Development of a Novel Deployable Solar Panel and Mechanism for 6U Cubesat of STEP Cube Lab-II

aerospace

Article Development of a Novel Deployable Solar Panel and Mechanism for 6U CubeSat of STEP Cube Lab-II

Shankar Bhattarai 1, Ji-Seong Go 1, Hongrae Kim 2 and Hyun-Ung Oh 1,*

1 Space Technology Synthesis Laboratory, Department of Smart Vehicle System Engineering, Chosun University, 309, Pilmun-daero, Dong-gu, Gwangju 61452, Korea; [email protected] (S.B.); [email protected] (J.-S.G.) 2 Soletop Co. Ltd., 409 Expo-ro, Yuseong-gu, Daejeon 34051, Korea; [email protected] * Correspondence: [email protected]

Abstract: The structural safety of solar cells mounted on deployable solar panels in the launch vibration environment is a signiﬁcant aspect of a successful CubeSat mission. This paper presents a novel highly damped deployable solar panel module that is effective in ensuring structural protection of solar cells under the launch environment by rapidly suppressing the vibrations transmitting through the solar panel by constrained layer damping achieved using printed circuit board (PCB)- based multilayered thin stiffeners with double-sided viscoelastic tapes. A high-damping solar panel demonstration model with a three-pogo pin-based burn wire release mechanism was fabricated and tested for application in the 6U CubeSat “STEP Cube Lab-II” developed by Chosun University, South Korea. The reliable release function and radiation hardness assurance of the mechanism in an in-orbit environment were conﬁrmed by performing solar panel deployment tests and radiation tests,

respectively. The design effectiveness and structural safety of the proposed solar panel module were validated by launch vibration and in-orbit environment tests at the qualiﬁcation level.

Citation: Bhattarai, S.; Go, J.-S.; Kim, H.; Oh, H.-U. Development of a Keywords: STEP Cube Lab-II CubeSat; solar panel module; multilayered stiffener; holding and Novel Deployable Solar Panel and release mechanism; launch vibration Mechanism for 6U CubeSat of STEP Cube Lab-II. Aerospace 2021, 8, 64. https://doi.org/10.3390/ aerospace8030064 1. Introduction In recent years, the onboard power demand of CubeSat has steadily increased as Academic Editor: Paolo Tortora the capability of the platform for advanced missions has significantly improved owing to advances in technology miniaturization [1–3]. To meet the onboard power demand, Received: 5 February 2021 deployable solar panels have commonly been adopted in CubeSats that encompass the Accepted: 1 March 2021 extension of surface areas for solar cell installation and also allow orientation or articulation Published: 5 March 2021 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 Publisher’s Note: MDPI stays neutral satellite’s deployable appendages [4–6]. The excessive dynamic deflection and acceleration with regard to jurisdictional claims in of a deployable solar panel under a lunch vibration environment causes stress on mounted published maps and institutional affil- iations. 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 Copyright: © 2021 by the authors. solar cells’ structural safety is an important factor for mission success. Licensee MDPI, Basel, Switzerland. Printed circuit board (PCB) substrate-based flight-proven deployable solar panels of This article is an open access article various configurations have been produced owing to the advantages of speedy fabrication distributed under the terms and conditions of the Creative Commons and easy electrical interconnection with the mounted solar cells. To date, the mechanical Attribution (CC BY) license (https:// design strategies commonly used to minimize panel dynamic deflection under launch creativecommons.org/licenses/by/ vibration loads have increased the solar panel eigenfrequency by including additional 4.0/). stiffeners made up of aluminum or fiberglass-laminate [7,8]. For instance, ISISpace [7]

Aerospace 2021, 8, 64. https://doi.org/10.3390/aerospace8030064 https://www.mdpi.com/journal/aerospace Aerospace 2021, 8, 64 2 of 24

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. Park 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 disadvantageous 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. Lim 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. Recently, to overcome the above- mentioned technical issues, multiple HRMs or an HRM with an additional launch restraint mechanism has been applied to provide surplus mechanical fixation points on a solar panel to minimize dynamic deflections [12,13]. For example, GomSpace [12] developed a multi- array deployable solar panel made up of aluminum (AL6082 T6-51) for application in 3U and 6U CubeSats, where two sleds with spring-based burn wire cutting HRMs were used to reduce the dynamic response of the panel under launch environments. The maximum power generation capacity from a single panel is 20.7 W. The mass of the GomSpace deployable solar panel module was 449 g. In addition, various configurations of mass- effective multi-array-based deployable solar panels made up of graphite have recently been proposed by MMA design LLC [13], in which an additional launch restraint system was implemented in conjunction with a burn wire cutting release mechanism. However, the implementation of multiple HRMs or an additional launch restraint device in a single panel could increase system complications, solar panel development costs, and minimize the accessible accommodation area for solar cells on the panel. Furthermore, HRMs based on a burn wire release technique have been widely utilized in deployable solar panels of CubeSats owing to their simplicity, relatively low cost, and ease of mechanism reset [14]. However, as the solar panel size is increasing to meet the increased power demands of forthcoming advanced space missions, the mechanical design of conventional wire cutting release mechanisms should be improved to ensure high holding capability and to guarantee secure reliable release action in an in-orbit environment. To overcome the aforementioned drawbacks of the current state-of-the-art mechanical design strategies for minimizing solar panel dynamic deflection, we focused on constrained layer damping by the implementation of multilayered thin stiffeners with double-sided viscoelastic tapes. The applications of viscoelastic materials for structural vibration control have been widely studied and practiced in space engineering fields owing to their simplicity and cost-effectiveness. For instance, Steinberg [15] introduced constrained layer damping with a viscoelastic material as a mechanical engineering technique for vibration suppression on plate and beam structures. Minesugi et al. [16] examined the feasibility and applicability of polyimide tape with viscous lamina for vibration attenuation of onboard electrical devices or subsystems of a small satellite under launch vibration loads. Park et al. [17] Aerospace 2021, 8, 64 3 of 24

evaluated a PCB employing multilayered stiffeners attached by viscoelastic acrylic tapes in a wedge lock to enhance the fatigue life of solder joints under vibration environments. Bhattarai et al. [18] investigated the effectiveness of the constrained layer damping strategy for launch vibration attenuation on a panel by varying the number of stiffener attachment conditions. After the successful operation of 1U sized nano-class satellite Space Technology Exper- imental Project CubeSat Laboratory (STEP Cube Lab) [19] for technology demonstration missions in 2017, the Space Technology Synthesis Laboratory (STSL) of Chosun University is developing multispectral earth observation 6U CubeSat “STEP Cube Lab-II” as part of the 2019 cube satellite contest hosted by the Korea Aerospace Research Institute (KARI) that is financially supported by the Ministry of Science and ICT (MSIT) of the Republic of Korea. The CubeSat is scheduled to be launched into space in the fourth quarter of 2022 using the Korea Space Launch Vehicle (Nuri). The main objective of STEP Cube Lab-II is to carry out the nation’s first multi-band earth observation mission on various targets of the Korean peninsula through the CubeSat platform. The secondary objectives are the in-orbit verification of domestically developed space technologies for forthcoming space programs. The highly damped viscoelastic multilayered stiffener-based solar array (VMLSA) combined with an optimized pogo pin-based HRM (P-HRM) is one of the technologies to be validated by the STEP Cube Lab-II mission. In this study, a novel high-damping deployable solar panel module with a three- pogo pin-based burn wire cutting HRM was fabricated and experimentally tested for application in STEP Cube Lab-II 6U CubeSat. The key advantages of the solar panel module proposed herein are that the panel’s dynamic stresses and deflections in vibration loads can be effectively attenuated or minimized owing to the high damping characteristics accomplished by shear deformation of acrylic material of adhesive tapes with multilayered thin stiffeners. The solar panel’s holding and release actions were accomplished using a three-pogo pin-based wire cutting release mechanism. The mechanism has a high loading capability, a simpler electrical system, reliable release functionality, and ease in the wire knotting process that overcomes the limitations and drawbacks of conventional burn wire triggering release mechanisms. The functionality of the proposed mechanism was validated using solar panel deployment tests under various test environments. In addition, the radiation tests of electrical components used in the mechanism, such as the total ionizing dose (TID) as well as single event effect (SEE), were performed to ensure the radiation hardness of the P-HRM. The solar panel’s natural frequency and damping ratio in a rigidly mounted condition were obtained by performing free-vibration tests in an ambient room temperature environment. Furthermore, the design effectiveness and structural safety under launch loads were validated through sine and random vibration tests at the qualification level. To validate the structural safety and reliable release functionality of the solar panel module in an orbit environment, a thermal vacuum (TV) test was performed. Optical microphotographs of the solar panel side edge were taken for visual inspection after performing all the tests mentioned above. These test and inspection results validated the effectiveness of a highly damped novel deployable solar panel module for use in actual space missions.

2. The STEP Cube Lab-II CubeSat’s Overview 2.1. Mission Objectives and System Descriptions Figure1 illustrates the system architecture of the STEP Cube Lab-II CubeSat mission. The STEP Cube Lab-II CubeSat has been developed in a collaborative framework of university and start-up aerospace companies for educational and technological veriﬁca- tion purposes. The STSL of Chosun University and a consortium team of eight domestic organizations are involved in this project to verify domestically developed space technologies and demonstrate unique nighttime video mode remote sensing capability through CubeSat. The CubeSat is equipped with three commercial off-the-shelf (COTS) remote sensing instruments onboard as the primary payload for earth remote sensing, such as an Aerospace 2021, 8, x FOR PEER REVIEW 4 of 24

