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; shankar@chosun.kr (S.B.);
jeesung529@chosun.kr (J.-S.G.)
2 Soletop Co. Ltd., 409 Expo-ro, Yuseong-gu, Daejeon 34051, Korea; hrkim@soletop.com
* Correspondence: ohu129@chosun.ac.kr

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
published maps and institutional affil-
of a deployable solar panel under a lunch vibration environment causes stress on mounted
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
distributed under the terms and
various configurations have been produced owing to the advantages of speedy fabrication
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) com-
bined 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
Figure 1 illustrates the system architecture of the STEP Cube Lab-II CubeSat mis-
sion. 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 technolo-
gies 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-optical camera (EOC),
electro-optical camera (EOC),broadband
broadbandinfrared
infraredcamera
camera(BBIRC),
(BBIRC), and
and long-wave
long-wave infra-
infrared
red camera
camera (LWIRC).
(LWIRC). The primary
The primary objective
objective of theof the STEP
STEP Cube Lab-II
Cube Lab-II projectproject is toout
is to carry carry
the
out the nation’s
nation’s first multi-band
first multi-band earth observation
earth observation missionmission by utilizing
by utilizing CubeSat’s
CubeSat’s electro-
electro-optical
optical
(EO) and (EO) and infrared
infrared (IR) images
(IR) images and videos
and videos on various
on various targets
targets of theof Korean
the Korean penin-
peninsula,
sula, including
including Mt. Paektu,
Mt. Paektu, which which has recently
has recently shown shown
signssigns of a volcanic
of a volcanic eruption,
eruption, obser-
observation
of areas
vation ofaffected by forest
areas affected fires, and
by forest fires,the
androute of spread
the route in areas
of spread wherewhere
in areas forestforest
fires have
fires
occurred,
have analysis
occurred, of theofthermal
analysis islandisland
the thermal phenomenon
phenomenon in urban areas areas
in urban of metropolitan
of metropolitancities,
observation
cities, of ship
observation ofactivities, and monitoring
ship activities, and monitoring of theofnuclear power
the nuclear plant’s
power coolant
plant’s dis-
coolant
charge in seawater.
discharge in seawater.Furthermore,
Furthermore, one ofonetheofspace technologies
the space to be verified
technologies in this mission
to be verified in this
is an optimized
mission data compression
is an optimized method method
data compression in the payload data handling
in the payload system (PDHS)
data handling system
and payload data transmission system (PDTS) for video data
(PDHS) and payload data transmission system (PDTS) for video data processing and processing and transmission
to the groundtostation.
transmission the groundAdditionally, this missionthis
station. Additionally, alsomission
tests a high-damping VMLSA to
also tests a high-damping
ensure structural
VMLSA to ensureprotection
structural of mountedofsolar
protection cells insolar
mounted launchcellsvibration
in launch environments
vibration envi- by
suppressing
ronments by transmitted
suppressinglaunch loads on
transmitted the solar
launch loadspanel
on theandsolar
a three-pogo
panel and pin-based
a three-pogo burn
wire triggering
pin-based burnmechanism
wire triggeringfor high holding capability
mechanism for high atholding
the launch stowedat
capability configuration
the launch
of the solar panel and reliable release action in space. Furthermore,
stowed configuration of the solar panel and reliable release action in space. Furthermore,an optimized SADA for
the acquisition of solar power maximization is one of the secondary
an optimized SADA for the acquisition of solar power maximization is one of the second- payloads. This study
is mainly focused on the on-ground experimental validation of
ary payloads. This study is mainly focused on the on-ground experimental validation of the design effectiveness
anddesign
the structural safety of the
effectiveness andVMLSA combined
structural safety ofwith P-HRM. combined with P-HRM.
the VMLSA

