained, requiring them to be pre-folded during launch and deployed upon reaching orbit. Currently, the deployable mechanisms used for deploying spacecraft solar panels are mostly "W"-shaped, exhibiting problems such as a single configuration, a low fold-to-stretch ratio, and severely reduced stiffness over large areas. Origami is a handicraft that can transform large two-dimensional surfaces into small three-dimensional solids. Applying it to spacecraft deployable mechanisms can greatly enrich the folding and unfolding configurations of solar panels while addressing the issue of low fold-to-stretch ratios. Commonly used origami styles include Miura Origami, Flasher Origami, and Yoshimura Origami. However, the application of origami technology to spacecraft deployable structures still presents the following challenges.

[0003] a. The materials used in the deployable mechanisms of origami-derived spacecraft are no longer just soft origami. Considering practical engineering applications, the mechanisms must possess a certain level of strength, rigidity, and quality. b. The deployable mechanism of an origami-derived spacecraft must consider not only the folding process but also issues such as panel thickness and motion interference. c. The deployable mechanism of origami-derived spacecraft requires more actuators and a more complex actuator layout. Summary of the Invention

[0004] In response to the above technical problems, the present invention provides an origami-derived spacecraft deployable mechanism and deployment method. By applying Miura origami to the spacecraft deployable structure, the constructed deployable mechanism has the characteristics of a large folding and unfolding ratio and a simple structure, and has the characteristics of a single degree of freedom. Only one drive is required to complete the transition of the deployable structure from a folded state to an unfolded state.

[0005] To achieve the above object, the technical solution of the present invention is: An origami-derived spacecraft deployable mechanism includes: a solar wing deployment drive mechanism connected to a spacecraft platform; the solar wing deployment drive mechanism includes a solar wing panel assembly, a first drive mechanism, and a second drive mechanism. The solar wing panel assembly includes a plurality of solar wing panels, including at least a first solar wing panel and a second solar wing panel. The first solar wing panel is connected to the spacecraft platform via the first driving mechanism. The side edge of the first solar wing panel and the same end of the side edge of the second solar wing panel are respectively connected to the first hinge, and the other end is connected to the second driving mechanism. The first driving mechanism drives the solar wing panel assembly to rotate 90° relative to the spacecraft platform and unfold, and the second driving mechanism drives the second solar wing panel to unfold relative to the first solar wing panel, thereby driving the remaining solar wing panels in the solar wing panel assembly to unfold simultaneously based on the Miura folding principle.

[0006] Preferably, the solar wing panel assembly further includes a third solar wing panel, a fourth solar wing panel, a fifth solar wing panel and a sixth solar wing panel; When the solar wing panel assembly is in an unfolded state, the solar wing panel assembly comprises solar wing panels arranged in two columns and three rows, wherein the first solar wing panel, the second solar wing panel, and the third solar wing panel are arranged in one column, and the first solar wing panel and the third solar wing panel are respectively located on both sides of the first solar wing panel, and opposite ends of the first solar wing panel and the third solar wing panel are respectively connected by a second hinge; the fourth solar wing panel, the fifth solar wing panel, and the sixth solar wing panel are arranged in one column, the fifth solar wing panel and the first solar wing panel are arranged in one row, the sixth solar wing panel and the fourth solar wing panel are arranged on both sides of the fifth solar wing panel, opposite ends of the fifth solar wing panel and the fourth solar wing panel are respectively connected by a third hinge, and opposite ends of the fifth solar wing panel and the sixth solar wing panel are respectively connected by a fourth hinge; The fourth solar wing panel and the third solar wing panel are arranged in a row, and the opposite sides of the two are connected by a fifth hinge; the sixth solar wing panel and the second solar wing panel are arranged in a row, and the opposite sides of the two are connected by a sixth hinge.

[0007] Preferably, at least two pairs of pressing bases are installed on the side walls of the spacecraft platform, and when the solar wing deployment drive mechanism is in a folded state, the solar wing deployment drive mechanism is folded and relies on the pressing bases.

