A thermally triggered self-deployment mechanism for solar panels on a spaceborne antenna platform

Abstract

This invention relates to the field of spacecraft structure technology and discloses a thermally triggered self-deploying mechanism for a spaceborne antenna platform's solar panel. The mechanism includes: a solar panel frame, an antenna backplate, at least one drive hinge, and at least one thermal lock. This mechanism utilizes the orbital solar radiation environment as a trigger signal: a thermal delay sleeve absorbs and accumulates solar radiation heat, causing the internal locking tongue, made of a bimetallic strip, to undergo directional bending deformation when a temperature threshold is reached, thereby disengaging from the locking seat. After the lock is released, the elastic potential energy stored in the energy storage spring is released, driving the solar panel to smoothly deploy to a predetermined angle. This invention achieves completely passive and reliable environmentally adaptive triggering and deployment, simplifies the structure, avoids reliance on electronic systems, and integrates a check mechanism to ensure stable deployment.

Description

Technical Field

[0001] This invention relates to the field of spacecraft structure technology, specifically to a thermally triggered self-deployment mechanism for a spaceborne antenna platform solar panel. Background Technology

[0002] Existing deployment mechanisms for spaceborne solar arrays often rely on complex electronic control systems. These systems issue commands via controllers at specific times to drive motors or pyrotechnics, unlocking and deploying the solar panels. While this approach is technically mature, its inherent electronic complexity introduces significant risks: a power outage or control signal interruption onboard could cause the entire deployment process to fail, resulting in the loss of critical mission functions. Furthermore, to ensure post-deployment stability, existing designs often employ additional electric or shape-memory alloy locks. These locks have limited locking force and are susceptible to loosening under long-term on-orbit vibration, potentially causing the solar panels to swing back unexpectedly, affecting the continuity of energy supply.

[0003] To address the aforementioned issues, this application proposes a thermally triggered self-deployment mechanism for a spaceborne antenna platform solar panel. Summary of the Invention

[0004] The purpose of this invention is to provide a thermally triggered self-deployment mechanism for solar panels on a spaceborne antenna platform, in order to solve the problems of low reliability caused by electronic dependence and unstable locking with the risk of swinging back in the prior art mentioned in the background.

[0005] To achieve the above objectives, the present invention provides the following technical solution: a solar panel thermally triggered self-deployment mechanism for a spaceborne antenna platform, comprising: a solar panel frame, an antenna backplate, at least one drive hinge, and at least one thermal lock; The solar cell frame has a first end and a second end, the first end of which is rotatably connected to the antenna back plate via the drive hinge, and the second end of which is releasably connected to the antenna back plate via the thermal lock. The drive hinge includes a fixed base fixed to the antenna backplate, a rotating component fixed to the first end of the solar cell frame, and an energy storage spring connecting the fixed base and the rotating component; in the locked state, the energy storage spring is pre-tightened and stores elastic potential energy. The thermal lock includes a lock base disposed on the antenna backplate, a latch disposed at the second end of the solar cell frame and capable of engaging with the lock base, and a heat delay sleeve covering the outside of the latch. The heat delay sleeve is configured to cause the latch to deform and disengage from the lock base after absorbing and accumulating heat to a threshold value. When the latch disengages from the lock base, the energy storage spring releases elastic potential energy to drive the rotating component to rotate around the fixed base, thereby triggering the solar cell frame to unfold.

[0006] Preferably, the drive hinge further includes a rotating shaft; the rotating component is rotatably mounted in the fixed base via the rotating shaft, and the energy storage spring surrounds the outer wall of the rotating shaft.

[0007] Preferably, the heat delay sleeve is a multi-layer rectangular sleeve structure, comprising, from the outside to the inside: an outer encapsulation shell with a selective thermal control coating on its outer surface; a vacuum isolation cavity surrounding the inner side of the outer encapsulation shell; a phase change material layer disposed inside the vacuum isolation cavity, composed of a material with a phase change temperature between 40°C and 60°C; and a heat insulation core layer disposed inside the phase change material layer; wherein the locking tongue is wrapped within the central hole of the heat insulation core layer.

[0008] Preferably, the latch is a thermal bimetallic strip, which is composed of a manganese-copper-nickel alloy layer with a high coefficient of thermal expansion and an Invar alloy layer with a low coefficient of thermal expansion.

