Passive deployment high stiffness flexible solar wing and its deployment method

Abstract

The application provides a passive deployment high-rigidity flexible solar wing and a deployment method thereof. The solar wing comprises a lifting mechanism, a swing support, two frame-type solar wings and a reinforcing beam. The lifting mechanism is used for driving the deployment passively and lifting the frame-type solar wings to a certain distance from a spacecraft body when the constraint is released. The swing support is arranged on the top of the lifting mechanism, bears the frame-type solar wings and can rotate relatively. The two frame-type solar wings are symmetrically arranged on the two sides of the swing support and comprise foldable frame structures and flexible solar wing plates. The reinforcing beam is connected between the two frame-type solar wings and is used for reinforcing the connection between the connecting frames at the roots of the two frame-type solar wings, further improving the overall rigidity and playing a role of compacting the lifting mechanism in the folded state. The application simplifies the system structure, significantly improves the deployment rigidity of the solar wing and can effectively resist the in-orbit vibration load.

Description

Technical Field

[0001] This invention relates to the field of aerospace equipment technology, specifically to a passive deployable high-rigidity flexible solar array and its deployment method. Background Technology

[0002] As spacecraft power demands increase, the need for larger solar arrays is growing, leading to a decrease in deployment stiffness and posing new challenges to solar arrays. In recent years, with the rapid development of satellite constellations such as the Internet constellation, constellations like the State Grid have proposed a bamboo dragonfly-shaped solar array, which uses a lifting mechanism to raise the solar array above the satellite and deploy it, thus achieving a large-area solar array configuration.

[0003] However, existing bamboo-dragonfly-shaped solar arrays have the following drawbacks: First, insufficient stiffness, making them prone to vibration during on-orbit operation after deployment; second, complex system, as the deployment mechanism typically requires an active motor to drive the deployment, increasing system complexity and cost due to the motor and its control circuitry; third, existing lifting mechanisms often employ folding rod configurations, resulting in poor joint stiffness and, under the same envelope conditions, a small moment of inertia, affecting the overall stiffness of the solar array.

[0004] Therefore, there is an urgent need for a passive deployable flexible solar array solution that can achieve design goals such as high deployment stiffness, high reliability, and low cost while reducing system complexity. Summary of the Invention

[0005] To address the shortcomings of existing technologies, the purpose of this invention is to provide a passive deployable high-rigidity flexible solar array and its deployment method.

[0006] A passive deployable high-stiffness flexible solar array provided by the present invention includes: The lifting mechanism is used to passively deploy and lift the frame-type solar array to a set distance from the spacecraft body when the restraints are released; A swing bracket, located at the top of the lifting mechanism, is used to support the frame-type solar panel and is capable of rotating relative to the lifting mechanism; Two frame-type solar panels are symmetrically arranged on both sides of the swing bracket. Each frame-type solar panel includes a foldable frame structure and a flexible solar panel plate laid on the frame structure. A reinforcing beam is connected between the two frame-type solar panels. The reinforcing beam is used to strengthen the connection between the root connecting frames of the two frame-type solar panels and to press the lifting mechanism in the retracted state.

[0007] Preferably, the lifting mechanism includes a base, a turntable, a scissor lift unit, a top seat, and a drive spring; The base is used for fixed installation on the spacecraft body; The turntable is rotatably connected to the base; the scissor lift unit is connected between the turntable and the top seat; the drive spring is used to store elastic potential energy in the retracted state and drive the scissor lift unit to unfold when the constraint is released, so as to lift the top seat. The swing bracket is positioned above the top seat.

[0008] Preferably, it further includes: a clamping and releasing mechanism, used to clamp and fix each component of the passive deployable high-rigidity flexible solar array in the retracted state, and to release the clamping constraint when deployed.

[0009] Preferably, it further includes: a constraint release mechanism, used to constrain the deployment degree of freedom of the passive deployable high-rigidity flexible solar array in the retracted state, and to release the deployment constraint when deployed.