[0008] Preferably, the second driving mechanism comprises: a first connecting plate, a second connecting plate, a permanent magnet, a first rotating shaft, a magnetic damper, a coil assembly including an iron core, and an outer cylinder frame, wherein the first rotating shaft is fixedly connected to a side edge of the first connecting plate; The permanent magnet, the magnetic damper, and the coil group including the iron core are arranged in the outer cylinder frame. The outer cylinder frame, the coil group including the iron core, the magnetic damper, and the second connecting plate are all connected to the first rotating shaft through bearings. The permanent magnet is sleeved on the first rotating shaft. The first connecting plate is fixedly connected to the second solar wing panel, and the second connecting plate is connected to the first solar wing panel.

[0009] Preferably, the first driving mechanism includes a seventh hinge and a connecting frame, one end of the connecting frame is fixedly connected to the first solar wing panel, and the other end is fixedly connected to the seventh hinge, and the seventh hinge is fixedly connected to the spacecraft platform.

[0010] Preferably, the seventh hinge includes a first male hinge, a first female hinge, a second rotating shaft, a power wheel, and a drive motor, wherein the first female hinge is fixedly connected to the second rotating shaft, the first male hinge is connected to the second rotating shaft via a bearing, the first female hinge is fixedly connected to the connecting frame, the first male hinge is fixedly connected to the spacecraft platform, and the power wheel is sleeved on one end of the second rotating shaft; The drive motor is connected to the power wheel, and the drive motor drives the power wheel, driving the first female hinge to rotate, and then driving the rotating frame and the solar wing panel assembly to rotate relative to the spacecraft platform.

[0011] Preferably, the first hinge, the second hinge, the third hinge, the fourth hinge, the fifth hinge and the sixth hinge have the same structure and all include a second male hinge, a second female hinge and a third rotating shaft. The second female hinge is fixedly connected to the third rotating shaft, and the second male hinge is connected to the third rotating shaft through a bearing.

[0012] Preferably, limit blocks are provided on both the first male hinge and the second male hinge, and the limit blocks are used to limit the solar wing panels from unfolding to a preset angle.

[0013] Preferably, the first hinge, the second hinge, the third hinge, the fourth hinge, the fifth hinge, the sixth hinge and the seventh hinge are all provided with a locking mechanism, and the locking mechanism is used to lock the deployment angle of the solar wing panel.

[0014] Based on the same inventive concept, the present invention also provides a method for deploying an origami-derived spacecraft deployable mechanism, comprising the following steps: S1: The first driving mechanism drives the solar wing panel assembly to rotate 90° relative to the spacecraft platform and lock it; S2: The second driving mechanism drives the second solar wing panel to unfold relative to the first solar wing panel, thereby driving the remaining solar wing panels to unfold synchronously into a working plane.

[0015] Due to the adoption of the above technical solution, the present invention has the following advantages and positive effects compared with the prior art: The present invention provides an origami-derived spacecraft deployable mechanism, wherein the solar wing assembly comprises multiple solar wing panels. These panels are rigid panels connected by hinges. Based on the Miura origami principle, a second drive mechanism only needs to drive the first and second solar wing panels to deploy in order to drive the deployment of the remaining panels. Therefore, the present invention's solar wing assembly, based on the Miura origami principle, utilizes rigid panels and hinges to achieve the single-degree-of-freedom motion of Miura origami, resulting in a large fold-to-expand ratio and possessing sufficient strength, rigidity, and quality. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 A schematic diagram of an origami-derived spacecraft deployable mechanism in a deployed state according to an embodiment of the present invention; Figure 2 Schematic diagram of the folded state of the deployable mechanism of the origami-derived spacecraft according to an embodiment of the present invention Figure 3 is a schematic diagram of a second driving mechanism in a deployable mechanism of an origami-derived spacecraft according to an embodiment of the present invention; Figure 4 Schematic diagram of the structure of the seventh hinge in the deployable mechanism of an origami-derived spacecraft according to an embodiment of the present invention; Figure 5 Schematic diagram of the structure of the first to sixth hinges in the deployable mechanism of an origami-derived spacecraft according to an embodiment of the present invention.