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

Table
Table 1 presents the detailed specifications
specifications of of the
the CubeSat
CubeSatsystem.
system. For
For mission
missionopera-
oper-
ation, thesatellite
tion, the satelliteorbit
orbitisisselected
selected as
as aa sun-synchronous
sun-synchronous orbit of 700 km altitude from the
earth surface
earth surface and
andananinclination
inclinationangle
angleof 98.2 ◦ withwith
of 98.2° a local
a time
local oftime
descending node (LTDN)
of descending node
of 16:00 of
(LTDN) p.m. Thep.m.
16:00 ground station contact
The ground time of time
station contact the satellite calculated
of the satellite from the
calculated orbital
from the
analysis
orbital is approximately
analysis 7.5 min,
is approximately 7.5and
min,access to the to
and access ground stationstation
the ground occursoccurs
six times a day.
six times
aThe CubeSat
day. mission
The CubeSat operation
mission life is set
operation lifetoisone year.
set to 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
Orbital parameter Altitude: 700 km
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
system UHF band: 9600 bps Gaussian minimum
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

Figure 2a,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 observa-
tion 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 pan-
els, 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 de-
ployable 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.
Figure STEP Cube
2. STEP Cube Lab-II’s
Lab-II’s mechanical
mechanical configurations:
configurations: (a)
(a) solar
solar panel
panel stowed
stowed and
and (b)
(b) solar
solar panel
panel deployed.
deployed.

2.3. Design
2.3. Design of Solar
of Solar PanelPanel
Module Module withAccommodation
with the the Accommodation of Solar
of Solar Cell Cell
A deployable
A deployable solar solar
panelpanel
basedbased
on a on
PCB a PCB substrate
substrate usingusing multilayered
multilayered thin stiffeners
thin stiffeners
attached
attached usingusing double-sided
double-sided viscoelastic
viscoelastic tapes istapes is proposed
proposed for STEPforCube
STEPLab-II
CubeCubeSat
Lab-II CubeSat
as
a novel design approach to ensure structural protection of solar cells by suppressing the the
as a novel design approach to ensure structural protection of solar cells by suppressing
transmitted
transmitted vibration
vibration loadsloads on a panel
on a panel in a launch
in a launch environment.
environment. The holding
The holding and releasing
and releasing
action of the solar panel is accomplished using an optimized version
action of the solar panel is accomplished using an optimized version of the pogo pin-based of the pogo pin-based
burn burn wire cutting
wire cutting releaserelease mechanism.
mechanism. The solarThepanel
solardeployment
panel deploymentis set to is set to
occur occur1within
within h
of the1 orbital
h of the orbital of
injection injection
STEP Cubeof STEP
Lab-IICube
fromLab-II from the
the P-POD. TheP-POD.
passive The passive
torsional torsional
force of
force of the
the torsional torsional
hinges deployshinges
and deploys
latches theandpanel
latches theprojected
at its panel at its projected
position. position.
Twenty AZUR Twenty
AZUR space 30% efficiency triple-junction GaAs solar cells (3G30C)
space 30% efficiency triple-junction GaAs solar cells (3G30C) [9] can be mounted on the front [9] can be mounted
onof
surface the
thefront surface
proposed of the
solar panelproposed solardesign
because this panelstrategy
becausedoesthis not
design
reduce strategy does not
the availa-
ble area for solar cell accommodation. Thus, the power generation of the proposed solarof the
reduce the available area for solar cell accommodation. Thus, the power generation
panelproposed solar compared
is maximized panel is maximized
to the same compared to the
solar panel same
area by solar panel area
employing by employing
the aforemen-
tionedthe aforementioned
design strategies todesign
ensurestrategies to ensure
the structural safetytheof structural
solar cells.safety of solar cells.
The energy balance analysis (EBA) of the CubeSat
The energy balance analysis (EBA) of 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]. Table 2 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 Table 1. 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

Table 3 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
Figure 3a,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 Figure 3,
Aerospace 2021, 8,
Aerospace 2021, 8, 64
x FOR PEER REVIEW 88 of
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24