[0010] Preferably, the frame structure includes a longitudinal bar assembly, a foldable triangular frame, and a first folding bar; The foldable triangular frames are distributed on both sides of the vertical rod assembly and connected by the vertical rod assembly, with the foldable triangular frames on each side arranged in series. The first folding rod is connected between two adjacent foldable triangular frames and is used to constrain the relative position between the two adjacent foldable triangular frames when unfolded.

[0011] Preferably, the clamping release mechanism includes a clamping beam, a clamping rope, and a hot knife assembly; The clamping beam is located on the retracted outer side of the passive deployable high-rigidity flexible solar array. The clamping rope passes through the clamping beam to clamp and fix each component of the passive deployable high-rigidity flexible solar array to the spacecraft body. The hot knife assembly is used to cut the clamping rope during deployment to release the clamping constraint. Preferably, the restraint release mechanism includes a support and a restraint rope; The support is fixed to the lowest foldable triangular frame in the folded state. One end of the constraint rope is connected to the support, and the other end is connected to the uppermost foldable triangular frame in the folded state. In the folded state, the constraint rope is in a taut state, constraining the degree of freedom of the passive deployable high-rigidity flexible solar array. When deployed, the constraint rope is cut or released to release the deployment constraint on the passive deployable high-rigidity flexible solar array.

[0012] A method for deploying a passive, high-stiffness, flexible solar array according to the present invention includes: Step S1: After the satellite enters orbit, the clamping and releasing mechanism is unlocked and released in batches. The lifting mechanism and the second folding rod are deployed and locked under the action of their own torsion spring torque, raising the solar array to a predetermined height above the spacecraft body. Step S2: The constraint release mechanism is powered on to release the constraint on the frame unit of the frame-type solar panel. The frame unit unfolds into place and locks under its own spring torque. Step S3: Under the action of the first and second drives of the lifting mechanism, the two solar arrays achieve solar orientation tracking.

[0013] Compared with the prior art, the present invention has the following beneficial effects: 1. This application adopts a frame-type solar array structure, which forms triangular cubic units by arranging foldable triangular frames in series along the longitudinal rod assembly. When unfolded, it forms a high-rigidity frame structure. With the help of reinforcing beams, the solar array has excellent structural rigidity in the unfolded state and can effectively resist on-orbit vibration loads.

[0014] 2. Compared with the folding rod, the scissor lift mechanism described in this application has a larger moment of inertia and better stiffness under the same envelope conditions.

[0015] 3. The lifting mechanism and solar array described in this invention are both passive torsion spring driven deployment, which reduces system complexity and improves deployment reliability. Attached Figure Description

[0016] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a planar schematic diagram of the passive deployable high-rigidity flexible solar array in the retracted state according to an embodiment of this application; Figure 2 This is a planar schematic diagram of the passive deployment high-rigidity flexible solar array after deployment during the lifting and flipping stage, according to an embodiment of this application. Figure 3 This is a planar schematic diagram of the passive deployable high-rigidity flexible solar array in the deployed state according to an embodiment of this application; Figure 4 This is a plan view of the lifting mechanism of the passive deployable high-rigidity flexible solar array according to an embodiment of this application; Figure 5 This is a plan view of the frame-type solar wing portion of the passive deployable high-rigidity flexible solar wing according to an embodiment of this application. Figure 6 This is a schematic diagram of the foldable triangular frame of the passive deployable high-rigidity flexible solar array according to an embodiment of this application.

[0017] Explanation of reference numerals in the attached figures: 1. Lifting mechanism; 1-1. Base; 1-2. Turntable; 1-3. Short rod; 1-4. Inner scissor bar; 1-5. Outer scissor bar; 1-6. Top seat; 1-7. Swing bracket; 1-8. First drive; 1-9. Second drive; 1-10. Drive spring; 1-11. Locking assembly; 2. Frame-type solar panel; 2-1. Longitudinal rod assembly; 2-2. Foldable triangular frame; 2-2-1. Composite material rectangular tube; 2-2-2. First folding plate; 2-2-3. Hinge shaft; 2-2-4. Torsion spring; 2-2-5. Locking hook; 2-2-6. Lock head; 2-3. First folding rod; 2-4. Connecting frame; 2-5. Second folding rod; 2-6. Flexible solar panel; 3. Reinforcing beam; 4. Compression release mechanism; 5. Restraint release mechanism. Detailed Implementation