Aerospace 2021, 8, 64 4 of 24 CubeSat. The CubeSat is equipped with three commercial off-the-shelf (COTS) remote sensing instruments onboard as the primary payload for earth remote sensing, such as an electro-opticalelectro-optical camera (EOC),(EOC), broadbandbroadband infrared infrared camera camera (BBIRC), (BBIRC), and and long-wave long-wave infrared infra- redcamera camera (LWIRC). (LWIRC). The primaryThe primary objective objective of the of STEP the STEP Cube Lab-IICube Lab-II project project is to carry is to out carry the outnation’s the nation’s first multi-band first multi-band earth observation earth observ missionation mission by utilizing by utilizing CubeSat’s CubeSat’s electro-optical electro- optical(EO) and (EO) infrared and infrared (IR) images (IR) images and videos and videos on various on various targets target of thes of Koreanthe Korean peninsula, peninsula,including including Mt. Paektu, Mt. Paektu, which which has recently has recently shown shown signs signs of a volcanic of a volcanic eruption, eruption, observation obser- vationof areas of affectedareas affected by forest by forest fires, andfires, the and route the route of spread of spread in areas in areas where where forest forest fires havefires haveoccurred, occurred, analysis analysis of the of thermal the thermal island island phenomenon phenomenon in urban in urban areas areas of metropolitan of metropolitan cities, cities,observation observation of ship of activities,ship activities, and monitoringand monitoring of the of nuclear the nuclear power power plant’s plant’s coolant coolant dis- dischargecharge in seawater.in seawater. Furthermore, Furthermore, one ofone the of space the space technologies technologies to be verified to be verified in this mission in this missionis an optimized is an optimized data compression data compression method me in thethod payload in the payload data handling data handling system (PDHS)system (PDHS)and payload and datapayload transmission data transmission system (PDTS) system for (PDTS) video data for processingvideo data and processing transmission and transmissionto the ground to station. the ground Additionally, station. Addition this missionally, this also mission tests a high-dampingalso tests a high-damping VMLSA to VMLSAensure structural to ensure protectionstructural ofprotection mounted of solar mounted cells in solar launch cells vibration in launch environments vibration envi- by ronmentssuppressing by transmittedsuppressing launch transmitted loads onlaunch the solar loads panel on the and solar a three-pogo panel and pin-based a three-pogo burn pin-basedwire triggering burn mechanismwire triggeri forng high mechanism holding capability for high atholdin the launchg capability stowed at configuration the launch stowedof the solar configuration panel and reliableof the solar release panel action and in reliable space. release Furthermore, action anin optimizedspace. Furthermore, SADA for anthe optimized acquisition SADA of solar for powerthe acquisition maximization of sola isr onepower of themaximization secondary payloads.is one of the This second- study aryis mainly payloads. focused This onstudy the is on-ground mainly focused experimental on the on-ground validation experimental of the design validation effectiveness of theand design structural effectiveness safety of theand VMLSA structural combined safety of with the VMLSA P-HRM. combined with P-HRM.

Figure 1. The system architecture of the Space Technology ExperimentalExperimental Project CubeSat Laboratory (STEP Cube Lab)-II CubeSat’s mission.

TableTable 11 presentspresents thethe detaileddetailed speciﬁcationsspecifications of of the the CubeSat CubeSat system. system. For For mission mission opera- oper- ation,tion, thethe satellitesatellite orbitorbit isis selectedselected asas aa sun-synchronoussun-synchronous orbitorbit of 700 km altitude from the earthearth surfacesurface andand an an inclination inclination angle angle of 98.2of 98.2°◦ with with a local a timelocal oftime descending of descending node (LTDN) node (LTDN)of 16:00 of p.m. 16:00 The p.m. ground The ground station contactstation contact time of time the satellite of the satellite calculated calculated from the from orbital the orbitalanalysis analysis is approximately is approximately 7.5 min, 7.5 and min, access and access to the to ground the ground station station occurs occurs six times six times a day. aThe day. CubeSat The CubeSat mission mission operation operation life is setlife to is oneset to year. one year.

Aerospace 2021, 8, 64 5 of 24

Table 1. Detailed system speciﬁcations of the STEP Cube Lab-II.

Items Speciﬁcations Satellite name STEP Cube Lab-II Size (mm) 366 (X) × 117.8 (Y) × 238.5 (Z) Weight (kg) 9.6 Orbit: Sun-synchronous orbit Altitude: 700 km Orbital parameter Orbital inclination angle: 98.2 deg. Local time of descending node (LTDN): 16:00 p.m Primary payloads: EOC, BBIRC, LWIRC Mission payload Secondary payloads: PDHS and PDTS, VMLSA, P-HRM, SADA Camera EOC BBIRC LWIRC GSD (m) 10 350 350 Camera performance Observation width (km) 40 × 40 220 × 160 220 × 160 Wave length range (µm) 0.38–0.94 3–14 9–12 Multi-band still imaging mode: 5.85 Observation mission time (s) BBIRC and LWIRC video mode: 10 Ground station contact 7.5 Minutes (Average) Uplink UHF band: 1200 bps Audio frequency shift keying (AFSK) Communication S-band: >8 Mbps UHF band: 9600 bps Gaussian minimum system Downlink 16-ary amplitude and phase shift keying shift keying (GMSK) (16-APSK) Attitude control system 3-axis Attitude control method Mission life 1 Year

2.2. Satellite’s Configuration and Design Descriptions Figure2a,b demonstrate the STEP Cube Lab-II mechanical configurations at the solar panel stowed and deployed states, respectively. The CubeSat’s mechanical dimensions and total weight are 366 mm (X-axis) × 117.8 mm (Y-axis) × 238.5 mm (Z-axis) and 9.6 kg, respectively, which are within the 6U standard. The CubeSat mainly consists of COTS components enclosed in domestically developed structures, while a few interface boards and mechanical subsystems were designed and assembled by the team. The primary payload comprises three optics and sensor pairs: EOC, BBIRC, and LWIRC, which are designed to operate in staring mode and are capable of capturing images and videos nearly simultaneously from the three cameras. The performance specification of the EOC of spectral 0.38–0.94 µm wavelength range is a 10 m ground sampling distance (GSD), which covers an area of 40 km × 40 km swath width. The GSD and swath width of the BBIRC and LWIRC were 350 m and 220 km × 160 km, respectively. The maximum estimated observation times for the multi-band still imaging mode and dual-IRC video mode on the Korean peninsula during an orbit period are 5.85 s and 10 s, respectively. All commands for uplink as well as the downlink of images and telemetry will be performed using the ultra-high frequency (UHF) amateur band at a frequency of 1200–9600 bps. Additionally, the commu- nications between CubeSat and ground stations for data transmission are accomplished by an S-band antenna with a capacity greater than 8 Mbps. For the redundancy of the satellite communication system, communication between the satellite and ground stations can be accomplished through the iridium communication network, which also enables satellite orbital determination through the iridium GPS navigation data. Furthermore, the satellite’s attitude control is accomplished by the three-axis attitude control method by the application of the XACT-15 attitude determination and control system (ADCS) [20]. To optimize the in-orbit power generation capability from the fixed-type deployable solar panels, the attitude of STEP Cube Lab-II will be set in such a way that deployable solar panels always point toward the sun direction by a 90-degree roll maneuver of the CubeSat from the nadir-pointing attitude through the ADCS, except for the periods of mission operation Aerospace 2021, 8, x FOR PEER REVIEW 6 of 24

Aerospace 2021, 8, 64 (ADCS) [20]. To optimize the in-orbit power generation capability from the fixed-type de-6 of 24 ployable solar panels, the attitude of STEP Cube Lab-II will be set in such a way that deployable solar panels always point toward the sun direction by a 90-degree roll maneuver of the CubeSat from the nadir-pointing attitude through the ADCS, except for the periods and ground contact. For mission accomplishment, the onboard ADCS roll maneuvers the of mission operation and ground contact. For mission accomplishment, the onboard satellite back to the nadir point attitude. ADCS roll maneuvers the satellite back to the nadir point attitude.

(a) (b)

Figure 2. STEP Cube Lab-II’s mechanical configurations: conﬁgurations: ( a) solar panel stowed and (b) solar panel deployed.

2.3. Design2.3. Design of Solar of SolarPanel PanelModule Module with the with Accommodation the Accommodation of Solar of SolarCell Cell A deployableA deployable solar solarpanel panel based based on a onPCB a PCBsubstrate substrate using using multilayered multilayered thin stiffeners thin stiffeners attachedattached using using double-sided double-sided viscoelastic viscoelastic tapes is tapes proposed is proposed for STEP for Cube STEP Lab-II Cube CubeSat Lab-II CubeSat as a novelas adesign novel designapproach approach to ensure to ensurestructural structural protection protection of solar of cells solar by cells suppressing by suppressing the the transmittedtransmitted vibration vibration loads loads on a panel on a panel in a la inunch a launch environment. environment. The holding The holding and releasing and releasing actionaction of the of solar the solarpanel panel is accomplished is accomplished using using an optimized an optimized version version of the of pogo the pogopin-based pin-based burn burnwire cutting wire cutting release release mechanism. mechanism. The solar The panel solar deployment panel deployment is set to isoccur set towithin occur 1 withinh of the1 orbital h of the injection orbital of injection STEP Cube of STEP Lab-II Cube from Lab-II the P-POD. from theThe P-POD. passive Thetorsional passive force torsional of the torsionalforce of thehinges torsional deploys hinges and deployslatches the and pa latchesnel at its the projected panel at itsposition. projected Twenty position. AZUR Twenty spaceAZUR 30% efficiency space 30% triple-junct efﬁciencyion triple-junction GaAs solar cells GaAs (3G30C) solar [9 cells] can (3G30C)be mounted [9] canon the be front mounted surfaceon of the the front proposed surface solar of thepanel proposed because this solar design panel strategy because does this not design reduce strategy the availa- does not ble areareduce for solar the available cell accommodation. area for solar cellThus, accommodation. the power generation Thus, the of powerthe proposed generation solar of the panelproposed is maximized solar compared panel is maximized to the same compared solar panel to the area same by solaremploying panel areathe aforemen- by employing tionedthe design aforementioned strategies to design ensure strategies the structural to ensure safety the of structuralsolar cells. safety of solar cells. The energyThe energy balance balance analysis analysis (EBA) (EBA) of the of CubeSat the CubeSat is performed is performed to verify to verify whether whether the the power generation from the proposed solar cell composition is well adjusted with power power generation from the proposed solar cell composition is well adjusted with power consumption by the subsystems and payloads in mission scenarios, power storage for consumption by the subsystems and payloads in mission scenarios, power storage for system operation in eclipse, and battery health over the mission life [21]. Table2 shows system operation in eclipse, and battery health over the mission life [21]. Table 2 shows the results of the EBA in the multi-band still imaging and IR video operation modes of the results of the EBA in the multi-band still imaging and IR video operation modes of earth observation. The earth observation period of the multi-band still imaging and IR earth observation. The earth observation period of the multi-band still imaging and IR video operation mode of the CubeSat is the most power-consuming phase relative to the video operation mode of the CubeSat is the most power-consuming phase relative to the other operating modes that were determined by the power on/off status of the hardware other operating modes that were determined by the power on/off status of the hardware in operational modes. The total power consumption by the CubeSat subsystems and in operational modes. The total power consumption by the CubeSat subsystems and pay- payloads in multi-band still imaging and IR video operation mode with a contingency loads in multi-band still imaging and IR video operation mode with a contingency margin margin of 10 percent is 18.38 W, which was calculated using the operating voltage and of 10 percent is 18.38 W, which was calculated using the operating voltage and current current information provided in each component’s datasheet. CubeSat’s power analysis was performed using the FreeFlyer software in accordance with the orbital information mentioned in Table1. The AZUR space 30% efﬁciency triple-junction GaAs solar cell

(3G30C) was considered in the analysis. The analysis result shows that the average power Aerospace 2021, 8, 64 7 of 24

generation in an orbit in the worst-case scenario would be 33.29 W. The EBA result indicates that the power margin of the satellite in an orbit is 2.25 Wh, which is 5.63% of the generated power in a daylight period. Furthermore, the maximum depth of discharge (DoD) of the battery in an orbit with a battery pack capacity of 77 Wh is 13.41%, which is within the maximum 20% system requirement. The results show that the CubeSat will have a sufﬁcient power margin and good battery health over the mission life period in the proposed composition of solar cells with the satellite’s sun-pointing attitude. The system requirement of battery DoD% less than 20 cannot be guaranteed with only body-mounted solar cells, which justiﬁes the choice of deployable solar panels for STEP Cube Lab-II to accomplish mission objectives over the 1-year mission operation lifetime.