[0009] Preferably, the drive hinge further includes a check mechanism; the check mechanism includes at least one movable pawl disposed on the rotating member and a ratchet tooth disposed on the inner wall of the fixed seat; the ratchet tooth has an inclined guide surface that gradually increases in height along the rotational direction in which the rotating member drives the solar cell frame to unfold.

[0010] Preferably, the movable pawl has a first end and a second end opposite to each other; its first end is hinged in the receiving groove of the rotating member, and the hinged end faces the rotation direction of the driving solar cell frame to unfold; its second end has a reset elastic member at its bottom, and the bottom of the reset elastic member is installed at the bottom of the receiving groove.

[0011] Preferably, it also includes an unfolding limiter, which includes a first stop block disposed at one end of the fixed base and a second stop block disposed on the rotating member; when the solar cell frame is unfolded to the maximum design angle, the first stop block and the second stop block abut against each other.

[0012] Preferably, the first stop is a protrusion, and the second stop is a stop that cooperates with the protrusion.

[0013] Preferably, the locking tongue is fixedly mounted to the second end of the solar cell frame via a mounting bracket.

[0014] Preferably, the drive hinge and the thermal lock are arranged in pairs and symmetrically on both sides of the solar cell frame.

[0015] Compared with the prior art, the beneficial effects of the present invention are: This mechanism employs a purely mechanical, passive drive system, with the driving force derived from an energy-storing spring pre-tensioned within the drive hinge before launch. This spring stores potential energy in its locked state; when the thermal lock is unlocked by heat, the potential energy is rapidly released, converting into torque that drives the rotating components, thus smoothly deploying the solar panel. This design completely eliminates reliance on onboard power supplies, circuitry, and active control systems, greatly simplifying the system configuration and significantly reducing its weight. Simultaneously, it fundamentally avoids deployment failures due to power outages or signal interruptions, significantly improving the mechanism's reliability and autonomy in complex space environments.

[0016] This mechanism integrates a reliable anti-return mechanism, consisting of a movable pawl and ratchet teeth on the inner wall of the fixed base. The inclined guide surface of the ratchet teeth allows the movable pawl to slide and press down during deployment, and at a predetermined angle, it is pushed by the bottom reset elastic element, stably engaging with the tooth groove. This one-way mechanical locking effectively resists the reverse torque caused by satellite maneuvers, space disturbances, or residual spring vibrations, rigidly preventing the solar panel frame from retracting. It ensures the permanent stability of the deployed posture, providing a solid foundation for continuous energy harvesting. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of the structure of the solar panel thermally triggered self-deploying mechanism of a spaceborne antenna platform in the locked state according to the present invention; Figure 2 This is a schematic diagram of the structure of the solar panel thermally triggered self-deploying mechanism of a spaceborne antenna platform in the deployed state according to the present invention; Figure 3 This is an exploded view of the driving hinge in the thermally triggered self-deploying mechanism of the solar panel on a spaceborne antenna platform according to the present invention. Figure 4 This is a cross-sectional view of the fixing seat in the thermally triggered self-deployment mechanism of the solar panel on a spaceborne antenna platform according to the present invention. Figure 5 This is an enlarged view of A in the thermally triggered self-deploying mechanism of the solar panel on a spaceborne antenna platform according to the present invention. Figure 6 This is an exploded view of the thermal lock in the thermally triggered self-deployment mechanism of the solar panel on a spaceborne antenna platform according to the present invention. Figure 7 This is a cross-sectional view of the thermal delay sleeve in the thermally triggered self-deployment mechanism of a solar panel on a spaceborne antenna platform according to the present invention. In the diagram: 1. Solar panel frame; 2. Antenna backplate; 3. Drive hinge; 4. Thermal lock; 5. First stop block; 6. Second stop block; 31. Fixing base; 32. Rotating component; 33. Rotating shaft; 34. Energy storage spring; 35. Reset elastic component; 36. Receiving groove; 37. Movable pawl; 38. Ratchet; 41. Lock seat; 42. Lock tongue; 43. Thermal delay sleeve; 44. Mounting bracket; 431. Outer enclosure; 432. Vacuum isolation cavity; 433. Phase change material layer; 434. Thermal insulation core layer; 435. Center hole. Detailed Implementation

[0018] The technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments.