Figure 3, were on
were mounted mounted
the rearon the rear
surface surface
of the PCB of the PCB
panel panel
through 3MTM 966 acrylic
through double-sided
double-sided 3MTM
966 acrylic
tapes tapes
[22]. The TM
3M[22].966
Theacrylic
3M tape
TM 966 acrylic tape is a high-temperature
is a high-temperature adhesive thatadhesive
has beenthat
usedhasin
aerospace
been usedapplications
in aerospacethat comply with
applications thatASTM
comply E596 lowASTM
with volatility
E596specification criteria
low volatility for
specifi-
space application
cation of theapplication
criteria for space National Aeronautics
of the Nationaland Aeronautics
Space Administration
and Space(NASA). For the
Administration
attenuation of transmitted launch loads on the panel, the fundamental principle
(NASA). For the attenuation of transmitted launch loads on the panel, the fundamental is based on
constrained
principle layer damping
is based accomplished
on constrained by the shear
layer damping deformation
accomplished bycharacteristics achieved
the shear deformation
between interlaminated
characteristics achieved thin stiffeners
between and adhesive
interlaminated thintapes. The basic
stiffeners specifications
and adhesive ofThe
tapes. the
materials
basic used in the
specifications ofsolar panel module
the materials used inarethe
listed inpanel
solar Table module
4. are listed in Table 4.

(a) (b)

Figure
Figure 3. Configurationsof
3. Configurations ofaahighly
highlydamped
dampeddeployable
deployablesolar
solarpanel
panel module’s
module’s demonstration
demonstration model:
model: (a) (a) stowed
stowed andand
(b)
(b) deployed.
deployed.

4. Specifications
Table The of materialsprocess
stiffener attachment used in the
on solar panel is
the panel 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
onmodulus
the four (Pa)
edge corners of the PCB panel 18.73 109
and×stiffeners. The
stiffener attachment process on the PCB
Density (kg/m ) panel
3 is shown in Figure 4. First, prepare
1850 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 conductivity
using a cloth moistened(W/mK) 0.29
with IPA to remove dust contamination. The
tape was cut according to the shape of the stiffener by a knife tip, put
Manufacturer 3M into the integration
Company
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
Adhesive
on the last stiffener attached to thickness (mm)
the tape, and 0.06
then the top guide jig has to clamp with M3
bolts fastened on the jig through a
Color torque of 1.1 Nm to create a compression
Transparentforce on the
Adhesive tape [22] specimen. As recommended by the adhesive tape datasheet [22], the adhesive bonding
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 Figure 4. 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
specimen. As recommended by the adhesive tape datasheet [22], the adhesive 9bonding of 24

resin of the attached tapes must be cured for up to 72 h for secure attachment.

Figure4.4.Integration
Figure Integration process
process of
of aa stiffeners
stiffenerson
onthe
thePCB
PCBpanel.
panel.

Thenumber
The number of attached
attachedstiffeners was
stiffeners wasdetermined
determined by the
by 3U
theCubeSat
3U CubeSatsolar panel’s ex-
solar panel’s
perimental test results [18] while considering dynamic clearance while accommodating
experimental test results [18] while considering dynamic clearance while accommodating the
P-POD.
the The
P-POD. thickness
The thickness of the solar
of the panel
solar after
panel thethe
after attachment
attachment of the stiffeners
of the is 3.7
stiffeners mm,
is 3.7 mm,
which provides a lateral edge gap margin of 3.3 mm for a dynamic clearance
which provides a lateral edge gap margin of 3.3 mm for a dynamic clearance on P-POD on P-POD [11].[11].