[0017] Explanation of reference numerals: 1-spacecraft platform; 2-pressing base; 3-solar wing panel assembly, 301-first solar wing panel; 302-second solar wing panel; 303-third solar wing panel; 304-fourth solar wing panel; 305-fifth solar wing panel; 306-sixth solar wing panel; 4-connecting frame; 5-second driving mechanism; 501-first connecting plate; 502-second connecting plate; 503-permanent magnet; 504-first rotating shaft; 505-magnetic resistance Body; 506 - coil assembly including an iron core; 507 - outer cylinder frame; 6 - first hinge; 601 - second male hinge; 602 - second female hinge; 603 - third rotating shaft; 7 - second hinge; 8 - third hinge; 9 - fourth hinge; 10 - fifth hinge; 11 - sixth hinge; 12 - seventh hinge; 1201 - first male hinge; 1202 - first female hinge; 1203 - second rotating shaft; 1204 - power wheel; 1205 - drive motor; 13-limiting block; 14-locking block; 15-leaf spring hook. DETAILED DESCRIPTION

[0018] The following is a detailed description of an origami-derived spacecraft deployable mechanism proposed by the present invention in conjunction with the accompanying drawings and specific embodiments. The advantages and features of the present invention will become more apparent from the following description.

[0019] See Figure 1-2 A paper-derived spacecraft deployable mechanism includes: a solar wing deployment drive mechanism connected to a spacecraft platform 1; the solar wing deployment drive mechanism includes a solar wing panel assembly 3, a first drive mechanism and a second drive mechanism 5, The solar wing panel assembly 3 includes a plurality of rigid solar wing panels, which are hingedly connected to each other, wherein the solar wing panel assembly 3 includes at least a first solar wing panel 301 and a second solar wing panel 302, the first solar wing panel 301 being connected to the spacecraft platform 1 via a first driving mechanism, the first solar wing panel 301 and the second solar wing panel 302 having opposite sides thereof having the same end connected to the first hinge 6, and the other end connected to the second driving mechanism 5. Preferably, the first solar wing panel 301 and the second solar wing panel 302 are arranged vertically, and the opposite sides of the first solar wing panel 301 and the second solar wing panel 302 have opposite sides thereof having the end away from the spacecraft platform 1 connected to the first hinge 6, and the end close to the spacecraft platform 1 connected to the second driving mechanism 5; The first driving mechanism drives the solar wing panel assembly 3 to rotate 90° relative to the spacecraft platform 1 and unfold, and the second driving mechanism 5 drives the second solar wing panel 302 to unfold relative to the first solar wing panel 301, thereby driving the remaining solar wing panels in the solar wing panel assembly 3 to unfold simultaneously based on the Miura folding principle.

[0020] This embodiment provides an origami-derived spacecraft deployable mechanism, wherein the solar wing assembly 3 comprises multiple solar wing panels. These panels are rigid panels connected by hinges. Based on the Miura origami principle, only the second drive mechanism 5 needs to deploy the first and second solar wing panels 301, 302 to drive the deployment of the remaining panels. Therefore, the solar wing assembly 3 based on the Miura origami principle utilizes rigid panels and hinges to achieve the single-degree-of-freedom motion of Miura origami, resulting in a large fold-to-expand ratio and possessing sufficient strength, rigidity, and quality.

