A control method of a large-size flexible solar wing driving device of a flat satellite, a storage medium and computer equipment

Abstract

The application provides a control method of a flat satellite large-size flexible solar wing driving device, the driving device comprising a satellite body, a double-shaft SADA, a swing rod, a hinge, a compression release mechanism and a ROSA solar wing; the control method comprising: sending a release signal to the compression release mechanism, releasing the compressed solar wing after the compression release mechanism receives the release signal, driving the solar wing to do passive unfolding of a first angle through the hinge at one end of the swing rod; calculating a second angle that the double-shaft SADA needs to rotate according to a target angle and the first angle; sending a rotating signal to the double-shaft SADA, driving the solar wing to continue rotating the second angle after the double-shaft SADA receives the rotating signal, and sending a rotating-in-place signal; sending an unfolding signal to the solar wing, and starting the battery array surface unfolding of the solar wing until completely unfolded. The control method improves the energy utilization efficiency of the whole satellite platform.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of satellite control, and in particular to a control method of a large-size flexible solar wing driving device of a flat plate satellite, a storage medium and computer equipment. BACKGROUND

[0002] The solar wing is a core energy component of a satellite in orbit, and its reliable deployment and accurate positioning are directly related to the mission life of the satellite. The size of the solar wing directly affects the power generation and the motion form, and directly affects the size of the satellite envelope. The traditional large-size flexible solar wing is mostly used in large cubic satellites and is placed on the outer surface of the satellite and is deployed outward after being directly unlocked. This method has the following disadvantages: a large amount of energy is consumed for full-time active driving, which is not conducive to the satellite platform with energy shortage; the requirement for the rotation angle range is high, which affects the product reliability; the space requirement for the satellite is large, which is not suitable for flat plate satellites and is not suitable for large-scale satellite systems. In addition, for large-size flexible structures, the stability control during the deployment process and the vibration suppression after the deployment also need to be solved.

[0003] Therefore, there is a need in the art for a high-efficiency and reliable control method that can fully utilize the initial kinetic energy released and smoothly connect with active driving, while adapting to the new flat plate satellite structure layout, reducing the overall size of the satellite, realizing large-scale stacked launch, and optimizing energy use and control performance. SUMMARY

[0004] The purpose of the present application is to provide a control method of a large-size flexible solar wing driving device of a flat plate satellite, a storage medium and computer equipment, to solve the problems of excessive dependence on active driving in the solar wing deployment process, high energy consumption, limited installation position, excessive space requirement, high requirement for SADA performance and non-optimal connection in the deployment process in the prior art.

[0005] In one aspect, the present application provides a control method of a large-size flexible solar wing driving device of a flat plate satellite, the driving device comprising: a satellite body, a double-axis SADA, a swing rod, a hinge, a compression and release mechanism and a ROSA solar wing; the solar wing is installed in the concave space in the middle of the satellite body, the compression and release mechanism is arranged at both ends of the concave space for compressing or releasing the solar wing, the double-axis SADA is arranged at the side of the satellite body, the swing rod connects the solar wing and the double-axis SADA, and the hinge is arranged at the connection between the swing rod and the double-axis SADA.

[0006] The control method comprises:

[0007] Step S101: Send a release signal to the clamping release mechanism. After receiving the release signal, the clamping release mechanism releases the clamped solar panel. The solar panel is passively deployed to a first angle by the hinge at one end of the swing arm.

[0008] Step S102: Calculate the second angle that the dual-axis SADA needs to rotate based on the target angle and the first angle;

[0009] Step S103: A rotation signal is sent to the dual-axis SADA. After receiving the rotation signal, the dual-axis SADA drives the solar array to continue rotating at the second angle and sends a rotation completion signal. A deployment signal is sent to the solar array, and the solar array begins to deploy the battery array until it is fully deployed.

[0010] Preferably, the solar array rotates from the center of the celestial body toward a direction away from the celestial body.

[0011] Preferably, the first angle is the maximum rotation angle of the hinge. This ensures that the mechanical characteristics of the hinge are fully utilized during the passive deployment phase, achieving an initial and reliable deployment action.

[0012] Preferably, the target angle is set to 180 degrees. This allows the solar array to eventually deploy to a position substantially parallel to the sidewalls of the star, maximizing the area exposed to sunlight.

[0013] Preferably, the first axis of the dual-axis SADA is connected to the celestial body, and the second axis of the dual-axis SADA is connected to the hinge. This dual-axis SADA structure provides two degrees of freedom of motion, enabling the solar array to perform more complex attitude adjustments.