[0018] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0019] Example 1 like Figures 1 to 6 As shown, this embodiment provides a passively deployable high-rigidity flexible solar array. The passively deployable high-rigidity flexible solar array includes a lifting mechanism 1, two frame-type solar arrays 2, a reinforcing beam 3, a clamping release mechanism 4, and a constraint release mechanism 5. The lifting mechanism 1 is used to passively drive the deployment and lift the frame-type solar arrays 2 to a certain distance from the spacecraft body when the constraint is released. The two frame-type solar arrays 2 are symmetrically arranged on both sides of the swing brackets 1-7. The reinforcing beam 3 is a rigid connector, connecting the root connecting frames 2-4 of the two frame-type solar arrays 2, strengthening the connection between the root connecting frames 2-4 of the two frame-type solar arrays 2, further improving the overall rigidity, and simultaneously clamping the lifting mechanism 1 in the retracted state. The clamping release mechanism 4 and the constraint release mechanism 5 are used to achieve clamping fixation and deployment constraint in the retracted state.

[0020] like Figure 4 As shown, the lifting mechanism 1 includes a base 1-1, a turntable 1-2, a short rod 1-3, an inner scissor bar 1-4, an outer scissor bar 1-5, a top seat 1-6, a swing bracket 1-7, a first drive 1-8, a second drive 1-9, a drive spring 1-10, and a locking assembly 1-11. The base 1-1 is fixedly mounted on the spacecraft body. The turntable 1-2 is rotatably connected to the base 1-1, allowing the turntable 1-2 to rotate relative to the base 1-1. The inner scissor bar 1-4 and the outer scissor bar 1-5 are hinged at their midpoints to form a scissor unit. The short rod 1-3 is connected at the end of the scissor unit between the turntable 1-2 or the top seat 1-6, constituting a scissor-type lifting structure. The top seat 1-6 is located above the scissor unit, and the swing bracket 1-7 is positioned above the top seat 1-6, used to support the frame-type solar array 2 and capable of rotating relative to the top seat 1-6.

[0021] The drive spring 1-10 is a torsion spring, sleeved on the hinge axis of the inner scissor lift 1-4 and the outer scissor lift 1-5. In the retracted state, the drive spring 1-10 stores elastic potential energy; when the constraint is released, the drive spring 1-10 releases the elastic potential energy, driving the inner scissor lift 1-4 and the outer scissor lift 1-5 to rotate around their hinge point, causing the scissor lift unit to unfold and raising the top seat 1-6 to a certain distance from the spacecraft body. The locking assembly 1-11 includes a locking pin and a locking hole. The locking pin engages with the locking hole under the action of elastic force, and is inserted into the locking hole when the scissor lift unit is unfolded into position, realizing the rigid locking of the lifting mechanism 1 to establish the unfolding stiffness of the lifting mechanism 1.

[0022] The first drive 1-8 drives the turntable 1-2 to rotate relative to the base 1-1, thereby achieving rotation in the azimuth direction; the second drive 1-9 drives the swing bracket 1-7 to rotate relative to the top seat 1-6, thereby achieving rotation in the pitch direction. Through the cooperation of the first drive 1-8 and the second drive 1-9, two-dimensional solar orientation of the two frame-type solar arrays 2 can be achieved, simplifying the system complexity.

[0023] like Figure 5 As shown, each frame-type solar panel 2 includes a longitudinal rod assembly 2-1, a foldable triangular frame 2-2, a first folding rod 2-3, a connecting frame 2-4, a second folding rod 2-5, and a flexible solar panel 2-6. The foldable triangular frames 2-2 are distributed on both sides of the longitudinal rod assembly 2-1 and connected by the longitudinal rod assembly 2-1. The foldable triangular frames 2-2 on each side are arranged in series. Each foldable triangular frame 2-2 has a folded state and an unfolded state. Adjacent foldable triangular frames 2-2 are connected by the first folding rod 2-3. The connecting frame 2-4 is located at the end of the longitudinal rod assembly 2-1 and is used to connect to the swing bracket 1-7. The second folding rod 2-5 connects the connecting frame 2-4 and the longitudinal rod assembly 2-1. The flexible solar panel 2-6 is laid on the frame surface formed by the unfolded longitudinal rod assembly 2-1 and the foldable triangular frames 2-2.