[0021] Preferably, the solar wing panel assembly 3 further includes a third solar wing panel 303, a fourth solar wing panel 304, a fifth solar wing panel 305 and a sixth solar wing panel 306. The first solar wing panel 301, the second solar wing panel 302, the third solar wing panel 303, the fourth solar wing panel 304, the fifth solar wing panel 305 and the sixth solar wing panel 306 have the same thickness, preferably 9-11 mm, such as 9, 10 and 11 mm. The first solar wing panel 301 and the fifth solar wing panel 305 are parallelograms of the same size and shape, and the second solar wing panel 302, the third solar wing panel 303, the fourth solar wing panel 304 and the sixth solar wing panel are right-angled trapezoids of the same size and shape. When the solar wing panel assembly 3 is in the unfolded state, the solar wing panel assembly 3 is composed of two columns and three rows of solar wing panels. The first solar wing panel 301, the second solar wing panel 302, and the third solar wing panel 303 are distributed in one column, close to the side of the spacecraft platform 1; the fourth solar wing panel 304, the fifth solar wing panel 305 and the sixth solar wing panel 306 are distributed in one column, away from the side of the spacecraft platform 1.

[0022] The second solar wing panel 302 and the third solar wing panel 303 are respectively located on both sides of the first solar wing panel 301, and the opposite sides of the first solar wing panel 301 and the third solar wing panel 303 are respectively connected by a second hinge 7; the fifth solar wing panel 305 is arranged in a row with the first solar wing panel 301, and the sixth solar wing panel 306 and the fourth solar wing panel 304 are arranged on both sides of the fifth solar wing panel 305, and the opposite sides of the fifth solar wing panel 305 and the fourth solar wing panel 304 are respectively connected by a third hinge 8, and the opposite sides of the fifth solar wing panel 305 and the sixth solar wing panel 306 are respectively connected by a fourth hinge 9; The fourth solar wing panel 304 and the third solar wing panel 303 are arranged in a row, and the opposite side ends of the two are away from the setting position of the second hinge 7 or the third hinge 8 are connected by the fifth hinge 10; the sixth solar wing panel 306 and the second solar wing panel 302 are arranged in a row, and the opposite side ends of the two are away from the setting position of the first hinge 6 or the fourth hinge 9 are connected by the sixth hinge 11.

[0023] In the solar wing panel assembly 3, the third solar wing panel 303, the first solar wing panel 301 and the second solar wing panel 302 in a row close to the spacecraft platform 1 and the fourth solar wing panel 304, the fifth solar wing panel 305 and the sixth solar wing panel 306 in a row away from the spacecraft platform 1 are mirror-symmetrical, wherein the third solar wing panel 303 and the second solar wing panel 302 are right-angled trapezoids, and the first solar wing panel 301 is a parallelogram. The smaller vertex angle C of the third solar wing panel 303 of the right-angled trapezoid is equal to the smaller vertex angle B of the first solar wing panel 301 of the parallelogram, and is also equal to the smaller vertex angle A of the second solar wing panel 302 of the right-angled trapezoid, and the vertex angle is less than 90°.

[0024] In this embodiment, by optimizing the shape, thickness, and hinge design of the solar wing panels, interference issues that may occur during the deployment or folding of the drive mechanism are avoided. Furthermore, a single drive mechanism is used to drive the deployment of the single-degree-of-freedom mechanism, reducing the complexity and weight of the drive system.

[0025] Preferably, at least two pairs of pressing bases 2 are installed on the side walls of the spacecraft platform 1 , and when the solar wing deployment drive mechanism is in a folded state, the solar wing deployment drive mechanism is folded and leans against the pressing bases 2 .

[0026] Preferably, if Figure 3 As shown, the second driving mechanism 5 includes: a first connecting plate 501, a second connecting plate 502, a permanent magnet 503, a first rotating shaft 504, a magnetic damper 505, a coil assembly 506 including an iron core, and an outer cylinder frame 507. The first rotating shaft 504 is fixedly connected to the side of the first connecting plate 501; The permanent magnet 503, the magnetic damper 505, and the coil assembly 506 including the iron core are disposed within the outer cylinder frame 507. The second connecting plate 502, the magnetic damper 505, and the coil assembly 506 including the iron core are fixedly connected to the outer cylinder frame 507. The outer cylinder frame 507, the coil assembly 506 including the iron core, the magnetic damper 505, and the second connecting plate 502 are all connected to the first rotating shaft 504 via bearings. The permanent magnet 503 is sleeved on the first rotating shaft 504. The first connecting plate 501 is fixedly connected to the second solar wing panel 302 , and the second connecting plate 502 is connected to the first solar wing panel 301 .