Table
3.2. 4. Specifications
Solar Panel Holding of and
materials used
Release in the solar panel module.
Mechanism
FigureItem
5a,b illustrate close-up views of the proposed three-pogo pin-based
Details mechanism
Value
in fully and partially stowed solar panel states,
Elastic respectively.
modulus (Pa) The mechanism
18.73 ×comprises
109
pogo pins, electrical interface PCB, brackets, Dyneema
Density (kg/m ) wire, resistor, and resistor
3 1850PCB. Park
FR4
et al. [23] and Bhattarai et al. [24] investigated
Poisson’s the applicability and effectiveness
ratio 0.136 of two
pogo pins in HRM, although the main Thermal conductivity
purpose (W/mK)
of using three pogo pins in the0.29proposed
mechanism is to simplify the electrical circuit Manufacturer
by reducing the electrical3Mcomponents
Company to
lower the cost of the product and increase Adhesive material for secure release action
its reliability Acrylic
in a harsh
space environment. The main design Adhesive
driver thickness
for the(mm) 0.06
use of the pogo pin (MP511-1111-
E03100A, CFE Corporation Co., Guangdong, Color China) [25] connectors is the Transparent
establishment of
a secure electromechanical Allowable
connection temperature
between range (°C)
the electrical interface PCB −40
andtothe232resistor
Adhesive tape [22]
Thermal conductivity at 41 °C (W/mK) 0.178
PCB integrated at the edge of the CubeSat structure and solar panel edge, respectively. The
Coefficient of thermal expansion (ppm/°C) 1.99
pogo pins provide electrical current to the resistor mounted on the resistor 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 restraint 0.93, 0.01 in the
sable material (CVCM) outgassing (%)
in-plane direction of the solar panel during the stowed configuration is accomplished by
ball
3.2.and
Solarsocket joints made
Panel Holding on theMechanism
and Release 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 effective-
ness 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 compo-
nents 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-
 stowed state is achieved by tightening the Dyneema wire on the brackets. Mechanical re-
straint in the in-plane direction of the solar panel during the stowed configuration is ac-
complished 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 working 10 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,
integrated
Taiwan) [26] on was
the resistor
integratedPCBonasthe
anresistor
actuator.
PCB Asasthe
anmechanism
actuator. Asisthe
activated
mechanism by the power
is acti-
supply
vated byin the
an electrical circuit,
power supply in the resistor cuts
an electrical thethe
circuit, tightened
resistorwire. Thetightened
cuts the panel deployment
wire. The is
initiated instantly by
panel deployment the compression
is initiated instantly force
by theofcompression
the pogo pin plunger,
force pogo±pin
of the 0.68 0.19 N, which
plunger,
could rapidly or simply interrupt the circuit
0.68 ± 0.19 N, which could rapidly or simply interruptpath to the resistor. Consequently, the
circuit path to the resistor. Con-hinge
torsional forces deploy the solar panel to its intended angle. Table 5 lists the
sequently, the hinge torsional forces deploy the solar panel to its intended angle. Table 5 specifications
oflists
thethe
hardware used in
specifications ofP-HRM.
the hardware used in P-HRM.

(a) (b)

Figure Configurations
5. 5.
Figure Configurationsofofthree-pogo
three-pogopin-based
pin-based HRM:
HRM: (a) solar panel
(a) solar panelfully
fullystowed
stowedand
and(b)
(b)solar
solar panel
panel partially
partially stowed.
stowed.

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

Items Items Details Details Value

Value
Manufacture CFE Corporation Co.
Manufacture
Voltage and current (V, A) CFE Corporation12, 3 Co.
Voltage andMax. electrical
current (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 Total
Qualification temperature range (◦(mm)
length C) −40 to985
Full stroke (mm) 1.4
Spring forceWorking
(N) stroke (mm) 0.68 ± 1.0
0.19
Total length (mm) Manufacture 9YGK
Material Dyneema
Wire Full stroke (mm) 1.4
Diameter (mm) 0.205
Working stroke Max.(mm)
allowable force (N) 1.088.2
Manufacture Manufacture Walsin technology
YGKCo., Taipei, Taiwan
Package SMD Type
Resistor Material
Electrical resistance (ohm) Dyneema 4.7
Wire
Diameter (mm)
Resistance tolerance (%) 0.205 ±1
Max. power dissipation (W) 0.25
Max. allowable force (N) 88.2
Manufacture
Figure 6a,b illustrate the electrical system ofWalsin technology
the three-pogo Co., Taipei,
pin-based Taiwan
mechanisms.
For mechanism Package
activation, the input voltage on the electrical circuit path
SMD Type to the resistor
Resistor Electrical resistance (ohm) 4.7
Resistance tolerance (%) ±1
Max. power dissipation (W) 0.25

Figure 6a,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 opto-
coupler
(PC817)(PC817) [27] transmits
[27] transmits electrical
electrical signals signals
by lightby light between
between two electrically
two electrically isolatedisolated
circuits,
circuits, thus preventing high-voltage electrical malfunctions in
thus preventing high-voltage electrical malfunctions in the circuit. The the circuit. The powercutoff
power cut-
off function
function on on
thethe mechanism
mechanism is accomplished
is accomplished instantly
instantly afterafter the deployment
the deployment of theofsolar
the solar
panel
panel with
with the the electrical
electrical circuitcircuit
shownshown in thediagram
in the block block diagram
withoutwithout
employingemploying a micro-
a microcontroller
controller
unit (MCU).unit (MCU).