Table 2. Energy balance analysis in mission operation mode of Earth observation.

Power Charge/ Time (min) Parameter Power (W) Remarks (Wh) Discharge (Wh) Generation 33.29 39.95 Daylight 72 Consumption −18.38 −22.06 Average power for charging 17.89 Actual power for charging (Considering conversion eff.) 11.44 Eclipse 27 Consumption −18.38 −9.19 −9.19 Power margin 2.25 Power budget (%) 5.63 Req.: >0 Battery depth of discharge (DoD) (%) 13.41 Req.: <20

Table3 summarizes the detailed mass budget of the solar panel module proposed in this study. The total mass of the proposed solar panel module is 306.5 g, which is comparatively lighter than that of the commercially available deployable solar panel for 6U CubeSat and those developed in an academic environment. For instance, the mass of the proposed solar panel module is lower by a factor of 2.04, compared to the 6U CubeSat’s solar panel module stiffened by applying additional high-pressure laminated G10 material [8].

Table 3. Detailed mass budget of the proposed solar panel model.

Items Mass (g) Printed circuit board (PCB) panel 201.5 Thin stiffeners 85 Viscoelastic tapes 3 Pogo pin-based HRM 6 Hinges and fasteners 11 Total 306.5

3. A PCB-Based High-Damping Deployable Solar Panel Module 3.1. Design Description Figure3a,b show the demonstration model of a highly damped deployable solar panel module’s stowed and deployed conﬁgurations, respectively. The solar panel module is comprised of a PCB panel with thin stiffeners attached by adhesive tapes, a three-pogo pin-based HRM, and torsional hinges. The PCB panel of dimensions 325.4 mm × 193 mm × 1.6 mm is made up of FR4 material and provides a mechanical interface for mounting solar cells and stiffeners. Five layers of thin 0.4 mm FR4 PCB stiffeners, as shown in Figure3, Aerospace 20212021,, 88,, 64x FOR PEER REVIEW 88 of of 24

Figurewere mounted 3, were onmounted the rear on surface the rear of thesurface PCB of panel the PCB through panel double-sided through double-sided 3MTM 966 acrylic 3MTM 966tapes acrylic [22]. Thetapes 3M [22].TM 966The acrylic 3MTM tape966 acrylic is a high-temperature tape is a high-temperature adhesive that adhesive has been that used has in beenaerospace used applicationsin aerospace that applications comply with that ASTM comply E596 with low ASTM volatility E596 speciﬁcation low volatility criteria specifi- for cationspace applicationcriteria for space of the application National Aeronautics of the National and Aeronautics Space Administration and Space (NASA). Administration For the (NASA).attenuation For of the transmitted attenuation launch of transmitted loads on the la panel,unch theloads fundamental on the panel, principle the fundamental is based on principleconstrained is based layer dampingon constrained accomplished layer damping by the shear accomplished deformation by characteristics the shear deformation achieved characteristicsbetween interlaminated achieved thinbetween stiffeners interlaminated and adhesive thin tapes. stiffeners The basicand adhesive speciﬁcations tapes. of The the basicmaterials specifications used in the ofsolar the materials panel module used in are the listed solar in panel Table 4module. are listed in Table 4.

(a) (b)

Figure 3. ConfigurationsConﬁgurations of of a ahighly highly damped damped deployable deployable solar solar panel panel module’s module’s demonstration demonstration model: model: (a) (stoweda) stowed and and (b) (deployed.b) deployed.

TableThe 4. Speciﬁcations stiffener attachme of materialsnt process used in theon solarthe panel panel is module. considerably simple, although the uniform bonding strength distribution in the workmanship should be controlled with in- Item Details Value tensive care. Therefore, for symmetrical attachment and easy integration of the stiffeners, four guide holes were madeElastic on modulus the four (Pa)edge corners of the PCB panel18.73 and× stiffeners.109 The stiffener attachment processDensity on the (kg/m PCB panel3) is shown in Figure 4. First,1850 prepare the ma- FR4 terials and integration tools such as PCB panel, stiffeners, 3M966 tape, integration jig, iso- Poisson’s ratio 0.136 propyl alcohol (IPA), torque wrench, knife, and cleaning cloth, then wipe the stiffeners and PCB panel byThermal using a cloth conductivity moistened (W/mK) with IPA to remove dust contamination. 0.29 The tape was cut according to theManufacturer shape of the stiffener by a knife tip, put 3M into Company the integration jig through the guide hole interfaces, and the process was repeated up to the specified Adhesive material Acrylic number of stiffeners that had to be attached to the panel. The PCB panel should be placed on the last stiffener attachedAdhesive to thicknessthe tape, and (mm) then the top guide jig has to 0.06 clamp with M3 bolts fastened on the jig throughColor a torque of 1.1 Nm to create a compression Transparent force on the specimen. As recommended by the adhesive tape datasheet [22], the adhesive bonding Adhesive tape [22] Allowable temperature range (◦C) −40 to 232 resin of the attached tapes must be cured for up to 72 h for secure attachment. Thermal conductivity at 41 ◦C (W/mK) 0.178 Coefﬁcient of thermal expansion (ppm/◦C) 1.99 Adhesive strength to steel (N/100 mm) 159 Total mass loss (TML) and Collected volatile 0.93, 0.01 condensable material (CVCM) outgassing (%)

The stiffener attachment process on the panel is considerably simple, although the uniform bonding strength distribution in the workmanship should be controlled with intensive care. Therefore, for symmetrical attachment and easy integration of the stiffeners,

Aerospace 2021, 8, 64 9 of 24

four guide holes were made on the four edge corners of the PCB panel and stiffeners. The stiffener attachment process on the PCB panel is shown in Figure4. First, prepare the materials and integration tools such as PCB panel, stiffeners, 3M966 tape, integration jig, isopropyl alcohol (IPA), torque wrench, knife, and cleaning cloth, then wipe the stiffeners and PCB panel by using a cloth moistened with IPA to remove dust contamination. The tape was cut according to the shape of the stiffener by a knife tip, put into the integration jig through the guide hole interfaces, and the process was repeated up to the speciﬁed number of stiffeners that had to be attached to the panel. The PCB panel should be placed on the last stiffener attached to the tape, and then the top guide jig has to clamp with M3 bolts fastened on the jig through a torque of 1.1 Nm to create a compression force on the Aerospace 2021, 8, x FOR PEER REVIEW 9 of 24 specimen. As recommended by the adhesive tape datasheet [22], the adhesive bonding resin of the attached tapes must be cured for up to 72 h for secure attachment.

FigureFigure 4. 4.Integration Integration processprocess of a stiffeners stiffeners on on the the PCB PCB panel. panel.

TheThe numbernumber of attached attached stiffeners stiffeners was was dete determinedrmined by the by 3U the CubeSat 3U CubeSat solar panel’s solar panel’s ex- experimentalperimental test test results results [18] [18 while] while consider consideringing dynamic dynamic clearance clearance while while accommodating accommodating the theP-POD. P-POD. The The thickness thickness of the of the solar solar panel panel after after the the attachment attachment of the of thestiffeners stiffeners is 3.7 is 3.7mm, mm, whichwhich provides provides aa laterallateral edge ga gapp margin margin of of 3.3 3.3 mm mm for for a adynamic dynamic clearance clearance on onP-POD P-POD [11]. [11 ].

3.2.Table Solar 4. Specifications Panel Holding of andmaterials Release used Mechanism in the solar panel module. FigureItem5a,b illustrate close-up views of theDetails proposed three-pogo pin-basedValue mechanism in fully and partially stowed solar panelElastic states, modulus respectively. (Pa) The mechanism18.73 × comprises 109 pogo pins, electrical interface PCB, brackets,Density Dyneema (kg/m3) wire, resistor, and resistor1850 PCB. Park FR4 et al. [23] and Bhattarai et al. [24] investigatedPoisson’s the ratio applicability and effectiveness0.136 of two pogo pins in HRM, although the mainThermal purpose conductivity of using (W/mK) three pogo pins in the0.29 proposed mechanism is to simplify the electrical circuitManufacturer by reducing the electrical3M components Company to lower the cost of the product and increaseAdhesive its reliability material for secure release actionAcrylic in a harsh space environment. The main designAdhesive driver thickness for the (mm) use of the pogo pin (MP511-1111-0.06 E03100A, CFE Corporation Co., Guangdong,Color China) [25] connectors is theTransparent establishment of Allowable temperature range (°C) −40 to 232 a secureAdhesive electromechanical tape [22] connection between the electrical interface PCB and the resistor PCB integrated at the edge of theThermal CubeSat conductivity structure at 41 and °C solar (W/mK) panel edge, respectively.0.178 The pogo pins provide electricalCoefficient current to of thethermal resistor expansion mounted (ppm/°C) on the resistor1.99 PCB during Adhesive strength to steel (N/100 mm) 159 mechanism activation. The holding constraint on the solar panel at the stowed state is Total mass loss (TML) and Collected volatile conden- achieved by tightening the Dyneema wire on the brackets. Mechanical restraint0.93, 0.01 in the sable material (CVCM) outgassing (%) in-plane direction of the solar panel during the stowed conﬁguration is accomplished by ball3.2. andSolar socket Panel Holding joints made and Release on the Mechanism brackets that act as a mechanical limiter to avoid the adverse panel strike on pogo pins beyond the pogo pin’s plunger working stroke range. Figure 5a,b illustrate close-up views of the proposed three-pogo pin-based mecha- Additionally, the socket limits ball movement within a nominal gap, which averts the nism in fully and partially stowed solar panel states, respectively. The mechanism com- burden at the wire knot under vibration environments. To release the holding constraint on prises pogo pins, electrical interface PCB, brackets, Dyneema wire, resistor, and resistor the solar panel, a resistor (3216 SMD type, Walsin Technology Co., Taipei, Taiwan) [26] was PCB. Park et al. [23] and Bhattarai et al. [24] investigated the applicability and effectiveness of two pogo pins in HRM, although the main purpose of using three pogo pins in the proposed mechanism is to simplify the electrical circuit by reducing the electrical components to lower the cost of the product and increase its reliability for secure release action in a harsh space environment. The main design driver for the use of the pogo pin (MP511- 1111-E03100A, CFE Corporation Co., Guangdong, China) [25] connectors is the establishment of a secure electromechanical connection between the electrical interface PCB and the resistor PCB integrated at the edge of the CubeSat structure and solar panel edge, respectively. The pogo pins provide electrical current to the resistor mounted on the resistor PCB during mechanism activation. The holding constraint on the solar panel at the