[0027] When the coil group 506 containing the iron core is energized, it generates a magnetic field, driving the permanent magnet 503 to rotate, and then driving the rotation axis and the first connecting plate 501 to rotate, thereby driving the second solar wing panel 302 to unfold relative to the first solar wing panel 301. After the second solar wing panel 302 and the first solar wing panel 301 form a plane, the coil containing the iron core is de-energized, and the magnetic damper 505 attracts the permanent magnet 503 to lock the unfolding angle of the second solar wing panel 302 and the first solar wing panel 301.

[0028] The first driving mechanism includes a seventh hinge 12 and a connecting frame 4 . One end of the connecting frame 4 is fixedly connected to the first solar wing panel 301 , and the other end is fixedly connected to the seventh hinge 12 . The seventh hinge 12 is fixedly connected to the spacecraft platform 1 .

[0029] like Figure 4 As shown, the seventh hinge 12 includes: a first male hinge 1201, a first female hinge 1202, a second rotating shaft 1203, a power wheel 1204, and a drive motor 1205. The first female hinge 1202 is fixedly connected to the second rotating shaft 1203. The first male hinge 1201 is connected to the second rotating shaft 1203 through a bearing. The first female hinge 1202 is fixedly connected to the connecting frame 4. The connecting frame 4 is fixedly connected to the first solar wing panel 301. The first male hinge 1201 is fixedly connected to the spacecraft platform 1. The power wheel 1204 is mounted on one end of the second rotating shaft 1203. The drive motor 1205 is connected to the power wheel 1204. The drive motor 1205 drives the power wheel 1204, driving the first female hinge 1202 to rotate, and then driving the connecting frame 4 and the solar wing panel assembly 3 to rotate relative to the spacecraft platform 1.

[0030] A limit block 13 is provided on the first male hinge 1201 , and the limit block 13 is used to limit the rotation angle of the first female hinge 1202 .

[0031] A locking mechanism is provided on the seventh hinge, which includes a leaf spring hook 15 and a locking block 14. The locking block 14 is provided on the first male hinge 1201, and the leaf spring hook 15 is provided on the first female hinge 1202. After the first female hinge 1202 is unfolded to a preset angle, the leaf spring hook 15 hooks the locking block 14 to lock the hinge unfolding angle.

[0032] Preferably, the first hinge 6, the second hinge 7, the third hinge 8, the fourth hinge 9, the fifth hinge 10, and the sixth hinge 11 have the same structure. Figure 5 As shown, they all include a second male hinge 601, a second female hinge 602, and a third rotating shaft 603. The second female hinge 602 is fixedly connected to the third rotating shaft 603, and the second male hinge 601 is connected to the third rotating shaft 603 through a bearing.

[0033] A limit block 13 is provided on the second male hinge 601 , and the limit block 13 is used to limit the solar wing panel to be unfolded to a preset angle. Preferably, the unfolded state of the solar wing panel assembly 3 is a working plane.

[0034] Preferably, the first hinge 6, second hinge 7, third hinge 8, fourth hinge 9, fifth hinge 10, sixth hinge 11, and seventh hinge 12 are each equipped with a locking mechanism for locking the solar wing panel's deployment angle. The locking mechanism includes a leaf spring hook 15 and a locking block 14. The locking block 14 is provided on the second male hinge 601, and the leaf spring hook 15 is provided on the second female hinge 602. When the second female hinge 602 is deployed to a preset angle, the leaf spring hook 15 engages the locking block 14, locking the hinge's deployment angle.