(a)

(b)

Figure6.
Figure Three-pogo pin-based
6. Three-pogo pin-based mechanism’s
mechanism’selectrical
electricalsystem:
system:(a)
(a)front and
front rear
and view
rear of electrical
view interface
of electrical PCBPCB
interface (dimensions:
(dimen-
34 mm × 16 mm × 1 mm) (b) schematic block diagram.
sions: 34 mm × 16 mm × 1 mm) (b) schematic block diagram.

Figure7a,b
Figure 7a,billustrate
illustrateresistor
resistorPCB’s
PCB’sfront
front and
and rear
rear views,
views, respectively.
respectively. A resistor
A resistor as
as an actuator for solar panel release is soldered on the front side of the resistor
an actuator for solar panel release is soldered on the front side of the resistor PCB, which PCB,
which
faces thefaces the outward
outward directiondirection while integrating
while integrating on the
on the solar solar
panel. Topanel. To interconnect
interconnect the elec-
the electrical power coming from the pogo pins to the resistor, three surface-mounted
trical power coming from the pogo pins to the resistor, three surface-mounted electrodes
electrodes
were were
attached attached
to the to the of
rear surface rear
thesurface
resistorofPCB
the that
resistor
wasPCB that was connected
mechanically mechanicallyto
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
electron
scanningmicroscope (SEM) inspection
electron microscope of the electrode
(SEM) inspection of thepads after pads
electrode completing all the vibra-
after completing all
tion tests to evaluate the frictional impact or damage caused by the
the vibration tests to evaluate the frictional impact or damage caused by the tip of tip of the pogo
the pins,
pogo
even
pins,though the spring-loaded
even though pin has
the spring-loaded a low
pin has friction impactimpact
a low friction on vibration loads. loads.
on vibration The thick-
The
ness of theofelectrical
thickness contact
the electrical part part
contact or curvature tip on
or curvature tipthe
on electrodes waswas
the electrodes onlyonly
reduced
reduced by
23.88%
by 23.88% compared
compared to that of of
to that thethe60.37 μm
60.37 µminitial
initialthickness.
thickness.The
Thefrictional
frictionalimpact
impact on on the
the
electrode
electrode under
under vibration
vibration load
load isis minimal;
minimal; thus,
thus, aa secure
secure physical
physical contact
contact can
can be
be assured
assured
even
even after being exposed to severe launch vibration loads because of the 1.7 mm full
after being exposed to severe launch vibration loads because of the 1.7 mm full stroke
stroke
of
of the
the spring-loaded
spring-loaded pinpin of
of the
the pogo
pogo pins.
pins.

(a) (b)