[0019] Please see Figures 1-7 The present invention provides a technical solution: a solar panel thermally triggered self-deployment mechanism for a spaceborne antenna platform, comprising a solar panel frame 1, an antenna backplate 2, at least one drive hinge 3, and at least one thermal lock 4. The solar cell frame 1 has a first end and a second end, the first end of which is rotatably connected to the antenna back plate 2 via a drive hinge 3, and the second end of which is releasably connected to the antenna back plate 2 via a thermal lock 4. The drive hinge 3 includes a fixed base 31 fixed to the antenna back plate 2, a rotating member 32 fixed to the first end of the solar cell frame 1, and an energy storage spring 34 connecting the fixed base 31 and the rotating member 32; in the locked state, the energy storage spring 34 is pre-tightened and stores elastic potential energy. The thermal lock 4 includes a lock seat 41 disposed on the antenna back plate 2, a lock tongue 42 disposed at the second end of the solar cell frame 1 and capable of engaging with the lock seat 41, and a thermal delay sleeve 43 covering the outside of the lock tongue 42; the thermal delay sleeve 43 is configured to cause the lock tongue 42 to deform and disengage from the lock seat 41 after absorbing and accumulating heat to a threshold; when the lock tongue 42 disengages from the lock seat 41, the energy storage spring 34 releases elastic potential energy to drive the rotating member 32 to rotate around the fixed base 31, thereby triggering the solar cell frame 1 to unfold.

[0020] The mechanism achieves a rotatable connection between the first end of the solar panel 1 and the antenna backplate 2 via a drive hinge 3, and a releasable connection between the second end and a thermal lock 4, forming a complete retractable integrated structure. During the launch and transportation phases, in the locked state, the energy storage spring 34 in the drive hinge 3 is pre-tensioned, continuously storing the elastic potential energy needed to deploy the solar panel 1. Simultaneously, the locking tongue 42 of the thermal lock 4 engages with the locking seat 41, firmly locking the solar panel 1 in the retracted position to prevent accidental loosening. When the satellite moves from the shadow area to the sunlight area, the thermal environment undergoes a fundamental and continuous change. Stable and high-intensity solar radiation becomes a constant heat source. The thermal delay sleeve 43 outside the thermal lock 4 begins to continuously absorb solar radiation heat. When the absorbed heat accumulates to a preset threshold and penetrates the multiple thermal resistances of the thermal delay sleeve 43, it causes the internal locking tongue 42 to deform, causing the locking tongue 42 to disengage from the locking seat 41, thus officially releasing the locking state. At the moment of release, the elastic potential energy stored in the energy storage spring 34 is rapidly released, driving the rotating component 32 to rotate around the fixed base 31, thereby causing the entire solar cell frame 1 to smoothly unfold around the axis of the drive hinge 3, realizing the heat-triggered self-deployment function of the solar cell frame 1; if the satellite does not enter the sunlight area and is in the shadow area, the temperature is extremely low (approximately -150℃), and there is no continuous and stable heat source, the temperature of the locking tongue 42 cannot reach the unlocking threshold, and the mechanism always remains locked.

[0021] The drive hinge 3 also includes a rotating shaft 33; the rotating component 32 is rotatably mounted inside the fixed base 31 via the rotating shaft 33, and the energy storage spring 34 surrounds the outer wall of the rotating shaft 33. The rotating component 32 is rotatably mounted inside the fixed base 31 via the rotating shaft 33, forming a stable and reliable rotary kinematic pair, providing stable support and movement trajectory for the rotation and deployment of the solar cell frame 1, ensuring a smooth and unhindered deployment process. The energy storage spring 34 surrounds the outer wall of the rotating shaft 33, with its two ends fixedly connected to the fixed base 31 and the rotating component 32, respectively. This installation method allows the elastic force of the energy storage spring 34 to act directly on the center of rotation, efficiently converting the linear energy storage characteristics and elastic restoring force of the spring into torque to drive the rotation of the rotating component 32, thereby driving the solar cell frame 1 to deploy smoothly, ensuring the smoothness, concentration, and stability of the driving force transmission, avoiding problems such as driving force loss or transmission deviation, and ensuring the reliability of the deployment action.