Aerospace 2021, 8, x FOR PEER REVIEW 10 of 24

stowed state is achieved by tightening the Dyneema wire on the brackets. Mechanical restraint in the in-plane direction of the solar panel during the stowed configuration is accomplished by ball and socket joints made on the brackets that act as a mechanical limiter Aerospace 2021, 8, 64 to avoid the adverse panel strike on pogo pins beyond the pogo pin’s plunger working10 of 24 stroke range. Additionally, the socket limits ball movement within a nominal gap, which averts the burden at the wire knot under vibration environments. To release the holding constraint on the solar panel, a resistor (3216 SMD type, Walsin Technology Co., Taipei, integratedTaiwan) [26] on was the resistorintegrated PCB on as the an resistor actuator. PCB As as the an mechanismactuator. As isthe activated mechanism by the is acti- power supplyvated by in the an electricalpower supply circuit, in thean electrical resistor cutscircuit, the the tightened resistor wire.cuts the The tightened panel deployment wire. The is initiatedpanel deployment instantly by is initiated the compression instantly forceby the of compression the pogo pin force plunger, of the 0.68pogo± pin0.19 plunger, N, which could0.68 ± rapidly 0.19 N, which or simply could interrupt rapidly or the simply circuit interrupt path to the circuit resistor. path Consequently, to the resistor. the Con- hinge torsionalsequently, forces the hinge deploy torsional the solar forces panel deploy to its th intendede solar panel angle. to Tableits intended5 lists the angle. speciﬁcations Table 5 oflists the the hardware specifications used inof P-HRM.the hardware used in P-HRM.

(a) (b)

FigureFigure 5. 5.Conﬁgurations Configurations of of three-pogo three-pogo pin-based pin-based HRM:HRM: (a) solar panel panel fully fully stowed stowed and and (b ()b solar) solar panel panel partially partially stowed. stowed.

Table 5.TableSpeciﬁcations 5. Specifications of the hardwareof the hardware used in used the in three-pogo the three-pogo pin-based pin-based HRM. HRM.

ItemsItems DetailsDetails Value Value Manufacture CFE Corporation Co. ManufactureVoltage and current (V, A) CFE Corporation 12, 3 Co. Voltage andMax. current electrical (V, contact A) resistance (mΩ) 12, 350 Life test (Cycle) 100,000 Max. electrical contact resistance (mΩ) 50 Pogo pin Qualification temperature range (°C) −40 to 85 Life test (Cycle)Spring force (N) 100,000 0.68 ± 0.19 Pogo pin Qualiﬁcation temperatureTotal range length (◦ (mm)C) −40 to 9 85 Full stroke (mm) 1.4 ± Spring forceWorking (N) stroke (mm) 0.68 1.00.19 Total length (mm)Manufacture 9YGK Material Dyneema Wire Full stroke (mm) 1.4 Diameter (mm) 0.205 Working strokeMax. (mm) allowable force (N) 1.0 88.2 ManufactureManufacture Walsin technology YGK Co., Taipei, Taiwan Package SMD Type Material Dyneema Wire Resistor Electrical resistance (ohm) 4.7 Diameter (mm)Resistance tolerance (%) 0.205 ± 1 Max. allowableMax. force power (N) dissipation (W) 88.2 0.25

Figure 6a,bManufacture illustrate the electrical system of Walsin the three-pogo technology pin-based Co., Taipei, mechanisms. Taiwan For mechanism Packageactivation, the input voltage on the electrical circuit SMD Typepath to the resistor Resistor Electrical resistance (ohm) 4.7

Resistance tolerance (%) ±1 Max. power dissipation (W) 0.25

Figure6a,b illustrate the electrical system of the three-pogo pin-based mechanisms. For mechanism activation, the input voltage on the electrical circuit path to the resistor was applied to 8V, although the electrical components were supplied at 3.3 V. The optocoupler Aerospace 2021, 8, x FOR PEER REVIEW 11 of 24

Aerospace 2021, 8, 64 11 of 24

was applied to 8V, although the electrical components were supplied at 3.3 V. The optocoupler(PC817) (PC817) [27] transmits [27] transmits electrical electrical signals signals by light by between light between two electrically two electrically isolated isolated circuits, circuits,thus preventing thus preventing high-voltage high-voltage electrical electrical malfunctions malfunctions in the in circuit.the circuit. The The power power cutoff cut- offfunction function on on the the mechanism mechanism is accomplished is accomplished instantly instantly after after the deployment the deployment of the of solar the solar panel panelwith thewith electrical the electrical circuit circuit shown shown in the blockin the diagram block diagram without without employing employing a microcontroller a micro- controllerunit (MCU). unit (MCU).

(a)

(b)

FigureFigure 6. 6. Three-pogoThree-pogo pin-based mechanism’s electricalelectrical system: system: ( a()a front) front and and rear rear view view of electricalof electrical interface interface PCB PCB (dimensions: (dimensions:34 mm 34× mm16 mm× 16 ×mm1 mm)× 1 mm) (b) schematic(b) schematic block block diagram. diagram.

FigureFigure 7a,b7a,b illustrate illustrate resistor resistor PCB’s PCB’s front front and and rear rear views, views, respectively. respectively. A Aresistor resistor as anas actuator an actuator for solar for solar panel panel release release is soldered is soldered on the on front the side front of sidethe resistor of the resistorPCB, which PCB, faceswhich the faces outward the outward direction direction while integrating while integrating on the solar on thepanel. solar To panel. interconnect To interconnect the elec- tricalthe electrical power coming power from coming the pogo from pins the pogo to the pins resistor, to the three resistor, surface-mounted three surface-mounted electrodes wereelectrodes attached were to the attached rear surface to the of rear the surface resistor of PCB the that resistor was PCBmechanically that was connected mechanically to connected to the resistor internally. Additionally, two via hole interfaces were made near the resistor internally. Additionally, two via hole interfaces were made near the electrodes the electrodes that were physically connected to the electrodes. In general, thermal vias are that were physically connected to the electrodes. In general, thermal vias are well-known well-known methods used to enhance the heat dissipation of surface-mounted components methods used to enhance the heat dissipation of surface-mounted components in a PCB. in a PCB. However, the application of vias in the resistor PCB is to make electrodes more However, the application of vias in the resistor PCB is to make electrodes more securely securely attached to the PCB surface in order to avoid the risk of electrode pad detachment attached to the PCB surface in order to avoid the risk of electrode pad detachment due to due to pogo pin friction under launch vibration loads. Bhattarai et al. [24] performed a

Aerospace 2021, 8, x FOR PEER REVIEW 12 of 24

Aerospace 2021, 8, 64 12 of 24

pogo pin friction under launch vibration loads. Bhattarai et al. [24] performed a scanning electronscanning microscope electron microscope (SEM) inspection (SEM) inspection of the electrode of the pads electrode after padscompleting after completing all the vibra- all tionthe vibrationtests to evaluate tests to evaluatethe frictional the frictional impact or impact damage or damage caused causedby the bytip theof the tip ofpogo the pins, pogo evenpins, though even though the spring-loaded the spring-loaded pin has pin a low has friction a low friction impact impact on vibration on vibration loads. loads.The thick- The nessthickness of the of electrical the electrical contact contact part partor curvature or curvature tip on tip the on electrodes the electrodes was wasonly only reduced reduced by 23.88%by 23.88% compared compared to that to that of ofthe the 60.37 60.37 μmµm initial initial thickness. thickness. The The frictional frictional impact impact on on the the electrodeelectrode under under vibration vibration load load is is minimal; minimal; thus, thus, a a secure secure physical physical contact contact can can be be assured assured eveneven after after being being exposed exposed to to severe severe launch launch vibr vibrationation loads loads because because of the the 1.7 1.7 mm mm full full stroke stroke ofof the the spring-loaded spring-loaded pin pin of of the the pogo pogo pins. pins.