Figure7.
Figure Mechanicalconfiguration
7.Mechanical configurationof
ofthe
theburn
burn resistor
resistor PCB:
PCB: (a)
(a) front
front view
view and
and (b)
(b) 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. Table 6 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
proximately 0 Vdivider
voltage resistor until theis panel is deployed
significantly higherbecause theofresistance
than that value
the resistor usedofasthe
an resistors
actuator
used in the voltage resistor divider is significantly higher than that of the
for panel deployment. The electrical circuit over the resistor becomes open once the resistor used as
solar
an actuator for panel deployment. The electrical circuit over the resistor becomes
panel is deployed such that the output signal of the circuit becomes high, which confirms open
once the solar panel
the deployment is deployed
status of the solarsuch thatTherefore,
panel. the output signal ofto
compared the
thecircuit becomes
mechanism high,
[28], the
which confirms the deployment status of the solar panel. Therefore, compared
P-HRM proposed in this study has a relatively simpler electrical circuit and is well suited to the
mechanism [28],P-POD
to the available the P-HRM
lateralproposed
edge gap. in this study has a relatively simpler electrical cir-
cuit 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 Undeployed Deployed
Deployed
Vin Enable
Vin pin voltage Pogo
Pogo Enable pinvoltageOutput
Output Pogo pin Output
Pogo pin voltage Output
voltage
1 1 1 voltage0 voltage 1 voltage voltage
1
1 10 01 1 0 0 0 1 01
1 0 0 0 0 0
3.3. Wire-Tightening Process
The wire-tightening
3.3. Wire-Tightening procedure on the brackets for the solar panel’s holding mechan-
Process
ical 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 Figure 8. 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
steady
carriedtight tension
out at on thecorner
the bracket wire knot. For
for the reliable
final panel
knotting of release,
the wire,physical contact
which helps to of the
create
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
wire to the resistor
resistor is
is mandatory.
mandatory. Thus,
Thus, the
the notching
notching guide
guide rail
rail is
is made
made on on the
the bracket’s
bracket’s
surface to prevent wire misalignment from the resistor surface under launch vibration
surface to prevent wire misalignment from the resistor surface under launch vibration
loads. Additionally,
loads. Additionally, the
thecombination
combinationofofvertical and
vertical andhorizontal cross-pattern
horizontal cross-patternwinding of the
winding of
wire with two additional hole interfaces in the brackets creates a secure holding
the wire with two additional hole interfaces in the brackets creates a secure holding con- constraint
on the panel.
straint on theThe process
panel. The of wire tightening
process to stow thetosolar
of wire tightening stowpanel is therefore
the solar panel significantly
is therefore
simpler and more reliable than conventional mechanisms.
significantly simpler and more reliable than conventional mechanisms.

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

4.
4. Dynamic
Dynamic Characteristics
Characteristics Investigation
Investigation
In the constrained layer
In the constrained layer damping dampingstrategy,
strategy,thethe transmitted
transmitted vibration
vibration energy
energy dissipa-
dissipation
tion is highly dependent on the viscoelastic core layer thickness, dynamic
is highly dependent on the viscoelastic core layer thickness, dynamic modulus, thickness modulus, thick-
of
ness of constraining
constraining layers,
layers, and and vibration
vibration frequency frequency [30,31]. Therefore,
[30,31]. Therefore, a solar
a solar panel panel free-
free-vibration
vibration test was executed
test was executed at 25 °C ambient
at 25 ◦ C ambient room temperature
room temperature under the under the boundary
boundary condition con-
in
dition in which the interfaces of the hinge and HRM holes were rigidly
which the interfaces of the hinge and HRM holes were rigidly clamped. The roving hammer clamped. The rov-
ing hammer
method wasmethod
used to wasexciteused
theto excite
solar the solar
panel in its panel in its free An
free vibration. vibration. An accelerom-
accelerometer sensor
eter sensor was mounted at the solar panel center to obtain the
was mounted at the solar panel center to obtain the time-domain frequency responses. time-domain frequencyTo
responses.
investigate Totheinvestigate the design
design usefulness usefulnessfor
or practicality or rapid
practicality for rapid
suppression suppression
of panel chatteringof
panel chattering
vibration duringvibration
its in-orbit during its in-orbitthe
performance, performance,
solar panelthe solar panel
module’s module’sprofile
acceleration accel-
eration profile under
under deployment wasdeployment
also measured was also
in themeasured
same room in the same room
temperature temperatureThe
environment. en-
vironment. The dynamic
dynamic characteristics characteristics
investigation investigation
tests testspanel
of a typical PCB of a typical PCB size
of the same panel of the
without
same size additional
attaching without attaching
stiffeners additional
were alsostiffeners
performed were also performed
to compare to compare
the results obtainedthe re-
from
sults obtainedFigure
the VMLSA. from 9a,b
the VMLSA.
show theFigure 9a,b show
solar panel the solarinpanel
time histories time historiestests
the free-vibration in the
at
free-vibration tests atcondition
the rigidly mounted the rigidly andmounted condition
the acceleration andunder
profile the acceleration
deployment, profile under
respectively.
deployment, respectively.
The results illustrate The results
that stiffeners illustrate
with that stiffeners
viscoelastic withcan
acrylic tape viscoelastic
effectivelyacrylic tape
suppress
the transmitted
can vibrationthe
effectively suppress ontransmitted
the solar panel.
vibration on the solar panel.
Aerospace 2021, 8,
Aerospace 2021, 8, 64
x FOR PEER REVIEW 14 of
14 of 24
24

(a) (b)

Figure 9. Dynamic
Figure 9. Dynamic characteristics
characteristics investigation
investigation results
results of
of the
the solar
solar panels:
panels: (a)
(a) time
time histories
histories of
of free-vibration
free-vibration response
response in
in
rigidly
rigidly mounted
mounted condition
condition and
and (b)
(b) acceleration
acceleration profile
profile under
under deployment.
deployment.