[0022] The thermal delay sleeve 43 is a multi-layer rectangular sleeve structure, comprising, from the outside to the inside: an outer encapsulation shell 431 with a selective thermal control coating on its outer surface; a vacuum isolation cavity 432 surrounding the inner side of the outer encapsulation shell 431; a phase change material layer 433 disposed inside the vacuum isolation cavity 432, composed of a material with a phase change temperature between 40°C and 60°C; and a heat insulation core layer 434 disposed inside the phase change material layer 433; wherein the locking tongue 42 is encased within the central hole 435 of the heat insulation core layer 434. The thermal delay sleeve 43 achieves the core physical structure and working mechanism for delayed unlocking and shielding against instantaneous thermal interference. It adopts a multi-layer rectangular sleeve structure, comprising, from the outside to the inside: an outer encapsulation shell 431, a vacuum isolation cavity 432, a phase change material layer 433, and a heat insulation core layer 434. The locking tongue 42 is encased within the central hole 435 of the heat insulation core layer 434, and the phase change material layer 433 is composed of a material with a phase change temperature between 40°C and 60°C. During in-orbit operation, the selective thermal control coating on the outer surface of the outer casing 431 plays a crucial role. High-absorption coatings in some areas efficiently absorb solar radiation heat, while high-emissivity coatings in other areas efficiently radiate heat into the cold black space, establishing a clear heat input and output interface on the surface of the outer casing 431. When the satellite enters a solar-lit area, the absorbed heat must overcome three thermal resistance barriers in sequence: the first barrier is the vacuum isolation cavity 432, where the vacuum environment greatly hinders heat conduction and convection, allowing heat to transfer inward only through slow thermal radiation; the second barrier is the phase change material layer 433, the most critical thermal delay stage. When the material reaches its phase change point, it absorbs a large amount of latent heat to complete the phase change, during which the temperature remains almost constant, requiring a continuous input of large amounts of thermal energy to overcome this barrier; the third barrier is the thermal insulation core layer 434, composed of ultra-low thermal conductivity materials such as aerogel, which further impedes heat transfer. Only when the solar radiation lasts long enough, and the total input heat is sufficient to complete the phase change and penetrate all thermal insulation layers, can the heat finally reach and heat the central locking tongue 42. If the satellite does not enter the sunlight area and there is no continuous heat source, the heat that is short-lived or fluctuates will be absorbed, buffered and dissipated by multiple thermal resistances and cannot be transmitted to the locking tongue 42. The thermal delay sleeve 43 acts like a "thermal low-pass filter" and only allows slow and stable thermal signals to pass through, filtering out transient thermal noise and ensuring that the unlocking action only occurs in a continuous sunlight environment, thereby improving the reliability of the mechanism.

[0023] The latch 42 is a thermal bimetallic strip, composed of a high thermal expansion coefficient manganese-copper-nickel alloy layer and a low thermal expansion coefficient Invar alloy layer. When the ambient temperature reaches the preset unlocking threshold, due to the significant difference in thermal expansion between the two materials, the bimetallic strip will undergo bending deformation towards the Invar alloy layer, thereby driving the latch to disengage from the lock seat 41 and unlocking the device. The two alloys have significantly different thermal expansion coefficients, and the operation relies on the thermal deformation characteristics of the bimetallic strip. The unlocking temperature is compatible with the phase change material layer 433, and the preset unlocking threshold temperature is approximately 60°C. When the satellite enters the solar radiation zone, the heat transferred by the thermal delay sleeve 43 continuously accumulates, causing the temperature of the latch 42 to slowly rise from an extremely low temperature. When it finally stabilizes at the preset unlocking threshold of 60°C, due to the difference in thermal expansion between the manganese-copper-nickel alloy and the Invar alloy, the entire latch 42 will undergo directional bending deformation, with the deformation direction precisely towards the direction of disengagement from the lock seat 41. This deformation is an inherent physical property of the bimetallic strip. Without any electronic components, motors, or other external driving devices, it can directly provide the mechanical displacement and force required for unlocking, realizing a purely passive thermomechanical triggering. This simplifies the mechanism structure, reduces the risk of on-orbit failure, and the choice of bimetallic strip material ensures the stability and reliability of the deformation, avoiding unlocking failure or mis-locking due to fluctuations in material properties.