(a) (b)

FigureFigure 7. 7. MechanicalMechanical configuration configuration of of the the burn burn resistor resistor PCB: PCB: ( (aa)) front front view view and and ( (bb)) rear rear view. view. In the mechanism, pogo pins with the combination of voltage resistor divider are In the mechanism, pogo pins with the combination of voltage resistor divider are used to determine the panel deployment status, which provides telemetry of “1” or “0” used to determine the panel deployment status, which provides telemetry of “1” or “0” through a buffer IC based on the current flow status in the resistor’s circuit path, where “1” through a buffer IC based on the current flow status in the resistor’s circuit path, where implies voltage resistor divider’s output voltage. Table6 lists the truth tables of the circuit. “1” implies voltage resistor divider’s output voltage. Table 6 lists the truth tables of the Once the mechanism is activated, the current flows through the resistor and the voltage circuit. Once the mechanism is activated, the current flows through the resistor and the resistor divider. The output voltage at the voltage resistor divider becomes approximately voltage resistor divider. The output voltage at the voltage resistor divider becomes ap- 0 V until the panel is deployed because the resistance value of the resistors used in the proximatelyvoltage resistor 0 V divideruntil the is panel significantly is deployed higher be thancause that the of resistance the resistor value used of as the an resistors actuator usedfor panel in the deployment. voltage resistor The divider electrical is circuitsignifican overtly the higher resistor than becomes that of the open resistor once the used solar as anpanel actuator is deployed for panel such deployment. that the output The signal electric ofal the circuit circuit over becomes the resistor high, whichbecomes confirms open oncethe deployment the solar panel status is deployed of the solar such panel. that Therefore, the output compared signal of tothe the circuit mechanism becomes [28 high,], the whichP-HRM confirms proposed the in deployment this study has status a relatively of the simplersolar panel. electrical Therefore, circuit andcompared is well to suited the mechanismto the available [28], P-PODthe P-HRM lateral proposed edge gap. in this study has a relatively simpler electrical circuit and is well suited to the available P-POD lateral edge gap. Table 6. Circuit truth table for deployment status of solar panel. Table 6. Circuit truth table for deployment status of solar panel. Solar Panel Deployment Status Input Solar Panel Deployment Status Input Undeployed UndeployedDeployed Deployed Vin Enable Pogo pin voltage Pogo Output pin voltageOutput Pogo pin voltagePogo pin OutputOutput voltage Vin Enable 1 1 1 voltage 0 voltage 1 voltage voltage 1 1 10 01 1 0 0 0 1 0 1 1 0 0 0 0 0 3.3. Wire-Tightening Process 3.3. Wire-TighteningThe wire-tightening Process procedure on the brackets for the solar panel’s holding mechanical constraint is shown in Figure 8. Dyneema wire (YGK G-Soul) [29] (diameter 0.20 mm) The wire-tightening procedure on the brackets for the solar panel’s holding mechanical is wound on the brackets that can bear maximum strength of 88.2 N; the panel-holding constraint is shown in Figure8. Dyneema wire (YGK G-Soul) [ 29] (diameter 0.20 mm) capacity can be further increased with the wire-winding numbers. The surgeon’s knot is is wound on the brackets that can bear maximum strength of 88.2 N; the panel-holding carried out at the bracket corner for the final knotting of the wire, which helps to create a capacity can be further increased with the wire-winding numbers. The surgeon’s knot is steadycarried tight out attension the bracket on the corner wire knot. for the For final reliable knotting panel of release, the wire, physical which helpscontact to of create the

Aerospace 2021, 8, 64 13 of 24

Aerospace 2021, 8, x FOR PEER REVIEW 13 of 24

a steady tight tension on the wire knot. For reliable panel release, physical contact of the wire to the resistor is mandatory. Thus, the notching guide rail is made on the bracket’s surface to prevent wire misalignment from the resistor surface under launch vibrationvibration loads.loads. Additionally, thethe combination combination of of vertical vertical and and horizontal horizontal cross-pattern cross-pattern winding winding of the of thewire wire with with two two additional additional hole hole interfaces interfaces in the in brackets the brackets creates creates a secure a secure holding holding constraint con- strainton the panel.on the The panel. process The ofproces wires tightening of wire tightening to stow the to solar stow panel the solar is therefore panel signiﬁcantlyis therefore significantlysimpler and moresimpler reliable and more than reliable conventional than conventional mechanisms. mechanisms.

Figure 8. Solar panel’s wire-tightening process.

4. Dynamic Dynamic Characteristics Characteristics Investigation InIn the constrainedconstrained layerlayer damping damping strategy, strategy, the the transmitted transmitted vibration vibration energy energy dissipation dissipa- tionis highly is highly dependent dependent on the on viscoelastic the viscoelastic core layercore layer thickness, thickness, dynamic dynamic modulus, modulus, thickness thick- of nessconstraining of constraining layers, andlayers, vibration and vibration frequency frequency [30,31]. Therefore,[30,31]. Therefore, a solar panel a solar free-vibration panel free- vibrationtest was executedtest was executed at 25 ◦C ambientat 25 °C ambien room temperaturet room temperature under the under boundary the boundary condition con- in ditionwhich in the which interfaces the interfaces of the hinge of andthe hinge HRM and holes HRM were holes rigidly were clamped. rigidly The clamped. roving The hammer rov- ingmethod hammer was method used to was excite used the to solar excite panel the solar in its panel free vibration. in its free Anvibration. accelerometer An accelerom- sensor eterwas sensor mounted was at mounted the solar panelat thecenter solar panel to obtain center the time-domainto obtain the frequencytime-domain responses. frequency To responses.investigate To the investigate design usefulness the design or practicality usefulness for or rapidpracticality suppression for rapid of panel suppression chattering of panelvibration chattering during vibration its in-orbit during performance, its in-orbit the performance, solar panel the module’s solar panel acceleration module’s proﬁle accel- erationunder deploymentprofile under was deployment also measured was also in the me sameasured room in the temperature same room environment. temperature Theen- vironment.dynamic characteristics The dynamic investigation characteristics tests invest of a typicaligation PCBtestspanel of a typical of the samePCB sizepanel without of the sameattaching size additionalwithout attaching stiffeners additional were also stiffe performedners were to comparealso performed the results to compare obtained the from re- sultsthe VMLSA. obtained Figure from9 thea,b showVMLSA. the Figure solar panel 9a,b timeshow histories the solar in panel the free-vibration time histories tests in the at free-vibrationthe rigidly mounted tests at condition the rigidly and mounted the acceleration condition proﬁle and under the acceleration deployment, profile respectively. under deployment,The results illustrate respectively. that stiffenersThe results with illustrate viscoelastic that stiffeners acrylic tape with can viscoelastic effectively acrylic suppress tape canthe transmittedeffectively suppress vibration the on transmitted the solar panel. vibration on the solar panel.

Aerospace 8 Aerospace 20212021,, 8,, 64x FOR PEER REVIEW 1414 of of 24

(a) (b)

Figure 9.9. Dynamic characteristics investigation results of the solar panels: (a) timetime historieshistories ofof free-vibrationfree-vibration responseresponse inin rigidly mountedmounted conditioncondition andand ((bb)) accelerationacceleration proﬁleprofile underunder deployment.deployment.

Table7 7 summarizes summarizes the the solarsolar panel’spanel’s free-vibrationfree-vibration test test results results in in the the rigidlyrigidly mounted mounted condition. The results show that the VMLSA’s ﬁrst eigenfrequency and damping ratio condition. The results show that the VMLSA’s first eigenfrequency and damping ratio were 110.1 Hz and 0.141, higher by a factor of 1.33 and 3.9, respectively, compared to were 110.1 Hz and 0.141, higher by a factor of 1.33 and 3.9, respectively, compared to those those of the typical PCB solar panel. The results show that the application of viscoelastic of the typical PCB solar panel. The results show that the application of viscoelastic multi- multilayered thin stiffeners signiﬁcantly improved the vibration attenuation along with layered thin stiffeners significantly improved the vibration attenuation along with in- increased solar panel rigidity due to shear deformation characteristics and tough surface creased solar panel rigidity due to shear deformation characteristics and tough surface roughness between the interlaminated layers. Furthermore, this design strategy could also roughness between the interlaminated layers. Furthermore, this design strategy could also be advantageous for rapid attenuation of in-orbit vibration on the panel caused by rigid be advantageous for rapid attenuation of in-orbit vibration on the panel caused by rigid body motion of the satellite during slew maneuver as the panel chattering vibration under body motion of the satellite during slew maneuver as the panel chattering vibration under deployment was effectively reduced. deployment was effectively reduced. Table 7. Free-vibration test results of the solar panels in rigidly mounted condition. Table 7. Free-vibration test results of the solar panels in rigidly mounted condition.

SolarSolar PanelPanel 1st 1stEigenfrequency Eigenfrequency (Hz) (Hz) Damping Damping Ratio Ratio TypicalTypical PCBPCB panelpanel 82.6 82.6 0.036 0.036 VMLSAVMLSA 110.1110.1 0.1410.141

5. Experimental Validation 5.1. Solar Panel Deployment Test To validate the proposedproposed mechanism’smechanism’s stablestable release function,function, the deploymentdeployment tests of the the solar solar panel panel were were executed executed at an at ambien an ambientt room room temperature temperature of 25 of°C 25with◦C the with exper- the imentalexperimental test setup test setup shown shown in Figure in Figure 10. 10The. The qualification qualiﬁcation model model of of the the P-HRM P-HRM electrical system illustrated inin FigureFigure6 6was was electrically electrically connected connected to to a powera power source source for for the the triggering trigger- ingmechanism mechanism and and a data a data acquisition acquisition (DAQ) (DAQ) system system to determine to determine the solarthe solar panel panel deployment deploy- mentstatus. status. In addition, In addition, an accelerometer an accelerometer at the center at the of center the solar of the panel solar was panel attached was attached to measure to measurethe panel the acceleration panel acceleration responses. responses. The solar The panel’s solar random panel’s equivalent random equivalent static analysis static result anal- ysisshowed result that showed at least that triple at least wire triple winding wire winding is required is required to secure to a secure positive a positive margin margin with a withsafety a factorsafety offactor 3, which of 3, iswhich not shown is not shown here. Thus, here. to Thus, stow to the stow solar the panel solar as panel it is inas launchit is in launchconﬁguration, configuration, triple wire triple winding wire winding was performed was performed as described as described in the in above-mentioned the above-men- wire-tighteningtioned wire-tightening procedure. procedure.

Aerospace 2021, 8, 64 15 of 24 Aerospace 2021, 8, x FOR PEER REVIEW 15 of 24

Aerospace 2021, 8, x FOR PEER REVIEW 15 of 24

Figure 10. The solar panel deployment test setup. Figure 10. The solar panel deployment test setup. Figure 11 showsshows thethe VMLSAVMLSA release release function function test test results. results. The The time time history history of of the the input in- putvoltage voltageFigure in the in 11 mechanismthe shows mechanism the andVMLSA separationand releaseseparation signal function signal indicated test indicated results. that thethatThe solar thetime solar panel history panel is releasedof theis re- in- leasedatput 0.72 voltage at s from 0.72 in thes thefrom power mecha the triggering powernism a ndtriggering inseparation the circuit. in the signal The circuit. solarindicated The panel’s solar that acceleration panel’sthe solar acceleration panel response is re- responseindicatesleased at thatindicates 0.72 the s from complete that thethe deploymentpowercomplete triggering deployment of the in panel the of tookcircuit.the panel 1.58 The s took after solar 1.58 the panel’s releases after accelerationthe action release was actioncompleted.response was indicates completed. The test that results The the validatedtestcomplete results thedeployment validated release function the of therelease panel of the function took mechanism. 1.5 of8 sthe after Asmechanism. the intended release Asinaction theintended design, was completed.in thethe deploymentdesign, The the test deployment status results of validated the status solar of panelthe the releasesolar was panel conﬁrmed function was confirmedof according the mechanism. accord- to the ingelectricAs tointended the current electric in ﬂowthe current design, status flow the through deployment status the through resistor status the mounted of resistor the solar on mounted thepanel resistor wa ons confirmedthe PCB resistor using accord- pogoPCB usingpinsing andto pogo the a voltage electricpins and dividercurrent a voltage circuit.flow divider status circuit. through the resistor mounted on the resistor

Figure 11. The solar panel’s deployment test results. FigureFigure 11. 11.The The solar solar panel’s panel’s deployment deployment test test results. results.