[0035] A method for deploying an origami-derived spacecraft deployable mechanism comprises the following steps: S1: The first driving mechanism drives the solar wing panel assembly 3 to rotate 90° relative to the spacecraft platform 1 and lock it; The drive motor 1205 drives the power wheel 1204, driving the first female hinge 1202 to rotate, and then driving the connecting frame 4 and the solar wing panel assembly 3 to rotate relative to the spacecraft platform 1. The limit block 13 limits the rotation of the solar wing panel assembly 3 by 90 degrees. After the solar wing panel rotates into place, the leaf spring hook 15 hooks the locking block 14 to fix the solar wing panel assembly 3 in this position.

[0036] S2: The second driving mechanism 5 drives the second solar wing panel 302 to unfold relative to the first solar wing panel 301: When the coil assembly 506 including the iron core is energized, it generates a magnetic field, driving the permanent magnet 503 to rotate, thereby driving the rotation axis and the first connecting plate 501 to rotate, thereby driving the second solar wing panel 302 to unfold relative to the first solar wing panel 301. After the second solar wing panel 302 and the first solar wing panel 301 form a plane, the coil assembly including the iron core is de-energized, and the magnetic damper 505 attracts the permanent magnet 503 to lock the unfolding angle of the second solar wing panel 302 and the first solar wing panel 301. When the first solar wing panel 301 and the second solar wing panel 302 are unfolded, based on the Miura origami principle, the remaining solar wing panels are unfolded synchronously to form a working plane.

[0037] During the unfolding process, the limit blocks 13 on the first hinge 6, the second hinge 7, the third hinge 8, the fourth hinge 9, the fifth hinge 10 and the sixth hinge 11 limit the unfolding angle to 180 degrees. After unfolding to the full extent, the leaf spring hook 15 hooks the locking block 14, thereby keeping the unfolding angle of the solar wing panel at 180 degrees.

[0038] The embodiments of the present invention have been described in detail above with reference to the accompanying drawings, but the present invention is not limited to the above embodiments. Even if various changes are made to the present invention, if these changes fall within the scope of the claims of the present invention and their equivalents, they still fall within the scope of protection of the present invention.

Claims

1. An origami-derived spacecraft deployable mechanism, characterized in that: include: a solar wing deployment drive mechanism connected to a spacecraft platform; The solar wing deployment drive mechanism includes a solar wing panel assembly, a first drive mechanism and a second drive mechanism. The solar wing panel assembly includes a plurality of solar wing panels, including at least a first solar wing panel and a second solar wing panel. The first solar wing panel is connected to the spacecraft platform via the first driving mechanism. The first solar wing panel and the second solar wing panel have opposite sides with the same end connected to the first hinge and the other end connected to the second driving mechanism. The first driving mechanism drives the solar wing panel assembly to rotate 90° relative to the spacecraft platform and unfold, and the second driving mechanism drives the second solar wing panel to unfold relative to the first solar wing panel, thereby driving the remaining solar wing panels in the solar wing panel assembly to unfold simultaneously based on the Miura folding principle.

2. The origami-derived spacecraft deployable mechanism according to claim 1, characterized in that: The solar wing panel assembly further includes a third solar wing panel, a fourth solar wing panel, a fifth solar wing panel and a sixth solar wing panel; When the solar wing panel assembly is in an unfolded state, the solar wing panel assembly comprises solar wing panels arranged in two columns and three rows, wherein the first solar wing panel, the second solar wing panel, and the third solar wing panel are arranged in one column, and the first solar wing panel and the third solar wing panel are respectively located on both sides of the first solar wing panel, and opposite ends of the second solar wing panel and the third solar wing panel are respectively connected by a second hinge; the fourth solar wing panel, the fifth solar wing panel, and the sixth solar wing panel are arranged in one column, the fifth solar wing panel and the first solar wing panel are arranged in one row, the sixth solar wing panel and the fourth solar wing panel are arranged on both sides of the fifth solar wing panel, opposite ends of the fifth solar wing panel and the fourth solar wing panel are respectively connected by a third hinge, and opposite ends of the fifth solar wing panel and the sixth solar wing panel are respectively connected by a fourth hinge; The fourth solar wing panel and the third solar wing panel are arranged in a row, and the opposite sides of the two are connected by a fifth hinge; the sixth solar wing panel and the second solar wing panel are arranged in a row, and the opposite sides of the two are connected by a sixth hinge.

