Conceptual design and finite element method validation of a new type of self-locking hinge for deployable CubeSat solar panels

Abstract

Deployable mechanisms in CubeSat satellites have many problems with the system that provides the anchor position. The main defect of the traditional deployment mechanisms for solar panels in CubeSats is the lack of position system to block the back-driving of the panel when it reaches the final phase of the deployment. This generates spurious oscillations in the panel, affecting the photovoltaic process as well as generating fatigue in the mechanical elements of the mechanism (hinge or pin). In this work, the design, analysis and manufacture of a deployment mechanism for CubeSat solar panels is shown. A finite element method analysis was carried out in a hinge with an integrated blocking system as well as a double torsion spring, which can be used on CubeSats. The outcome shows the layout of the described anchor hinge and the used double-torsion spring, which provides a positive direction torque transfer. Likewise, the performed numerical analyses on the designed system, reduce the weight and optimise the geometry of the mechanism, showing its feasibility as well as the potential applications and further research in the area.

Modeling and simulation of the kinematic behavior of the deployment mechanism of solar array for a 1‐U CubeSat

“…To keep weight and spatial dimension within the standard values, 30,31 heavier electromechanical deploying drivers are replaced with a lesser weight strings as a proposed option for future possible implementation. The absence of holdback locks to prevent panels from bouncing back to compress the torsion spring that could set in vibration causes lower power generation and fatigue at the hinges Arturo et al 16 The proposed modification in this article aims at the component sizing and kinematic parameters for space and weight optimization, system stability during panel array deployment and operation, and photovoltaic surface projection /orientation towards solar rays for maximum power generation.…”

“…This condition is well achieved as the model end‐to‐end arm length is 90.5% of the standard maximum cross‐sectional dimension. A similar model launched in 2013 by Ecuadorian Civilian Space Agency (ECSA) ‐ the NEE‐01 PEGASUS, Arturo 16 have a dimension of 10 × 10 × 10 cm 3 (in launch configuration) and a mass of ˜1.26 kg having two solar panel wings with three panels on each side for a total of six panels per wing. The deployed dimensions are 10 × 10 × 75 cm 3 and 1.5 mm panel thickness.…”

“…This is an indication that the CubeSats will satisfy the maximum space and weight optimum criteria for its proper operation. The solar array assembly accounts for about 49.2% of the CubeSat weight and thus, imbalance in its operations will result to low power output and fatigue at the hinges, as reported in Reference 16.…”

“…There are, currently, five types of CubeSats: 0.5, 1, 2, 3, and 6 U. The number corresponds to the (approximate) length of the CubeSat in decimeters and the widths limited to 1decimeter, NASA 31 and Arturo 16 . The design of the structural system includes the main body (CubeSats frame), the arms, and the solar panels.…”

“…The solar array is released and deployed when activated by the heat of the sun using artificial muscle and memory metals respectively. It was placed in sun‐synchronous orbit at an altitude of 630 × 657 km 2 , inclination of 98.04 and period of 97.45 min with adorable features 99.98% pure titanium and 1.5 mm thickness, but lacks holding‐back mechanism to minimize vibration, 16 and thus exposed to high deployment induced vibration. Mirshams et al 17 designed a deployable array using Torsion spring for deployment and paraffin actuator as deployment rate controller.…”