Aerospace 2021, 8, x FOR PEER REVIEW 16 of 24

Aerospace 2021, 8, 64 16 of 24 The function of the deployment mechanism must be confirmed by conducting at least 10 repeatability checks at room temperature, as several deployment tests are required during theThe development function of the of deploymentCubeSat and mechanism testing phases must such be confirmed as run-in by and conducting acceptance at testing [32].least 10The repeatability mechanism checks was at therefore room temperature, repeatedly as severalreleased deployment in the same tests test are configuration required as shownduring thein Figure development 10. The of main CubeSat objectives and testing of the phases tests such were as run-into verify and the acceptance repeatability test- of the mechanisming [32]. The mechanismand to identify was therefore the release repeatedly time releaseddifference in thebetween same test single configuration and triple wire windings.as shown in The Figure release 10. The times main are objectives shown of in the Figure tests were12. The to verify system the repeatabilityfunctioned well of in 10 the mechanism and to identify the release time difference between single and triple wire repetitive activations without electrical malfunction on the circuit and no failure on the windings. The release times are shown in Figure 12. The system functioned well in 10 mountedrepetitive activationsresistor. However, without electrical the release malfunction time was on noticeably the circuit andvaried no failurein repetitive on the release testsmounted because resistor. of the However, variable the workmanship release time was of applied noticeably tension varied on in the repetitive knot of release the wire. The mechanism’stests because of average the variable release workmanship time with ofone, applied two, tensionand three on thewire knot windings of the wire. in 10 The repetitive triggersmechanism’s for each average case release were 0.77 time s, with 0.89 one, s, and two, 1.1 and s, threerespectively. wire windings Thus, inP-HRM 10 repetitive has adequate repetitivetriggers for release each case capability were 0.77 s,for 0.89 application s, and 1.1 s, in respectively. actual space Thus, missions. P-HRM hasThere adequate is no fast re- leaserepetitive requirement, release capability but it foris advantageous application in actual to cut space the missions.wire as soon There as is possible no fast release after the ini- tiationrequirement, of burn but wire it is advantageoustriggering in tothe cut mechanism the wire as soonto prevent as possible electrical after the malfunction initiation in the of burn wire triggering in the mechanism to prevent electrical malfunction in the circuit. circuit. Furthermore, Park et al. [8] performed solar panel release tests as a function of the Furthermore, Park et al. [8] performed solar panel release tests as a function of the wire wirethickness thickness within within the qualification the quali temperaturefication temperature range of − range40 to +60 of ◦−C.40 The to +60 release °C. timeThe release timeof the of panel the panel was slightly was slightly higher athigher low temperatures at low temperatures than that inthan the that high-temperature in the high-tempera- tureenvironment environment because because of the environmental of the environmental effect on the effect variation on the in variation the heating in time the of heating the time ofresistor. the resistor. However, However, the release the time release variation time from variation 0 to +60 from◦C was0 to nominal. +60 °C was nominal.

Number of windings = 1 Number of windings = 2 Number of windings = 3 2

1.5

1 Release (s) time

0.5

0 246810 Number of tests

Figure 12.12.Release Release times times of of the the proposed proposed mechanism mechanism in repetitive in repetitive release release function function test. test. Relative to the space radiation environment, the radiation hardness of electronic devicesRelative is deﬁned to the according space radiation to two key environment, aspects: cumulative the radiation effects andhardness single of event electronic ef- de- vicesfects [ 33is ].defined The most according important to parameter two key to aspects: consider cumulative when selecting effects electrical and componentssingle event effects [33]. The most important parameter to consider when selecting electrical components for space application is the tolerance of the cumulative effects of ionizing radiation. The TID test of the electrical interface PCB was conducted under the low-level energy spectrum of

Aerospace 2021, 8, 64 17 of 24

for space application is the tolerance of the cumulative effects of ionizing radiation. The TID test of the electrical interface PCB was conducted under the low-level energy spectrum of 1.17 to 1.33 MeV by Cobalt-60 (60Co) radiation source for radiation hardness assurance of the P-HRM in a harsh space environment. The dose rate of 60Co was 1.67 krad per hour. The TID test duration was set to 1 h because the solar panel deployment is generally initiated within 1 h of the orbital injection of CubeSat, which gives conﬁdence that the mechanism will not have in-orbit issues to release the panel due to radiation exposure. The voltage on the pogo pin was monitored during the test period to observe the electrical malfunction of the circuit due to the cumulative radiation exposure. The TID test result did not report any drastic variation in voltage in the circuit, while the electrical interface PCB was exposed to a 1 h cumulative radiation dose. However, the 1-year TID estimated by the space environment information system (SPENVIS) software for STEP Cube Lab-II in accordance with the system description is 10 krad. The mechanism will be activated within 1 h of the orbital injection of STEP Cube Lab-II for solar panel deployment, although further TID testing of P-HRM is scheduled to be executed under the projected 1-year total ionizing dose of 10 krad in order to assure radiation hardness over the mission life. Furthermore, the SEE test of the electrical interface PCB of P-HRM was performed to ensure that the electrical components were robust to the in-orbit expected radiation environment. The test specimen was subjected to a high-level energy spectrum of 100 MeV in a 5 krad radiation environment. The electrical current at the pogo pin was measured during the test. The result exhibited steadiness of the current, which revealed that no SEE event occurred in the circuit. The results of these radiation tests did not report a malfunction or electrical failure in the electrical system of the P-HRM.

5.2. Launch Vibration Environment Test The launch vibration tests of the proposed solar panel module were performed at ambient room temperature (20 ◦C) under the qualification-level sinusoidal vibration [34] and random vibration [35] launch loads to verify the damping performance and structural safety in a launch environment. Figure 13 illustrates an example of the launch vibration test setup of the VMLSA on a vibration shaker (J260/SA7M, IMV Corp., Osaka, Japan). To monitor the input vibration loads on the specimen, accelerometers were mounted on the test jig of the solar panel and the slip table of the vibration shaker. During the test, the output acceleration responses of the solar panel module were obtained using an accelerometer attached to the middle of the solar panel. The structural safety of the solar panel under vibration loads was validated by comparing first eigenfrequencies in low-level sine sweep (LLSS) tests executed before and after full-level vibration tests; the variation of first eigenfrequency in the LLSS tests must be within 5%. Additionally, after performing all vibration tests, a solar panel deployment test was executed to determine mechanism reliability. A modal survey test was conducted before and after the aforementioned full-level vibration tests to determine the first eigenfrequency of the solar panel by a low-level sinusoidal vibration excitation on the specimen with an amplitude of 0.5 g. The z-axis LLSS response of the solar panels in the same axis excitation conducted prior to the full-level vibration tests is shown in Figure 14. The first eigenfrequency of the VMLSA in the launch stowed configuration was 75.0 Hz, which is higher than that of a typical PCB solar panel by a factor of 1.53. Aerospace 2021, 8, x 64 FOR PEER REVIEW 1818 of of 24 24 Aerospace 2021, 8, x FOR PEER REVIEW 18 of 24

Figure 13. An example of launch vibration test setup of VMLSA on a vibration shaker. FigureFigure 13. 13. AnAn example example of of launch launch vibration vibration test test setup setup of of VMLSA VMLSA on on a vibration shaker. Input profile (20~500 Hz) Input profile (20~500st Hz) Typical PCB (1 st Eigenfreq. 48.9 Hz) Typical PCBst (1 Eigenfreq. 48.9 Hz) VMLSA (1st Eigenfreq. 75.0 Hz) 10 VMLSA (1 Eigenfreq. 75.0 Hz) 10

1 1

Acceleration (g) 0.1

0.01 20 100 500 0.01 Frequecncy (Hz) 20 100 500 Figure 14. Low-level sine sweepFrequecncy test results in z-axis (Hz) excitation.

FigureFigure 14. 14. Low-levelLow-level sine sine sweep sweep test test results results in in zz-axis-axis excitation. excitation.

Aerospace 2021, 8, 64x FOR PEER REVIEW 1919 of of 24

TheThe VMLSA’s VMLSA’s corresponding corresponding axis axis sine sine vibration vibration test test results results during during the the x-,x y-,-,y and-, and z- axisz-axis excitations excitations are are shown shown in inFigure Figure 15. 15 The. The qualification qualiﬁcation level level of full-level of full-level sinusoidal sinusoidal vibrationvibration loads loads for for small small satellites, satellites, as preferre as preferredd by the by theQB50 QB50 System System requirements requirements and andrec- ommendationsrecommendations [34], [34 was], was applied applied along along the the axis axis of of the the solar solar panel, panel, but but the the zz-axis-axis was was the the mostmost important important axis axis because because it it produced produced the the highest highest dynamic dynamic deflecti deﬂectionon in incomparison comparison to theto the others. others. The The solar solar panel’s panel’s corresponding corresponding x-, yx-,-, andy-, and z-axesz-axes maximum maximum acceleration acceleration re- sponsesresponses under under a sinusoidal a sinusoidal vibration vibration input input load load of of 2.5 2.5 g gin in each each excitation excitation axis axis were were 2.5, 2.5, 2.5,2.5, and and 10.26 10.26 g, g, respectively. respectively. A maximum 10.2 10.266 g resonance response was observed at 75 HzHz on on the the zz-axis-axis of of the the VMLSA VMLSA during during the the same axis excitation.