[0014] In some embodiments, a fiber Bragg grating sensor network is arranged on the surface of the solar array. The fiber Bragg grating sensors detect the strain distribution on the solar array surface, enabling real-time monitoring of the deformation of the flexible structure. This provides a data foundation for monitoring the deployment process and subsequent vibration control.

[0015] In some embodiments, piezoelectric ceramic sensors are arranged on the surface of the solar array to monitor the vibration state of the solar array in real time. The piezoelectric ceramic actuator acts as an active control element, suppressing vibration by applying a reverse torque. This achieves active suppression of vibration after the solar array is deployed.

[0016] Preferably, there are two solar panels, which are installed side by side in the recessed space in the middle of the star.

[0017] In a second aspect, the present invention provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the control method described in any one of the first aspects above.

[0018] Thirdly, the present invention provides a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the control method described in any one of the first aspects above.

[0019] In this embodiment, the solar array drive device features a unique design that places the solar array in a recessed space in the middle of the satellite. By dividing the solar array deployment into two stages—a "passive deployment" stage of the clamping and release mechanism and an "active drive" stage of the dual-axis SADA—it fully utilizes the initial mechanical energy after the solar array is released to complete the initial deployment. This significantly reduces the drive energy and rotational stroke required by the dual-axis SADA, improves the energy efficiency of the entire satellite platform, reduces restrictions on the SADA installation location and rotation angle, thereby increasing the reliability of SADA use. It also provides the ability to install large-size ROSA solar arrays, which is advantageous for flat-panel satellites, increases the satellite's mission capabilities, and reduces the satellite's envelope space, enabling the satellite to stack multiple satellites for a single launch.

[0020] It should be understood that the above general description and the following specific embodiments are merely exemplary and illustrative, and do not limit the scope of the invention. Attached Figure Description

[0021] The accompanying drawings, which are part of the specification of this invention, illustrate exemplary embodiments of the invention. The drawings, together with the description in the specification, serve to illustrate the principles of the invention.

[0022] Figure 1 This is a flowchart illustrating the control method for a large-size flexible solar array drive device for a flat-panel satellite according to an embodiment of the present invention.

[0023] Figure 2 This is a schematic diagram of the structure of the large-size flexible solar array drive device for a flat-panel satellite in the embodiment of the present invention, with the solar array deployed to the first angle.

[0024] Figure 3 This is a schematic diagram of the structure of the large-size flexible solar array drive device for flat-panel satellite in the solar array retracted state according to an embodiment of the present invention.

[0025] Figure 4 This is a schematic diagram of the compression and release mechanism in the large-size flexible solar array drive device of the flat satellite in an embodiment of the present invention.

[0026] Figure 5 This is a schematic diagram of the structure of the large-size flexible solar array drive device for a flat-panel satellite in the embodiment of the present invention, when the solar array is deployed to the target angle.

[0027] Explanation of reference numerals in the attached figures:

[0028] 1-Star body; 11-Recessed space; 2-Compression release mechanism; 21-Memory alloy expansion joint assembly; 22-Star support; 23-Y-type gripper; 3-Solar wing; 31-Drum; 32-Baseboard; 4-Dual-axis SADA; 5-Swing rod; 51-Hinge. Detailed Implementation

[0029] The features and exemplary embodiments of various aspects of the present invention will now be described in detail. To make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only configured to explain the present invention and to exemplify the principles of the present invention, and are not configured to limit the present invention. In addition, the structural components in the drawings are not necessarily drawn to scale. For example, the dimensions of some structural components or regions in the drawings may be enlarged for other structural components or regions to aid in the understanding of the embodiments of the present invention.

[0030] The directional terms used in the following description refer to the directions shown in the figures and are not intended to limit the specific structure of the embodiments of the present invention. In the description of the present invention, it should be noted that, unless otherwise stated, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand the specific meaning of the above terms in the present invention according to the specific circumstances.

[0031] Furthermore, the terms "comprising," "including," "having," or any other variations thereof are intended to cover non-exclusive inclusion, such that a structure or component that includes a list of elements includes not only those elements but also other structural elements that are not expressly listed or inherent to the structure or component. Without further limitations, an element defined by the phrase "comprising..." does not exclude the presence of other identical elements in the article or apparatus that includes the element.

[0032] Spatial relation terms such as "below," "under," "under," "low," "above," "on," and "high" are used for descriptive convenience to explain the positioning of one element relative to a second element, indicating that these terms are intended to cover different orientations of the device, in addition to those different from those shown in the figure. Furthermore, phrases such as "one element on / below another element" can indicate that two elements are in direct contact, or that there are other elements between the two elements. In addition, terms such as "first" and "second" are also used to describe individual elements, areas, parts, etc., and should not be considered limiting. Similar terms are used throughout the description to refer to similar elements.