[0024] like Figure 6As shown, the foldable triangular frame 2-2 consists of two equal-length composite rectangular tubes 2-2-1 and one foldable metal plate assembly. The two ends of the composite rectangular tubes 2-2-1 are equipped with metal joints. The foldable metal plate assembly includes two first folding plates 2-2-2, which are hinged together by a hinge shaft 2-2-3. A torsion spring 2-2-4 is installed on the hinge shaft 2-2-3, allowing the two first folding plates 2-2-2 to unfold 180 degrees from a folded state to an unfolded state. A locking hook 2-2-5 is installed on one side of the folding plate of the hinge shaft 2-2-3, and a lock head 2-2-6 is provided on the other side. When the unfolded position is reached, the locking hook 2-2-5 engages with the lock head 2-2-6 to lock the two first folding plates 2-2-2. The other two ends of the two first folding plates 2-2-2 are respectively rotatably connected to the two composite rectangular tubes 2-2-1. In the unfolded state, the two equal-length composite rectangular tubes 2-2-1 and one foldable metal plate assembly form a stable triangular structure.

[0025] The first folding rod 2-3 includes two first sub-folding rods, which are hinged together by a hinge shaft. A torsion spring is mounted on this hinge shaft, allowing the two first sub-folding rods to unfold 180 degrees from a folded state to an unfolded state. A locking hook is provided on one side of the folding plate of this hinge shaft, and a lock head is provided on the other side. When the unfolded position is reached, the locking hook engages with the lock head to lock the two first sub-folding rods. The hinge shaft of the first folding rod 2-3 is spatially orthogonal to the hinge shaft 2-2-3 of the foldable triangular frame 2-2; that is, the axial direction of the hinge shaft of the first folding rod 2-3 is perpendicular to the longitudinal rod assembly 2-1, and the axial direction of the hinge shaft 2-2-3 of the foldable triangular frame 2-2 is parallel to the longitudinal rod assembly 2-1. In the folded state, the first folding rod 2-3 folds and retracts to compress the envelope size of the frame-type solar panel 2 in the length direction.

[0026] The reinforcing beam 3 is a rigid connector, connecting the root connecting frames 2-4 of the two frame-type solar panels 2. The reinforcing beam 3 is fixedly connected to the connecting frames 2-4 of the left and right frame-type solar panels 2, strengthening the connection between the root connecting frames 2-4 of the two frame-type solar panels 2 and further improving overall rigidity. Simultaneously, in the retracted state, the reinforcing beam 3 acts to press the lifting mechanism 1, working in conjunction with the pressing and releasing mechanism 4 to press and fix the lifting mechanism 1 and the frame-type solar panels 2 to the spacecraft body.

[0027] The clamping release mechanism 4 is a hot-blade unlocking device, comprising a clamping beam, clamping ropes, and a hot-blade assembly. The clamping beam is located on the retracted outer side of the passively deployable high-rigidity flexible solar array. The clamping ropes pass through the clamping beams to clamp and fix the various components of the passively deployable high-rigidity flexible solar array to the spacecraft body. The hot-blade assembly is used to cut the clamping ropes during deployment to release the clamping constraint. In this embodiment, eight clamping release mechanisms 4 are preferably used, unlocking and releasing in batches.

[0028] The constraint release mechanism 5 is a hot-blade unlocking device, comprising a support and a constraint rope. The support is fixed to the lowest foldable triangular frame in the folded state. One end of the constraint rope is connected to the support, and the other end is connected to the uppermost foldable triangular frame in the folded state. In the folded state, the constraint rope is taut, constraining the frame unit (triangular cube unit) of the frame-type solar wing 2. When deployed, the constraint rope is cut or released to release the deployment constraint on the frame-type solar wing 2. This embodiment preferably employs four constraint release mechanisms 5.