Input profile (max. 2.5 g) x-axis res. (max. 2.5 g) y-axis res. (max. 2.5 g) z-axis res. (max. 10.26 g) 100

Acceleration (g) 1

0.1 10 100 Frequency (Hz)

FigureFigure 15. 15. TheThe results results of of VMLSA’s VMLSA’s corresponding corresponding axis axis sine sine vibration vibration tests tests in inthe the x-,x y-,-,y and-, and z-axisz-axis excitations.excitations.

TheThe VMLSA’s corresponding axis random vibr vibrationation test results for each excitation axisaxis are shown in Figure 1616.. The Grmsrms valuesvalues in in each each corresponding axis calculated from thethe PSD PSD acceleration acceleration profiles proﬁles under under the the x-,, y-,, and and z-axis-axis excitations excitations were 16.57, 16.36, and 13.51, respectively. Compared toto thethe inputinput levellevel randomrandom vibrationvibration proﬁleprofile ofof 14.114.1 GGrmsrms, the GGrmsrms ofof VMLSA VMLSA was was lower lower by by a a factor factor of of 1.04 1.04 in in the the zz-axis-axis owing owing to to the high damping achievedachieved by by the the attached attached stiffeners.

AerospaceAerospace 20212021, 8, 8x, FOR 64 PEER REVIEW 20 of 24 20 of 24

Input profile (max. 14.1 Grms ) x-axis res. (max. 16.57 Grms ) y-axis res. (max. 16.36 Grms ) z-axis res. (max. 13.51 Grms ) 10

1 2

0.1

PSD Acceleration(g PSD /Hz) 0.01

0.001 100 1000 Frequency (Hz)

FigureFigure 16. TheThe results ofof VMLSA’sVMLSA’s corresponding corresponding axis axis random random vibration vibration tests tests in the in xthe-, y x-,-, and y-, and z- axisz-axis excitations. excitations.

The resultsresults of of the the LLSS LLSS tests tests conducted conducted before before and afterand theafter full-level the full-level sinusoidal sinusoidal and and random vibration tests are summarized in Table8. The ﬁrst eigenfrequency shift of the random vibration tests are summarized in Table 8. The first eigenfrequency shift of the VMLSA was within 4.85% throughout the test sequences, which is within the 5% criterion. VMLSAThe visual was inspection within 4.85% of the solarthroughout panel module the test performed sequences, after which all the is vibration within the tests 5% did criterion. Thenot reportvisual dissociation inspection and of the plastic solar deformation panel module of the performed stiffeners. Thus, after the all structuralthe vibration safety tests did notof the report solar dissociation panel under launchand plastic vibration deformation loads was of validated the stiffeners. at the qualiﬁcation Thus, the structural level. safety of theFurthermore, solar panel inunder the middle launch of vibration the VMLSA, loads the was maximum validated relative at the dynamic qualification displace- level. ment under the random vibration load calculated by the 3-sigma value of acceleration Grms Tableresponse 8. Results was 0.12 of mm.the VMLSA’s The dynamic low-level displacement sine sweep of the(LLSS) VMLSA tests isconducted reduced by before a factor and of after full- level4.3 relative vibration to atests. typical PCB solar panel owing to superior damping resulting from shear deformation of viscoelastic adhesive tapes. Excitation Corresponding Axis Frequency Shift Test Status Axis 1st Eigenfrequency (Hz) Difference (%) Before 835.2 x 0.19 After 833.6 Before 602.4 Sine vibration y 0 After 602.4 Before 75.0 z 0.93 After 74.3 Before 833.6 x 0.19 After 832.0 Before 602.4 Random vibration y 0 After 602.4 Before 74.3 z 4.85 After 70.7

Aerospace 2021, 8, 64 21 of 24

Table 8. Results of the VMLSA’s low-level sine sweep (LLSS) tests conducted before and after full-level vibration tests.

Corresponding Axis Frequency Shift Test Excitation Axis Status 1st Eigenfrequency (Hz) Difference (%) Before 835.2 x 0.19 After 833.6 Before 602.4 Sine vibration y 0 After 602.4 Before 75.0 z 0.93 After 74.3 Before 833.6 x 0.19 After 832.0

Random vibration Before 602.4 y 0 After 602.4 Before 74.3 z 4.85 After 70.7

5.3. Thermal Vacuum Test A TV test of the VMLSA was performed by exposing six thermal cycles at a −40 to 60 ◦C qualification temperature range in a φ1 m TV chamber with a pressure lower than 10−5 torr to verify the solar panel’s structural safety and the release function of the P-HRM in the space environment. To measure the specimen’s temperature, thermocouples were mounted on the solar panel and electrical interface PCB. In order to determine the stabilized target temperature of the solar panel module, a thermocouple attached to the electrical interface PCB was taken as a temperature reference point (TRP). On the specimen, the target temperature stabilization was achieved by controlling the TV chamber’s shroud temperature at the 1 ◦C per hour rate. However, at the hot and cold plateaus of each cycle, the length of dwell time was fixed to 1 h. The P-HRM’s state of health (SOH) check was conducted by visual inspection at each dwelling time. The solar panel deployment test was performed at −20 ◦C during the third cold soak phase, as it is the worst state for triggering the mechanism after satellite ejection into an orbit or trajectory. The solar panel was released without any anomalies 5.20 s after the input voltage was triggered in the mechanism. However, owing to the difference in the heating time of the resistor to cut the Dyneema wire, the release time was marginally higher than that in ambient room conditions. In addition, the electrical power dissipation from the power supply wire to the mechanism within the chamber at an extreme cold temperature and the relay time delay of the solar panel deployment signal to the DAQ system are also factors that increase the release time [36]. However, the onboard batteries of CubeSats will have sufficient power to deploy solar panels, as release action is typically carried out at the initial stage of orbital injection. After the solar panel deployment test, the remaining cycles were completed within the qualification temperature range of the solar panel. After accomplishment of the TV test, the solar panel deployment test was performed at 25 ◦C room environment and the result is presented in Figure 17. The results of the release function tests of the mechanism obtained before and after vibration tests and during the TV test are plotted in the same Figure 17 to compare the release time in each event. The release time of the mechanism after the TV test was 1.5 s; the minimal difference in release time compared to that determined before the TV test is due to the thermal overstress of the resistor in six thermal cycles. Thus, these release function test results guarantee solar panel deployment in an orbit environment. Aerospace 8 Aerospace 20212021,, 8,, x 64 FOR PEER REVIEW 2222 of of 24 24

Input voltage Separation signal 10 During TV test: 5.20 s 8 After vibration tests: 0.91 s

6 Before vibration After TV tests: 1.5 s test: 0.72 s 4 Voltage Voltage (V) Burn wire 2 triggering initiation (power supply on)

0 0 5 10 15 20 Time (s)

FigureFigure 17. 17. SummarySummary of of release release time time of of the the mechanism mechanism at at each each test test event. event.

TheThe solar solar panel panel was was examined examined microscopically microscopically using using optical optical microphotographs, microphotographs, fol- fol- lowinglowing all all the the above-mentioned above-mentioned tests. tests. The The si siderealdereal edge edge optical optical microphotographs microphotographs taken taken onon the the solar solar panel panel before before and and after after the the TV TV te testssts did did not not show show cracks, cracks, plastic plastic deformation, deformation, oror dissociation dissociation in in the the attached attached stiffeners. stiffeners. Th Theseese qualification-level qualification-level tests tests and and inspection inspection resultsresults validated validated the the structural structural safety safety of the of theVMLSA VMLSA under under launch launch and andin-orbit in-orbit thermal thermal environments.environments. The The optimized optimized design design of ofthe the HRM, HRM, simplified simplified electrical electrical circuit, circuit, and and solar solar panelpanel deployment experimentsexperiments carried carried out out under unde differentr different test test conditions conditions led toled confidence to confi- dencein the in performance the performance of the of deployment the deployment system. system. The highly The highly damped damped deployable deployable solar panelsolar panelmodule module proposed proposed herein herein is effective is effective for guaranteeing for guaranteeing the structural the structural safety safety of solar of solar cells cellsunder under launch launch environment environment without without reducing reduci theng area the of area solar of cell solar attachment cell attachment and assuring and assuringreliable releasereliable action release of action the panel of the in apanel space in environment. a space environment. 6. Conclusions 6. Conclusions In this study, a 6U sized highly damped PCB-based deployable solar panel module In this study, a 6U sized highly damped PCB-based deployable solar panel module combined with an optimized pogo pin-based burn HRM was proposed and tested for use combined with an optimized pogo pin-based burn HRM was proposed and tested for use in the STEP Cube Lab-II CubeSat. The proposed solar panel ensures the structural safety in the STEP Cube Lab-II CubeSat. The proposed solar panel ensures the structural safety of solar cells under severe launch environments by reducing the solar panel’s dynamic of solar cells under severe launch environments by reducing the solar panel’s dynamic acceleration and deflection owing to the high damping characteristics achieved by thin acceleration and deflection owing to the high damping characteristics achieved by thin stiffeners and adhesive tapes. The novel three-pogo pin-based HRM has several advantages, stiffenersincluding and high adhesive loading capabilitytapes. The in novel launch three-pogo configuration, pin-based simpler HRM electrical has system,several andad- vantages,guaranteed including solar panel high deploymentloading capability in a space in launch environment. configuration, In addition, simpler electrical the radiation system,test resultsand guaranteed of the electrical solar panel interface deployment PCB confirmed in a space the environment. radiation hardness In addition, of the P-HRM the ra- diationto deploy test the results solar of panel. the electrical Further interfac TID testinge PCB of confirmed the mechanism the radiation will be hardness performed of atthe a P-HRMsystem to level deploy of P-HRM the solar under panel. the Further estimated TID testing one-year of the total mechanism ionizing dose will ofbe 10performed krad for atSTEP a system Cube Lab-IIlevel of to P-HRM assure radiation under the hardness estimated over one-year the mission total life. ionizing Compared dose toof a10 typical krad forPCB STEP solar Cube panel, Lab-II the relativeto assure dynamic radiation displacement hardness over at the centermission of life. the Compared VMLSA under to a typicalrandom PCB vibration solar panel, loads wasthe decreasedrelative dynamic by a factor displacement of 4.3. The at solar the panel center deployment of the VMLSA tests underconducted random after vibration the vibration loads andwas TV decreased tests validated by a factor the reliabilityof 4.3. The of solar the mechanism.panel deploy- In mentthe near tests future, conducted the solar after panel the vibration module and flight TV model tests validated will be manufactured the reliability and of the validated mechanism.again byIn carryingthe near future, out launch the solar vibration panel and module TV tests flight at model the acceptance will be manufactured level. This design and validatedstrategy couldagain by also carrying be advantageous out launch vibration for rapid and attenuation TV tests ofat the in-orbit acceptance vibration level. on This the designpanel causedstrategy by could the satellite also be duringadvantageous the slew for maneuver, rapid attenuation which could of in-orbit significantly vibration reduce on thethe panel performance caused by degradation the satellite of during satellites the where slew maneuver, the rapid acquisition which could of significantly the target point re- duceis required. the performance degradation of satellites where the rapid acquisition of the target point is required.