[0033] It will be apparent to those skilled in the art that the present invention can be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention.

[0034] Figure 1 This is a flowchart of the control method for a large-size flexible solar array drive device for a flat-panel satellite according to an embodiment of the present invention; Figure 2 This is a schematic diagram of the structure of the large-size flexible solar array drive device for a flat-panel satellite in the embodiment of the present invention when the solar array is deployed to the first angle. Figure 3 This is a schematic diagram of the structure of the large-size flexible solar array drive device for flat-panel satellite in the solar array retracted state according to an embodiment of the present invention; Figure 4 This is a schematic diagram of the clamping and releasing mechanism in the large-size flexible solar array drive device of the flat satellite in an embodiment of the present invention; Figure 5 This is a schematic diagram of the structure of the large-size flexible solar array drive device for a flat-panel satellite in the embodiment of the present invention, when the solar array is deployed to the target angle.

[0035] Example 1: Control Method

[0036] like Figures 1 to 5 As shown, the control method of the present invention is applied to a large-size flexible solar array drive device for a flat-panel satellite.

[0037] like Figure 2 As shown, Star 1 is the main structure of the satellite, and a recessed space 11 is provided at its center. This recessed space 11 is used to accommodate and house the ROSA solar array 3 during the launch phase, so that it does not protrude from the upper surface of Star 1, thereby forming a flat or nearly flat top profile.

[0038] Two sets of clamping and releasing mechanisms 2 are provided, located at both ends of the length of the recessed space 11, and fixedly installed on the star body 1. For example... Figure 4 As shown, the clamping and releasing mechanism 2 specifically includes an on-board support 22, a Y-shaped gripper 23, and a shape memory alloy expansion joint assembly 21. The on-board support 22 is fixed to the satellite body 1, and the Y-shaped gripper 23 is adapted to the shape of the drum 31 of the ROSA solar array 3. The shape memory alloy expansion joint assembly 21 passes through the on-board support 22 and the Y-shaped gripper 23, and before launch, a preload is applied to firmly clamp the drum 31 between the on-board support 22 and the Y-shaped gripper 23. Upon receiving the unlocking command, the shape memory alloy inside the shape memory alloy expansion joint assembly 21 is energized, deforms, and disengages from the locking device, thereby releasing the constraint on the drum 31 and releasing the ROSA solar array 3.

[0039] The ROSA solar panel 3 is housed within the recessed space 11 and is constrained by the clamping release mechanism 2. For example... Figure 3As shown, the ROSA solar array 3 includes a roll 31, a substrate 32, and a solar array. The solar array is made of flexible material and is wound around the roll 31 in its retracted state. The substrate 32 is the main support structure for the solar array, and the roll 31 is fixedly installed at the base of the substrate 32.

[0040] The dual-axis SADA 4 is fixedly mounted on the side of the star body 1. The dual-axis SADA 4 has a first axis and a second axis. The housing of the first axis is fixedly connected to the star body 1, providing a mounting base for the entire drive mechanism. The second axis is connected to a hinge 51 at one end of the pendulum 5. The dual-axis SADA 4 is capable of driving the second axis and the pendulum 5 connected thereto to rotate about two orthogonal axes.

[0041] One end of the swing arm 5 is connected to the second axis of the dual-axis SADA 4 via hinge 51, and the other end is connected to the base plate 32 of the ROSA solar array 3. Hinge 51 allows for a certain degree of freedom of movement between the swing arm 5 and the dual-axis SADA 4. Hinge 51 typically incorporates a torsion spring or has a limiting function.

[0042] The solar array drive device has high space utilization and a compact structure. By setting a recessed space 11 in the middle of the main body 1 and storing the solar array in the space, the external envelope size of the satellite in the launch state is greatly reduced, and a highly integrated design is achieved.

[0043] like Figure 1 As shown, the execution flow of this control method is as follows:

[0044] Step S101: Trigger release and passive deployment.

[0045] After the satellite successfully enters orbit, the onboard computer, such as the controller, sends a release signal to the clamping release mechanism 2. Upon receiving the signal, the clamping release mechanism 2 executes an unlocking action, such as detonating a pyrotechnic device or triggering a shape memory alloy, to release the restraints on the ROSA solar array 3.

[0046] Once freed, the ROSA solar wing 3 begins passive deployment, propelled by a drive source (such as a torsion spring) built into the hinge 51. The ROSA solar wing 3 rotates via the swing arm 5 around the axis of the hinge 51 in a predetermined direction, that is, from the center of the star 1 towards a direction away from the star.