[0024] The drive hinge 3 also includes a check mechanism; the check mechanism includes at least one movable pawl 37 on the rotating member 32 and a ratchet 38 on the inner wall of the fixed seat 31; the ratchet 38 has an inclined guide surface that gradually increases in height along the rotational direction in which the rotating member 32 drives the solar cell frame 1 to unfold. The drive hinge 3 is equipped with a check mechanism, the core function of which is to prevent the solar cell frame 1 from accidentally retracting after it has been unfolded into place, thus ensuring the stability of the unfolded state. The check mechanism consists of two parts: at least one movable pawl 37 on the rotating member 32 and a ratchet 38 on the inner wall of the fixed seat 31 that is adapted to the movable pawl 37, wherein the ratchet 38 has an inclined guide surface that gradually increases in height along the rotational direction in which the rotating member 32 drives the solar cell frame 1 to unfold. When the solar panel 1 unfolds under the driving force of the energy storage spring 34, the rotating component 32 rotates synchronously, causing the movable pawl 37 to move together. At this time, the movable pawl 37 will contact the inclined guide surface of the ratchet 38 and slide along the guide surface. The inclined structure of the guide surface will exert downward pressure on the movable pawl 37, forcing the movable pawl 37 to be pressed down and retracted. When the rotating component 32 drives the solar panel 1 to rotate to the predetermined unfolding angle, the movable pawl 37 is exactly aligned with the tooth groove of the ratchet 38. At this time, the downward pressure of the guide surface on the movable pawl 37 disappears, and the movable pawl 37 bounces upward under its own restoring force, forming a stable engagement with the tooth groove of the ratchet 38. This engagement structure can effectively prevent the rotating component 32 from reversing in the direction of retraction of the solar panel 1, thereby reliably locking the unfolded state of the solar panel 1 and avoiding the retraction problem caused by external force interference or spring rebound, ensuring that the mechanism always maintains a stable unfolding posture in complex on-orbit environments.

[0025] The movable pawl 37 has a first end and a second end. Its first end is hinged within the receiving groove 36 of the rotating member 32, and this hinged end faces the rotational direction that drives the solar cell frame 1 to unfold. The bottom of its second end is provided with a reset elastic element 35, the bottom of which is installed at the bottom of the receiving groove 36. This hinged design better complements the guiding surface of the ratchet 38, ensuring that the movable pawl 37 can move flexibly when in contact with the ratchet 38. The bottom of its second end is provided with a reset elastic element 35, the bottom of which is fixedly installed at the bottom of the receiving groove 36, always in a pre-tightened state, continuously providing elastic reset force to the movable pawl 37. When the rotating component 32 rotates in the unfolding direction, and the movable pawl 37 contacts the inclined guide surface of the ratchet 38, the force of the guide surface overcomes the elastic force of the reset elastic element 35, forcing the movable pawl 37 to retract into the receiving groove 36 and smoothly pass over the ratchet 38. After the movable pawl 37 passes over the ratchet 38, the force of the guide surface disappears, the reset elastic element 35 releases its elastic potential energy, and pushes the movable pawl 37 to reset outward, maintaining contact with the ratchet 38, preparing for the next passing or final engagement. This structural design ensures the flexibility of the movable pawl 37's movement and the reliability of its reset, avoiding problems such as the movable pawl 37 getting stuck, unable to reset, or not engaging securely, further ensuring the working stability of the check mechanism.

[0026] It also includes an unfolding limiter, which comprises a first stop 5 disposed at one end of the fixed base 31 and a second stop 6 disposed on the rotating member 32; when the solar cell frame 1 is unfolded to its maximum designed angle, the first stop 5 and the second stop 6 abut against each other. The first stop 5 is a protrusion, and the second stop 6 is a stop that cooperates with the protrusion. The unfolding limiter consists of two cooperating components: the first stop 5 disposed at one end of the fixed base 31 and the second stop 6 disposed on the rotating member 32, wherein the first stop 5 is a protrusion structure and the second stop 6 is a stop structure adapted to the protrusion. When the solar cell frame 1 is extended under the drive of the energy storage spring 34 and approaches the maximum design angle, the second stop 6, which rotates together with the rotating component 32, will have its trajectory intersect with the first stop 5 on the fixed base 31. When the solar cell frame 1 is fully extended to the maximum design angle, the first stop 5 and the second stop 6 will make hard contact, and through mechanical interference, the rotating component 32 will be forcibly prevented from continuing to rotate, thereby precisely limiting the final extension angle of the solar cell frame 1. This effectively avoids problems such as component deformation, damage or loose connection caused by over-extension. At the same time, in conjunction with the anti-return mechanism, it jointly ensures the stability and structural integrity of the solar cell frame 1 after it is extended.