Table 77 summarizes
Table summarizes the
the solar
solar panel’s
panel’s free-vibration
free-vibration test
test results
results in
in the
the rigidly
rigidly 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.
Solar
SolarPanel
Panel 1st 1st
Eigenfrequency
Eigenfrequency(Hz)
(Hz) Damping Ratio
Damping Ratio
Typical
TypicalPCB
PCB panel
panel 82.6
82.6 0.036
0.036
VMLSA
VMLSA 110.1
110.1 0.141
0.141

5.
5. Experimental
Experimental Validation
Validation
5.1.
5.1. Solar
Solar Panel
Panel Deployment
Deployment Test Test
To validate the proposed
proposed mechanism’s
mechanism’s stable release function, function, the deployment
deployment tests
of the
thesolar
solarpanel
panelwerewereexecuted
executed at an
at ambient
an ambient room roomtemperature
temperatureof 25 of
°C25with◦ Cthe exper-
with the
imental test setup
experimental shown
test setup shownin Figure
in Figure10.10.
TheThequalification
qualification model
modelofofthe
theP-HRM
P-HRM electrical
system illustrated
illustratedin inFigure
Figure66was waselectrically
electricallyconnected
connectedtoto a power
a power source
source forfor
thethe trigger-
triggering
ing mechanism
mechanism and and a data
a data acquisition
acquisition (DAQ)(DAQ)
systemsystem to determine
to determine the solar
the solar panel panel deploy-
deployment
ment
status.status. In addition,
In addition, an accelerometer
an accelerometer at theof
at the center center of the
the solar solar
panel panel
was was attached
attached to measure to
measure
the panelthe panel acceleration
acceleration responses. responses.
The solarThe solarrandom
panel’s panel’sequivalent
random equivalent staticresult
static analysis anal-
ysis result
showed showed
that that
at least at least
triple wiretriple wire winding
winding is required is required
to securetoasecure
positivea positive
margin margin
with a
safetya factor
with safetyof 3, which
factor of 3, is not shown
which here. Thus,
is not shown here.toThus,
stowto the solar
stow thepanel
solaraspanel
it is in
aslaunch
it is in
configuration,
launch triple wire
configuration, triplewinding was performed
wire winding was performed as described in the in
as described above-mentioned
the above-men-
wire-tightening
tioned wire-tighteningprocedure.
procedure.
Aerospace 2021, 8, 64 15 of 24
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Aerospace 2021, 8, x FOR PEER REVIEW 15 of 24

Figure 10.
Figure 10. The
The solar
solar panel deployment test
panel deployment test setup.
setup.
Figure 10. The solar panel deployment test setup.
Figure 11 shows
shows the theVMLSA
VMLSAreleasereleasefunction
functiontest
testresults.
results.
TheThe time
time history
history of of
thethe in-
input
put voltage
voltage Figure in the mechanism
in the11mechanism
shows the andVMLSA and separation
releasesignal
separation signal
function indicated
test results.
indicated that thethat the
Thesolar solar
timepanel panel
history is re-
of the in-
is released
leased
atput
0.72 at
voltage0.72in
s from s the
the from thetriggering
powerand
mechanism
power triggering
inseparationin the
the circuit. circuit.
signal
The The
indicated
solar solar
panel’s that panel’s
the solaracceleration
acceleration panel is re-
response
response
indicates indicates
leased atthat0.72the that the
s from
completethe deployment
complete
power deployment
triggering
of the in oftook
the
panel the panel
circuit. s took
1.58The 1.58
solar
after s after acceleration
thepanel’s
release the release
action was
action
completed.
responsewasindicates
completed.
The test that The
results test results
thevalidated
complete validated
the the
release function
deployment release
of the function
of the
panel of sthe
mechanism.
took 1.58 mechanism.
As
after intended
the release
As intended
inaction
the design, in the
the deployment
was completed. design, thetest
The deployment
status
results status
solarofpanel
of validated
the the release
the solar
waspanel was confirmed
confirmed
function ofaccording accord-
to the
the mechanism.
ing to
electric the electric
current current
flow statusflow status
through thethrough
resistor the resistor
mounted onmounted
the on
resistor
As intended in the design, the deployment status of the solar panel was confirmed accord- the
PCB resistor
using PCB
pogo
pins
ing and
using topogoa voltage
the pins and
electric divider circuit.
a voltage
current flow divider
statuscircuit.
through the resistor mounted on the resistor