[0029] Example 2 This embodiment provides a passive deployment method for a high-stiffness flexible solar array. This method is applied to the passive deployment high-stiffness flexible solar array described in Embodiment 1. The solar array deployment process is divided into two stages: a lifting and flipping stage and a solar array deployment stage, specifically including the following steps: Step S1: After the satellite enters orbit, the compression release mechanism is unlocked and released in batches. The lifting mechanism and the second folding rod are deployed and locked under the action of their own torsion spring torque, raising the solar array to a predetermined height above the spacecraft body.

[0030] Specifically, such as Figure 2 As shown, after the satellite enters orbit, the eight clamping release mechanisms 4 are unlocked and released in batches. The hot knife assembly cuts the clamping ropes, releasing the clamping fixation on each component of the passively deployable high-rigidity flexible solar array. At this time, because the constraint rope of the constraint release mechanism 5 is still in a taut state, the frame unit of the frame-type solar array 2 is still constrained, and the mechanism will not immediately fully deploy. The drive spring 1-10 of the lifting mechanism 1 releases elastic potential energy, driving the inner scissor bar 1-4 and the outer scissor bar 1-5 to rotate around their hinge point, causing the scissor unit to deploy and lifting the top seat 1-6 and the swing bracket 1-7 to a certain distance from the spacecraft body. After the scissor unit is deployed in place, the locking pin of the locking assembly 1-11 is inserted into the locking hole under the action of elastic force, realizing the rigid locking of the lifting mechanism 1. At the same time, the second folding rod 2-5 is deployed in place and locked under the action of its own torsion spring torque. After completion, the corresponding Figure 2 The state.

[0031] Step S2: The constraint release mechanism is powered on to release the constraint on the frame unit of the frame-type solar panel. The frame unit unfolds into place and locks under its own spring torque.

[0032] Specifically, such as Figure 3 As shown, after the lifting and flipping stage is completed, power is supplied to the four constraint release mechanisms 5, and the hot knife assembly cuts the constraint ropes, releasing the constraint on the frame unit of the frame-type solar wing 2. The foldable triangular frame 2-2 of the frame-type solar wing 2 unfolds under the drive of the torsion spring 2-2-4, the first folding plate 2-2-2 unfolds 180 degrees, and the locking hook 2-2-5 hooks the lock head 2-2-6 to lock it; the first sub-folding rod of the first folding rod 2-3 unfolds 180 degrees under the drive of the torsion spring and locks. The frame unit unfolds into place and locks under its own spring torque. After completion, the corresponding... Figure 3 The state.

[0033] Step S3: Under the action of the first and second drives of the lifting mechanism, the two solar arrays achieve solar orientation tracking.

[0034] Specifically, after being deployed, under the action of the first drive 1-8 and the second drive 1-9 of the lifting mechanism 1, the turntable 1-2 is driven to rotate relative to the base 1-1 by the first drive 1-8 to achieve azimuth orientation towards the sun, and the swing bracket 1-7 is driven to rotate relative to the top seat 1-6 by the second drive 1-9 to achieve pitch orientation towards the sun, thus realizing the sun orientation and tracking of the two solar wings.

[0035] It should be noted that the specific parameters involved in the above embodiments, such as the number of scissor unit stages of the lifting mechanism 1, the number of foldable triangular frames 2-2 of the frame-type solar wing 2, and the number of clamping release mechanisms 4 and restraint release mechanisms 5, are merely illustrative examples and are not intended to limit the technical solution of this application. For example, the number of scissor unit stages of the lifting mechanism 1 can be adjusted to a single stage or multiple stages according to the lifting height requirements; the number of foldable triangular frames 2-2 can be adjusted according to the unfolded area requirements; and the number of clamping release mechanisms 4 and restraint release mechanisms 5 can be adjusted according to the clamping force requirements. As long as its function, effect, and result are substantially the same as the technical solution described in the claims of this application, they should all be considered as equivalent implementations of this application.