[0027] The locking tongue 42 is fixedly mounted to the second end of the solar cell frame 1 via a mounting bracket 44. The locking tongue 42 is securely fixed to the second end of the solar cell frame 1 via a rigid mounting bracket 44. This connection method ensures that the relative position between the locking tongue 42 and the solar cell frame 1 remains fixed and will not shift due to factors such as mechanical vibration or on-orbit environmental interference, ensuring the accuracy of the engagement between the locking tongue 42 and the locking seat 41. In the retracted state, the locking tongue 42 and the locking seat 41 are engaged, and the mounting bracket 44 securely locks the second end of the solar cell frame 1 to the antenna backplate 2, ensuring the reliability of the retracted state. When the satellite enters the solar radiation zone, the locking tongue 42 deforms due to heat and disengages from the locking seat 41. The mounting bracket 44 moves along with the solar cell frame 1, completely freeing the second end of the solar cell frame 1 from the constraint of the antenna backplate 2, ensuring that the solar cell frame 1 can unfold smoothly. This ensures the stability and reliability of the locking and releasing actions, avoiding mechanical failures caused by loose connections.

[0028] Both the drive hinge 3 and the thermal lock 4 are paired and symmetrically arranged on both sides of the solar cell frame 1. The core principle of this symmetrical arrangement is to achieve mechanical balance and synchronized movement during deployment. The drive hinges 3 on both sides can simultaneously provide equal and symmetrical driving torques, preventing torsional torque caused by excessive force on one side when the solar cell frame 1 is deployed, thus preventing torsional deformation of the mechanism. The thermal locks 4 on both sides can simultaneously absorb solar radiation heat under the same thermal environment and simultaneously reach the unlocking threshold, achieving synchronous unlocking. This ensures that both ends of the solar cell frame 1 can simultaneously detach from the constraint of the antenna backplate 2, avoiding asynchronous deployment and structural distortion caused by delayed unlocking on one side.

[0029] Working principle: When the device is launched with the satellite, it is in a retracted and locked state. The second end of the solar panel 1 is securely connected to the antenna backplate 2 via a thermal lock 4. At this time, the locking tongue 42 remains rigid at room temperature, and its end forms an interference fit with the lock seat 41, constituting the main mechanical lock. At the same time, the energy storage spring 34 in the drive hinge 3 has been pre-tightened, storing the elastic potential energy required for deployment, but its release path is completely blocked by the aforementioned mechanical lock. This state can withstand the severe vibration and overload during launch. In the early stages of satellite orbit insertion, it may pass through a shadowed area with extremely low temperatures, and the mechanism remains locked. When the satellite enters the sunlight area, the thermal delay sleeve 43 begins to work. The high-absorption coating area on the surface of its outer encapsulation shell 431 continuously absorbs solar radiation heat, while the high-emissivity coating area dissipates heat into space. The heat must slowly penetrate the vacuum isolation cavity 432, undergo the phase change heat absorption process of the phase change material layer 433, and finally pass through the heat insulation core layer 434. This multi-layer structure results in a significant delay in heat transfer (usually requiring continuous exposure to sunlight for tens of minutes). Only stable and continuous sunlight can allow heat to eventually reach and heat the locking tongue 42 at the center. Brief thermal disturbances cannot break through this thermal resistance network, thus preventing false triggering. When the temperature of the locking tongue 42 slowly rises to its preset unlocking threshold (approximately 60°C), the bimetallic strip undergoes directional bending deformation towards the direction of disengagement from the lock seat 41 due to the different expansion amounts of the alloys on both sides. This deformation directly causes the locking tongue 42 to mechanically disengage from the lock seat 41, releasing the main lock. Instantly, the constraint of the energy storage spring 34 in the drive hinge 3 disappears, and its stored potential energy is immediately released, converting into torque that drives the rotating component 32 to rotate around the axis 33. Driven by the spring torque, the rotating component 32 drives the solar cell frame 1 to unfold smoothly. During the unfolding process, the movable pawl 37 fixed on the rotating component 32 moves accordingly, and its free end slides, presses down, and bounces up along the inclined guide surface of the ratchet 38 on the inner wall of the fixed seat 31, generating damping and producing a meshing sound until stable engagement is achieved at a predetermined angle, forming a check mechanism to prevent rotation. At the final stage of deployment, the second stop 6 on the rotating component 32 rigidly abuts against the first stop 5 on the fixed base 31, mechanically limiting and precisely terminating the deployment action. The drive hinges 3 and thermal locks 4, symmetrically arranged on both sides of the battery rack, ensure force balance and synchronization during the deployment process. Thus, the equipment completes a fully automated, passive deployment from launch locking, on-orbit intelligent thermal criterion triggering, to final mechanical locking deployment.