3. The origami-derived spacecraft deployable mechanism according to claim 1, characterized in that: At least two pairs of pressing bases are installed on the side walls of the spacecraft platform. When the solar wing deployment drive mechanism is in a folded state, the solar wing deployment drive mechanism is folded and leans against the pressing bases.

4. The origami-derived spacecraft deployable mechanism according to claim 1, characterized in that: The second driving mechanism includes: a first connecting plate, a second connecting plate, a permanent magnet, a first rotating shaft, a magnetic damper, a coil assembly including an iron core, and an outer cylinder frame, wherein the first rotating shaft is fixedly connected to a side edge of the first connecting plate; The permanent magnet, the magnetic damper, and the coil group including the iron core are arranged in the outer cylinder frame. The coil group including the iron core, the magnetic damper, and the second connecting plate are fixedly connected to the outer cylinder frame. The outer cylinder frame is connected to the first rotating shaft through a bearing. The permanent magnet is sleeved on the first rotating shaft. The first connecting plate is fixedly connected to the second solar wing panel, and the second connecting plate is connected to the first solar wing panel.

5. The origami-derived spacecraft deployable mechanism according to claim 2, characterized in that: The first driving mechanism includes a seventh hinge and a connecting frame, one end of the connecting frame is fixedly connected to the first solar wing panel, and the other end is fixedly connected to the seventh hinge, and the seventh hinge is fixedly connected to the spacecraft platform.

6. The origami-derived spacecraft deployable mechanism according to claim 5, characterized in that: The seventh hinge includes a first male hinge, a first female hinge, a second rotating shaft, a power wheel, and a drive motor, wherein the first female hinge is fixedly connected to the second rotating shaft, the first male hinge is connected to the second rotating shaft via a bearing, the first female hinge is fixedly connected to the connecting frame, the first male hinge is fixedly connected to the spacecraft platform, and the power wheel is sleeved on one end of the second rotating shaft; The drive motor is connected to the power wheel, and the drive motor drives the power wheel, driving the first female hinge to rotate, and then driving the rotating frame and the solar wing panel assembly to rotate relative to the spacecraft platform.

7. The origami-derived spacecraft deployable mechanism according to claim 2, characterized in that: The first hinge, second hinge, third hinge, fourth hinge, fifth hinge and sixth hinge have the same structure and all include a second male hinge, a second female hinge and a third rotating shaft. The second female hinge is fixedly connected to the third rotating shaft, and the second male hinge is connected to the third rotating shaft through a bearing.

8. The origami-derived spacecraft deployable mechanism according to claim 6 or 7, characterized in that: The first male hinge and the second male hinge are both provided with limit blocks, and the limit blocks are used to limit the solar wing panels from unfolding to a preset angle.

9. The origami-derived spacecraft deployable mechanism according to claim 6 or 7, characterized in that: The first hinge, the second hinge, the third hinge, the fourth hinge, the fifth hinge, the sixth hinge and the seventh hinge are all provided with a locking mechanism, and the locking mechanism is used to lock the deployment angle of the solar wing panel.

10. A method for deploying an origami-derived spacecraft deployable mechanism according to any one of claims 1 to 9, characterized in that: The following steps are involved: S1: The first driving mechanism drives the solar wing panel assembly to rotate 90° relative to the spacecraft platform and lock it; S2: The second driving mechanism drives the second solar wing panel to unfold relative to the first solar wing panel, thereby driving the remaining solar wing panels to unfold synchronously into a working plane.

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