[0036] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0037] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A passively deployable high-rigidity flexible solar array, characterized in that, include: The lifting mechanism is used to passively deploy and lift the frame-type solar array to a set distance from the spacecraft body when the restraints are released; A swing bracket, located at the top of the lifting mechanism, is used to support the frame-type solar panel and is capable of rotating relative to the lifting mechanism; Two frame-type solar panels are symmetrically arranged on both sides of the swing bracket. Each frame-type solar panel includes a foldable frame structure and a flexible solar panel plate laid on the frame structure. A reinforcing beam is connected between the two frame-type solar panels. The reinforcing beam is used to strengthen the connection between the root connecting frames of the two frame-type solar panels and to press the lifting mechanism in the retracted state.

2. The passive deployable high-rigidity flexible solar array according to claim 1, characterized in that, The lifting mechanism includes a base, a turntable, a scissor lift unit, a top seat, and a drive spring; The base is used for fixed installation on the spacecraft body; The turntable is rotatably connected to the base; the scissor lift unit is connected between the turntable and the top seat; the drive spring is used to store elastic potential energy in the retracted state and drive the scissor lift unit to unfold when the constraint is released, so as to lift the top seat. The swing bracket is positioned above the top seat.

3. The passive deployable high-rigidity flexible solar array according to claim 1, characterized in that, Also includes: The clamping and releasing mechanism is used to clamp and fix the components of the passive deployable high-rigidity flexible solar array in the retracted state, and to release the clamping constraint when it is deployed.

4. The passive deployable high-rigidity flexible solar array according to claim 1, characterized in that, Also includes: The constraint release mechanism is used to constrain the deployment degree of freedom of the passive deployable high-rigidity flexible solar array in the retracted state and release the deployment constraint when it is deployed.

5. The passive deployable high-rigidity flexible solar array according to claim 1, characterized in that, The frame structure includes a longitudinal bar assembly, a foldable triangular frame, and a first folding bar. The foldable triangular frames are distributed on both sides of the vertical rod assembly and connected by the vertical rod assembly, with the foldable triangular frames on each side arranged in series. The first folding rod is connected between two adjacent foldable triangular frames and is used to constrain the relative position between the two adjacent foldable triangular frames when unfolded.

6. The passive deployable high-rigidity flexible solar array according to claim 3, characterized in that, The clamping and releasing mechanism includes a clamping beam, a clamping rope, and a hot knife assembly; The clamping beam is located on the retracted outer side of the passive deployable high-rigidity flexible solar array. The clamping rope passes through the clamping beam to clamp and fix each component of the passive deployable high-rigidity flexible solar array to the spacecraft body. The hot knife assembly is used to cut the clamping rope during deployment to release the clamping constraint.

7. The passive deployable high-rigidity flexible solar array according to claim 4, characterized in that, The restraint release mechanism includes a support and a restraint rope; The support is fixed to the lowest foldable triangular frame in the folded state. One end of the constraint rope is connected to the support, and the other end is connected to the uppermost foldable triangular frame in the folded state. In the folded state, the constraint rope is in a taut state, constraining the degree of freedom of the passive deployable high-rigidity flexible solar array. When deployed, the constraint rope is cut or released to release the deployment constraint on the passive deployable high-rigidity flexible solar array.

8. A method for deploying a passively deployable high-stiffness flexible solar array, based on the passively deployable high-stiffness flexible solar array according to any one of claims 1-7, characterized in that, include: Step S1: After the satellite enters orbit, the clamping and releasing mechanism is unlocked and released in batches. The lifting mechanism and the second folding rod are deployed and locked under the action of their own torsion spring torque, raising the solar array to a predetermined height above the spacecraft body. Step S2: The constraint release mechanism is powered on to release the constraint on the frame unit of the frame-type solar panel. The frame unit unfolds into place and locks under its own spring torque. Step S3: Under the action of the first and second drives of the lifting mechanism, the two solar arrays achieve solar orientation tracking.

Images