[0030] Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

Claims

1. A thermally triggered self-deploying mechanism for a spaceborne antenna platform solar panel, characterized in that: include: Solar cell frame (1), antenna backplate (2), at least one drive hinge (3) and at least one thermal lock (4); The solar cell frame (1) has a first end and a second end opposite to each other. The first end is rotatably connected to the antenna back plate (2) via the drive hinge (3), and the second end is releasably connected to the antenna back plate (2) via the thermal lock (4). The drive hinge (3) includes a fixed base (31) fixed to the antenna back plate (2), a rotating part (32) fixed to the first end of the solar cell frame (1), and an energy storage spring (34) connecting the fixed base (31) and the rotating part (32); in the locked state, the energy storage spring (34) is pre-tightened and stores elastic potential energy. The thermal lock (4) includes a lock seat (41) disposed on the antenna back plate (2), a lock tongue (42) disposed at the second end of the solar cell frame (1) and capable of engaging with the lock seat (41), and a heat delay sleeve (43) covering the outside of the lock tongue (42); the heat delay sleeve (43) is configured to cause the lock tongue (42) to deform and disengage from the lock seat (41) after absorbing and accumulating heat to a threshold; when the lock tongue (42) disengages from the lock seat (41), the energy storage spring (34) releases elastic potential energy to drive the rotating member (32) to rotate around the fixed base (31) thereby triggering the solar cell frame (1) to unfold.

2. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: The drive hinge (3) also includes a rotating shaft (33); the rotating component (32) is rotatably mounted in the fixed base (31) via the rotating shaft (33), and the energy storage spring (34) surrounds the outer wall of the rotating shaft (33).

3. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: The heat delay sleeve (43) is a multi-layer rectangular sleeve structure, which includes at least the following from the outside to the inside: an outer encapsulation shell (431) with a selective thermal control coating on its outer surface; a vacuum isolation cavity (432) surrounding the inner side of the outer encapsulation shell (431); a phase change material layer (433) disposed inside the vacuum isolation cavity (432) and made of a material with a phase change temperature between 40°C and 60°C; and a heat insulation core layer (434) disposed inside the phase change material layer (433); wherein the locking tongue (42) is wrapped in the central hole (435) of the heat insulation core layer (434).

4. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 3, characterized in that: The latch (42) is a thermal bimetallic sheet, which is composed of a manganese-copper-nickel alloy layer with a high thermal expansion coefficient and an Invar alloy layer with a low thermal expansion coefficient.

5. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: The drive hinge (3) also includes a check mechanism; the check mechanism includes at least one movable pawl (37) provided on the rotating member (32) and a ratchet (38) provided on the inner wall of the fixed seat (31); the ratchet (38) has an inclined guide surface that gradually increases in the rotational direction in which the rotating member (32) drives the solar cell frame (1) to unfold.

6. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 5, characterized in that: The movable pawl (37) has a first end and a second end opposite to each other; its first end is hinged in the receiving groove (36) of the rotating member (32), and the hinged end is oriented toward the rotation direction of the driving solar cell frame (1) to unfold; its second end is provided with a reset elastic member (35) at the bottom, and the bottom of the reset elastic member (35) is installed at the bottom of the receiving groove (36).

7. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: It also includes an unfolding limiter, which includes a first stop (5) disposed at one end of the fixed base (31) and a second stop (6) disposed on the rotating member (32); when the solar cell frame (1) is unfolded to the maximum design angle, the first stop (5) and the second stop (6) abut against each other.

8. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 7, characterized in that: The first stop (5) is a protrusion, and the second stop (6) is a stop that cooperates with the protrusion.

9. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: The locking tongue (42) is fixedly installed at the second end of the solar cell frame (1) by a mounting bracket (44).

10. The self-deploying mechanism for a solar panel on a spaceborne antenna platform according to claim 1, characterized in that: The drive hinge (3) and the thermal lock (4) are both arranged in pairs and are symmetrically arranged on both sides of the solar cell frame (1).

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