[0047] During this stage, the ROSA solar array 3 will deploy to a first angle θ1. As a preferred embodiment, this first angle θ1 is limited by the mechanical structure of the hinge 51, that is, after deploying to the maximum rotation angle of the hinge 51, it is stopped by a limiting device. Figure 2This diagram illustrates the state of the ROSA solar array 3 after it has deployed to the first angle θ1. This process does not consume power from the dual-axis SADA 4. A microswitch sensor is installed on the clamping and releasing mechanism 2 to send a solar array unlocking and separation signal to the onboard computer. A microswitch or angle sensor can be installed on the hinge 51 to send a positioning command to the onboard computer after rotating to the correct position.

[0048] Step S102: Calculate the driving angle.

[0049] The controller has a preset target angle θ. target This angle is the final angle required for the ROSA solar array 3 to function properly. In one embodiment, the target angle θ is required for the solar array to fully deploy and achieve maximum illumination area. target It is set to 180 degrees. In one embodiment, the selected diagram is 180 degrees. This scheme's capability includes, but is not limited to, 180 degrees. This angle value is calculated based on the satellite's capabilities and usage requirements.

[0050] The controller then calculates: the second angle θ2 = θ target - θ1. This second angle θ2 is the remaining angle that the dual-axis SADA 4 needs to drive the ROSA solar array 3 to rotate.

[0051] Step S103: Active driving and full deployment.

[0052] The controller sends a rotation signal to the dual-axis SADA 4, which contains instructions to rotate by a second angle θ2.

[0053] The dual-axis SADA 4 is activated, driving the ROSA solar array 3 to continue rotating in the same direction as the passive deployment. Through hinge 51 and lever 5, the partially deployed solar array 3 substrate continues to rotate by a second angle θ2. During rotation, the solar array 3 moves from the recessed space 11 in the center of the star 1 towards the side away from the star 1. When the dual-axis SADA 4 drives the solar array 3 to reach the predetermined position (i.e., the total angle reaches the target angle θ), the rotation continues. target When the rotation is in place, a signal is issued.

[0054] Subsequently, the control computer sends a deployment signal to the ROSA Solar Array 3 itself. The deployment mechanism inside the Solar Array 3 begins to operate, gradually unfolding the folded flexible battery array until the battery array is completely flattened and enters the working state.

[0055] This control method can meet the motion angle requirements of different solar panel installation positions, and complete different rotation strokes according to different stages of solar panel deployment. This improves the energy utilization efficiency of the entire satellite platform, reduces the requirements for satellite SADA capabilities, and reduces the satellite envelope space. The two-stage deployment strategy of "passive deployment with clamping and releasing mechanism + active drive with dual-axis SADA" combines the reliability of passive deployment with the precision of active drive, ensuring the smoothness and controllability of the large flexible solar panel deployment process.

[0056] To further improve system performance, a monitoring and control system can be integrated into the battery array or support structure of the solar array 3.

[0057] One approach is to deploy a fiber Bragg grating sensor network. Multiple fiber Bragg grating sensors are attached to or embedded in the surface of the solar array 3 or key load-bearing components in a network form. During and after deployment in step S103, these sensors sense the strain distribution on the surface of the solar array 3 in real time, transmit the data back to the onboard computer, and can be used to monitor the stability of the deployment process, determine whether it has been fully deployed, and monitor the deformation of the flexible structure caused by thermal load or maneuvering in orbit.

[0058] Another approach is to use piezoelectric ceramics as sensing and actuation elements. Piezoelectric ceramic sensors are placed on the surface of solar array 3 to monitor vibration states (such as characteristic frequencies and amplitudes). When harmful vibrations are detected, the control algorithm drives piezoelectric ceramic actuators placed at specific locations to generate a reverse torque or force, actively counteracting the vibration energy, thereby suppressing the vibration of solar array 3 and ensuring the stability of the satellite's attitude.

[0059] In this embodiment, by introducing a fiber optic grating sensing network and / or a piezoelectric ceramic sensor, real-time monitoring and active suppression of the deformation and vibration state of the flexible solar array structure are achieved, significantly improving the attitude stability and reliability of the satellite during its on-orbit operation.

[0060] In a preferred embodiment of the present invention, there are two ROSA solar arrays 3, symmetrically arranged side-by-side in the recessed space 11 in the middle of the satellite body 1. The control method described above can be executed in parallel for these two solar arrays, achieving a balance of the overall satellite's rotational inertia, which is beneficial for satellite attitude control. Alternatively, they can be deployed sequentially according to a predetermined order. The two solar arrays 3 are housed within the recessed space of the satellite body, suitable for flat-panel satellite configurations, with a compact layout. After release, they are driven to unfold via a lever 5 and a dual-axis SADA4. The structural design is reasonable and suitable for the installation and control of large-size flexible solar arrays.

