# Development of a Novel Deployable Solar Panel and Mechanism for 6U CubeSat of STEP Cube Lab-II
## 1 Introduction  
In recent years, the onboard power demand of CubeSat has steadily increased as the capability of the platform for advanced missions has significantly improved owing to advances in technology miniaturization [1; 2; 3]. To meet the onboard power demand, deployable solar panels have commonly been adopted in CubeSats that encompass the extension of surface areas for solar cell installation and also allow orientation or articulation of the panel in the sun's direction using a combination of solar array drive assembly (SADA). However, severe launch vibration loads enforce dynamic stresses and deflections on the satellite's deployable appendages [4; 5; 6]. The excessive dynamic deflection and acceleration of a deployable solar panel under a lunch vibration environment causes stress on mounted solar cells and produces an undesirable burden on the holding and release mechanism (HRM), which may eventually lead to detachment or fracture of those cells. This problem becomes more severe when a larger solar panel is adopted to meet the onboard power demand of advanced missions. Thus, the minimization of the deployable solar panel's dynamic deflection and stress under severe launch vibration environments to ensure the solar cells' structural safety is an important factor for mission success. Printed circuit board (PCB) substrate-based flight-proven deployable solar panels of various configurations have been produced owing to the advantages of speedy fabrication and easy electrical interconnection with the mounted solar cells. To date, the mechanical design strategies commonly used to minimize panel dynamic deflection under launch vibration loads have increased the solar panel eigenfrequency by including additional stiffeners made up of aluminum or fiberglass-laminate [7; 8]. For instance, ISISpace [7]has produced deployable solar panels for 6U CubeSat application, where a thin PCB made up of FR4 material of 0.18 mm thickness is stiffened by an aluminum panel. Only 14 triple-junction gallium arsenide (GaAs) solar cells from the AZUR space [9] can be mounted on the solar panel because a hold-down and release mechanism is implemented at the panel center. The ISIS 6U deployable solar panel can generate 17 W of power in LEO with a flight model mass of 720 g. [PERSON] et al. [8] developed an FR4 PCB-based deployable solar panel for 6U CubeSat, which was stiffened by using stiffeners made up of G10 high-pressure fiberglass laminate composite material. Seventeen triple-junction GaAs solar cells with 30% efficiency can be attached to the panel so that the expected power generation capacity from a single panel is 19.5 W, although the total mass of the solar panel module is 625 g. However, this strategy involves a trade-off between the solar panel's stiffness and weight, which could be disadvantage for the CubeSat platform because it has a limited mass budget. Moreover, the increased mass of the solar panel unavoidably increases panel excitation in vibration loads, which creates an adverse burden on the HRM. Instead, the carbon-fiber-reinforced plastic (CFRP) panel and honeycomb panel are comparatively lightweight and exhibit high rigidity. [PERSON] et al. [10] developed deployable solar panels for VELOX-II 6U CubeSat based on an aluminum honeycomb panel instead of the PCB to ensure the stiffness requirement specified by the launch provider. The two deployable solar panels of the VELOX-II could produce a peak power of 40.8 W. The VELOX-II CubeSat deployable solar panel module per unit mass is only 500 g. However, solar panels based on CFRP and honeycomb are relatively thick and expensive. As the poly picosatellite orbital deployer (P-POD) has a restricted lateral edge gap for solar panel accommodation [11] and as CubeSat has limited development cost, CFRP and honeycomb panels are less applicable for use in the CubeSat platform.