Spacecraft Solar Array Independent Beta Axis Rotation
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Solution Overview
Problem
Conventional spacecraft solar arrays are unable to independently rotate about the β axis after deployment due to their linkage with closed cable loop systems, which synchronizes deployment but prevents independent tilting, leading to potential damage during missions like landing on asteroids or docking with other spacecraft.
Innovation Solution
A mechanism is introduced that includes a hinge with a rotation driving mechanism and a CCL disengagement mechanism, allowing the solar array to rotate about the β axis independently by decoupling from the closed cable loop system, enabling independent re-orientation of the solar array after deployment.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Stability of the object's composition
If a closed cable loop system is used to synchronize deployment of the solar array, then deployment synchronization is improved, but independent articulation about the β axis is prevented
Solution Approach 1:
The solar array deployment system is segmented into two independent functional subsystems: the CCL system handles deployment synchronization along the α axis, while the hinge mechanism with motor enables independent articulation about the β axis. This segmentation allows each subsystem to perform its specific function without interfering with the other, resolving the contradiction between deployment synchronization and independent articulation.
Solution Approach 2:
The hinge mechanism acts as an intermediary component between the CCL system and the solar array panels. It receives the synchronized deployment motion from the CCL system while simultaneously enabling independent β axis rotation through its integrated motor, thus mediating between the two conflicting requirements.
2Stability of the object's composition
If the solar array is latched in a fixed position after deployment, then structural stability is improved, but the ability to re-orient for different mission requirements is lost
Solution Approach 1:
The system transitions from a static latched configuration to a dynamic reconfigurable system. The hinge mechanism with motor enables the solar array to dynamically adjust its orientation about the β axis while maintaining structural stability through controlled actuation, allowing the system to adapt to different mission requirements such as landing on asteroids or docking with other spacecraft.
Solution Approach 2:
The system changes the orientation parameter of the solar array dynamically. By controlling the motor in the hinge mechanism, the β axis rotation angle can be adjusted to optimize solar panel orientation for different gravitational environments and mission scenarios, while the system maintains structural integrity through controlled parameter changes.
3Area of moving object
If the solar array is deployed to extend a substantial distance from the spacecraft body, then solar collection area is improved, but vulnerability to damage during maneuvers is increased
Solution Approach 1:
The solar array is deployed to a substantial distance to maximize solar collection area, but the hinge mechanism with motor enables dynamic repositioning capability. This allows the array to be re-oriented or retracted during maneuvers such as landing or docking, reducing vulnerability to damage while maintaining large solar collection area during normal operations.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This solution allows for post-deployment β axis rotation of the solar array, enhancing its ability to tilt and re-orient without affecting other structures, thereby reducing the risk of damage during various space missions and enabling efficient solar array positioning in diverse gravitational environments.
Implementation Method 1
a distal appendage rotatably coupled with the proximal appendage by way of a hinge connecting a distal end of the proximal appendage with a proximal end of the distal appendage, the hinge being configured such that the distal appendage is rotatable with respect to the proximal appendage about a β axis that is not aligned with the α axis
Implementation Method 2
maximize collection of solar radiation by photovoltaic solar cells disposed on a surface of the solar array
Data Source
AI summary
Techniques for two-axis articulation of a deployed spacecraft solar array are disclosed. In one aspect, an arrangement mechanically coupling a solar array with a sidewall of a body of a spacecraft includes a proximal appendage, a distal appendage rotatably coupled with the proximal appendage by way of a hinge, and a closed cable loop (CCL) system coupled with the proximal appendage and the distal appendage. In an on-orbit configuration, a long axis of the proximal appendage defines an α axis that is substantially orthogonal to the sidewall. The hinge includes CCL disengagement mechanism configured to de-couple the CCL system from the proximal appendage and the distal appendage and a rotation driving mechanism configured to cause the distal appendage to rotate about a β axis when the proximal appendage is in the on-orbit configuration, the β axis being not aligned with the α axis.