[0061] The control method in this embodiment clearly defines the complete sequence and connection logic from release to full deployment. The passive deployment phase provides a definite initial state for subsequent active actuation, making the entire deployment process smoother and more coherent. This avoids the startup shock or uncertainty that may result from relying entirely on a single drive source, improving mission reliability and making the control process smooth and reliable. Furthermore, the control method in this embodiment reduces restrictions on the SADA installation location and rotation angle, thereby increasing the reliability of SADA use. It also provides the capability for mounting large-size ROSA solar panels on flat-panel satellites, increasing the satellite's mission capabilities, while reducing the satellite's envelope space, enabling the stacking of multiple satellites for a single launch.

[0062] For dual-axis SADA, this embodiment proposes multiple drive modes (dual-axis coordination, single-axis drive), which can select the optimal drive strategy based on factors such as the actual working conditions of the satellite in orbit, energy level, and mechanical lifespan, thereby enhancing the adaptability and redundancy of the system, making the control strategy more flexible, and enabling the solar array to have the ability to have the best lighting conditions in non-SSO orbits, thus increasing the satellite's range of use.

[0063] The control method in this embodiment ensures that the solar array can ultimately reach the predetermined target angle (such as 180 degrees when fully deployed) by accurately calculating the remaining angle after passive deployment and performing closed-loop or open-loop position control on the dual-axis SADA, thus meeting the high-precision requirements for sun orientation and satellite attitude control.

[0064] Example 2: Storage Media and Computer Equipment

[0065] The present invention also provides a computer-readable storage medium. This medium can be any tangible medium such as a read-only memory (ROM), random access memory (RAM), flash memory, or hard disk of an onboard computer. It stores a computer program (instructions) that, when executed by a processor (such as the CPU of a spacefaring management computer or attitude control computer), can automatically implement the steps of any of the control methods described in Embodiment 1 above.

[0066] Similarly, the present invention provides a computer device. This device includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the program, it can implement the control method described in any one of the first embodiments above. Specifically, this computer device can be integrated into a satellite's onboard computer system.

[0067] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A control method for a large-size flexible solar array drive device for a flat-panel satellite, characterized in that, The driving device includes: Star, dual-axis SADA, lever, hinge, clamping release mechanism and ROSA solar array; The solar array is installed in the recessed space in the middle of the star body. The clamping and releasing mechanism is located at both ends of the recessed space for clamping or releasing the solar array. The dual-axis SADA is located on the side of the star body. The pendulum connects the solar array and the dual-axis SADA. The hinge is located at the connection between the pendulum and the dual-axis SADA. There are two solar arrays, which are installed side by side in the recessed space in the middle of the star body. The control method includes: Step S101: Send a release signal to the clamping release mechanism. After receiving the release signal, the clamping release mechanism releases the clamped solar panel. The solar panel is passively deployed to a first angle by the hinge at one end of the swing arm. Step S102: Calculate the second angle that the dual-axis SADA needs to rotate based on the target angle and the first angle; Step S103: A rotation signal is sent to the dual-axis SADA. After receiving the rotation signal, the dual-axis SADA drives the solar array to continue rotating at the second angle and sends a rotation completion signal. A deployment signal is sent to the solar array, and the solar array begins to deploy the battery array until it is fully deployed.

2. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 1, characterized in that, The solar array rotates in the direction from the center of the celestial body away from it.

3. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 2, characterized in that, The first angle is the maximum rotation angle of the hinge.

4. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 1, characterized in that, The target angle is set to 180 degrees.

5. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 1, characterized in that, The first axis of the dual-axis SADA is connected to the celestial body, and the second axis of the dual-axis SADA is connected to the hinge.

6. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 5, characterized in that, A fiber optic grating sensor network is deployed on the surface of the solar array. The fiber optic grating sensors detect the strain distribution on the surface of the solar array, thereby enabling real-time monitoring of the deformation of the flexible structure.

7. The control method for the large-size flexible solar array drive device of a flat-panel satellite according to claim 5, characterized in that, Piezoelectric ceramic sensors are arranged on the surface of the solar array to monitor the vibration state of the solar array in real time. The piezoelectric ceramic actuator acts as an active control element to suppress vibration by applying a reverse torque.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the program is executed by the processor, it implements the control method as described in any one of claims 1-7.

9. A computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the control method as described in any one of claims 1-7.

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