Aerospace 2021, 8, 64 23 of 24

Author Contributions: Conceptualization, S.B., H.K. and H.-U.O.; methodology, S.B., J.-S.G., and H.-U.O.; software, S.B. and J.-S.G.; formal analysis, S.B. and J.-S.G.; validation, S.B., J.-S.G. and H.-U.O.; writing—original draft preparation, S.B.; writing—review and editing, H.-U.O.; supervision, H.K. and H.-U.O.; funding acquisition, H.-U.O. All authors have read and agreed to the published version of the manuscript. Funding: This research was funded by the Ministry of Science and ICT (MSIT). Institutional Review Board Statement: Not applicable. Informed Consent Statement: Not applicable. Data Availability Statement: The data used to support the findings of this study are available from the corresponding author upon request. Acknowledgments: This research was supported by the Korea Aerospace Research Institute (KARI) and funded by the Ministry of Science and ICT (MSIT). Conflicts of Interest: The authors declare no conflict of interest.

References 1. Poghosyan, A.; Golkar, A. CubeSat evolution: Analyzing CubeSat capabilities for conducting science missions. Prog. Aerosp. Sci. 2017, 88, 59–83. [CrossRef] 2. Santoni, F.; Piergentili, F.; Donati, S.; Perelli, M.; Negri, A.; Marino, M. An innovative deployable solar panel system for CubeSats. Acta Astronaut. 2014, 95, 210–217. [CrossRef] 3. Anigstein, A.; Pena, R.S.S. Analysis of solar panel orientation in low altitude satellites. IEEE Trans. Aerosp. Electron. Syst. 1998, 34, 569–578. [CrossRef] 4. Wijker, J.J. Spacecraft Structures, 1st ed.; Springer: Leiden, The Netherlands, 2008; pp. 1–504. 5. Ampatzoglou, A.; Kostopoulos, V. Design, analysis, optimization, manufacturing, and testing of a 2U CubeSat. Int. J. Aerosp. Eng. 2018, 2018.[CrossRef] 6. Chau, V.M.; Vo, H.B. Structural dynamics analysis of 3-U CubeSat. Appl. Mech. Mater. 2019, 894, 164–170. [CrossRef] 7. Small Satellite, CubeSat Solar Panels-ISISPACE-Innovative Solutions in Space, Small Satellite Solar Panels. Available online: https://www.isispace.nl/product/isis-cubesat-solar-panels/ (accessed on 28 December 2020). 8. Park, T.Y.; Chae, B.G.; Oh, H.U. Development of 6U CubeSat’s deployable solar panel with burn wire triggering holding and release mechanism. Int. J. Aerosp. Eng. 2019, 2019.[CrossRef] 9. AZUR SPACE Solar Power GmbH, SPACE Solar Cells. Available online: http://www.azurspace.com/index.php/en/products/ products-space/space-solar-cells (accessed on 28 December 2020). 10. Lim, L.S.; Bui, T.D.V.; Low, K.S.; Tissera, M.S.C.; Pham, V.H.P.; Abhishek, R.; Soon, J.J.; Lew, J.M.; Aung, H.; Goh, S.T.; et al. VELOX-II: Challenges of developing a 6U nanosatellite. In Proceedings of the AIAA SPACE Conference and Exposition, Long Beach, CA, USA, 13–16 September 2016; pp. 1–11. 11. Lee, S.; Hutputanasin, A.; Toorian, A.; Lan, W.; Munakata, R. CubeSat Design Speciﬁcation Rev. 12; California Polytechnic State University: San Luis Obispo, CA, USA, 2009; pp. 1–22. Available online: https://srl.utu.ﬁ/AuxDOC/tke/radmon/cubesat_ standard.pdf (accessed on 28 December 2020). 12. NanoPower DSP-GOMspace, Deployable Solar Panels for 3U and 6U Satellites. Available online: https://gomspace.com/shop/ subsystems/power/nanopower-dsp.aspx (accessed on 28 December 2020). 13. Solar Arrays-MMA Design LLC, Solar Arrays. Available online: https://mmadesignllc.com/products/solar-arrays/ (accessed on 25 January 2021). 14. Thurn, A.; Huynh, S.; Koss, S.; Oppenheimer, P.; Butcher, S.; Schlater, J.; Hagan, P. A nichrome burn wire release mechanism for CubeSats. In Proceedings of the 41st Aerospace Mechanisms Symposium, Pasadena, CA, USA, 16–18 May 2012; pp. 479–488. 15. Steinberg, D.S. Vibration Analysis for Electronic Equipment, 3rd ed.; John Wiley and Sons Inc.: New York, NY, USA, 2000; pp. 1–440. 16. Minesugi, K.; Onoda, J. Passive vibration suppression using thin tape with viscous lamina. In Proceedings of the AIAA 36th Structures, Structural Dynamics and Materials Conference, New Orleans, LA, USA, 10–13 April 1995; pp. 200–206. 17. Park, T.Y.; Shin, S.J.; Park, S.W.; Kang, S.J.; Oh, H.U. High-damping PCB implemented by multi-layered viscoelastic acrylic tapes for use of wedge lock applications. Eng. Fract. Mech. 2021, 241, 107370. [CrossRef] 18. Bhattarai, S.; Kim, H.; Oh, H.U. CubeSat’s deployable solar panel with viscoelastic multi-layered stiffener for launch vibration attenuation. Int. J. Aerosp. Eng. 2020, 2020.[CrossRef] 19. Kang, S.J.; Oh, H.U. On-orbit thermal design and validation of 1U standardized CubeSat of STEP Cube Lab. Int. J. Aerosp. Eng. 2016, 2016.[CrossRef] 20. Components-Blue Canyon Technologies, Attitude Control System, XACT-15. Available online: https://www.bluecanyontech. com/components (accessed on 28 December 2020). 21. Jang, S.S.; Kim, S.H.; Lee, S.R.; Choi, J. Energy balance and power performance analysis for satellite in low earth orbit. J. Astron. Space Sci. 2010, 27, 253–262. [CrossRef] Aerospace 2021, 8, 64 24 of 24

22. 3M Company, 3MTM Adhesive Transfer Tape 966-Technical Data Sheet. Available online: https://3m.citrination.com/pif/000314 ?locale=en-US (accessed on 28 December 2020). 23. Park, T.Y.; Kim, S.H.; Kim, H.; Oh, H.U. Experimental investigation on the feasibility of using spring-loaded pogo pin as a holding and release mechanism for CubeSat’s deployable solar panels. Int. J. Aerosp. Eng. 2018, 2018.[CrossRef] 24. Bhattarai, S.; Kim, H.; Jung, S.H.; Oh, H.U. Development of pogo pin based holding and release mechanism for deployable solar panel of CubeSat. Int. J. Aerosp. Eng. 2019, 2019.[CrossRef] 25. CFE Company, Spring-Loaded Connectors. Available online: https://www.cfeconn.com/2-54mm-pitch-through-hole-pogo-pin- catalog.html (accessed on 28 December 2020). 26. PSA Company, Resistors-Walsin Technology, Resistors 3216 SMD. Available online: http://www.passivecomponent.com/ products/resistors/ (accessed on 28 December 2020). 27. Optocoupler, Components101, PC817 Photo-Coupler IC. Available online: https://components101.com/ics/pc817-ic-pinout- equivalent-datasheet (accessed on 28 December 2020). 28. Bhattarai, S.; Go, J.S.; Kim, H.; Oh, H.U. Experimental validation of a highly damped deployable solar panel module with a pogo pin-based burn wire triggering release mechanism. Int. J. Aerosp. Eng. 2020, 2020.[CrossRef] 29. YGK Fishing Wire, YGK G-Soul Super Jigman Braid. Available online: https://www.tsourosmarine.gr/en/braids/1402-ygk-g- soul-super-jigmanbraid.html#/12-colormulticolored/4816-diameter_mm_cm-0_205_mm (accessed on 28 December 2020). 30. Tian, S.; Xu, Z.; Wu, Q.; Qin, C. Dimensionless analysis of segmented constrained layer damping treatments with modal strain energy method. Shock Vib. 2016, 2016.[CrossRef] 31. Liu, T.X.; Hua, H.X.; Zhang, Z. Robust control of plate vibration via active constrained layer damping. Thin Walled Struct. 2004, 42, 427–448. [CrossRef] 32. ECSS-E-ST-33-01C Rev.2-Mechanisms. Space Engineering Mechanisms, ECSS Secretariat ESA-ESTEC Requirements & Standards Division; European Cooperation for Space Standardization: Noordwijk, The Netherlands, 2019; pp. 1–97. Available online: https: //ecss.nl/standard/ecss-e-st-33-01c-rev-2-1-march-2019-space-engineering-mechanisms/ (accessed on 28 December 2020). 33. Duzellier, S. Radiation effects on electronic devices in space. Aerosp. Sci. Technol. 2005, 9, 93–99. [CrossRef] 34. Singarayar, F.; Reinhard, R.; Asma, C.; Thoemel, J.; Scholz, T.; Bernal, C.; Weggelaar, W.; Kataria, G.S.D.; Richard, M. QB50 System Requirements and Recommendations. 2013, pp. 1–28. Available online: https://www.qb50.eu/index.php/techdocs/category/ QB50_Systems_Requirements_issue_44835.pdf?download=30:qb50-docs (accessed on 28 December 2020). 35. GEVS: GSFC-STD-7000A. General Environmental Veriﬁcation Standard (GEVS) for GSFC Flight Programs and Projects; NASA Goddard Space Flight Center: Greenbelt, MD, USA, 2013; pp. 1–203. Available online: https://standards.nasa.gov/standard/gsfc/gsfc- std-7000 (accessed on 28 December 2020). 36. Malagoli, M.; Allewaert, Y. Improvement of the wire rating standards based on TV testing and thermal modeling. In Proceedings of the 48th International Conference on Environmental Systems, Albuquerque, NM, USA, 8–12 July 2018; pp. 1–15.