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

Figure11.
Figure 11.The
Thesolar
solarpanel’s
panel’sdeployment
deploymenttest
testresults.
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 leas

10 repeatability checks at room temperature, as several deployment tests are required dur
ing the
Thedevelopment
function of the ofdeployment
CubeSat and testing phases
mechanism must such as run-inbyand
be confirmed acceptance
conducting at testing
[32]. The
least 10 mechanism
repeatability wasattherefore
checks repeatedly
room temperature, released
as several in the same
deployment teststest
areconfiguration
required as
during the development of CubeSat and testing phases such as run-in
shown in Figure 10. The main objectives of the tests were to verify the repeatability of theand acceptance test-
ing [32]. The mechanism
mechanism and to identify was therefore repeatedly
the release time released
difference in the same test
between configuration
single and triple wire
as shown in Figure 10. The main objectives of the tests were to verify the repeatability of
windings. The release times are shown in Figure 12. The system functioned well 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
mounted resistor. However,
repetitive activations the release
without electrical time was
malfunction onnoticeably varied
the circuit and in repetitive
no failure on the release
tests because of the variable workmanship of applied tension
mounted resistor. However, the release time was noticeably varied in repetitive release on the knot of the wire. The
mechanism’s
tests because ofaverage release
the variable time withof
workmanship one, two,tension
applied and three wire
on the knot windings
of the wire.in 10Therepetitive
triggers
mechanism’s for each caserelease
average were 0.77
time s, 0.89one,
with s, and
two,1.1ands, three
respectively. Thus,in
wire windings P-HRM has adequate
10 repetitive
triggers for release
repetitive each casecapability
were 0.77 s,for
0.89 s, and 1.1 s,in
application respectively.
actual space Thus, P-HRM has
missions. adequate
There is no fast re
repetitive
lease release capability
requirement, but it for application in actual
is advantageous to cut space
the missions.
wire as soonThereasis possible
no fast release
after the ini
requirement, but it is advantageous to cut the wire as soon as possible
tiation of burn wire triggering in the mechanism to prevent electrical malfunction in the after the initiation
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
wire thickness
thickness within within the qualification
the qualification temperature temperature
range of −range40 to +60 of ◦−40 to +60
C. The °C.time
release The release
time of the panel was slightly higher at low temperatures than
of the panel was slightly higher at low temperatures than that in the high-temperature that in the high-tempera
ture environment
environment becausebecause of the environmental
of the environmental effect on the effect on the
variation in variation
the heatingin theofheating
time the time
However, the release time variation from 0 to +60 ◦
of the resistor. However, the release time variation from 0 to +60 °C was nominal.
resistor. C was nominal.

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

1.5
Release time (s)

0.5

0
2 4 6 8 10
Number of tests
12.Release
Figure 12.
Figure Releasetimes of of
times thethe
proposed mechanism
proposed in repetitive
mechanism releaserelease
in repetitive functionfunction
test. test.
Relative to the space radiation environment, the radiation hardness of electronic
Relative
devices to the
is defined space radiation
according to two keyenvironment, the radiation
aspects: cumulative hardness
effects and single of electronic
event ef- de
vices is defined
fects [33]. according
The most importantto two keytoaspects:
parameter considercumulative effects
when selecting and components
electrical single 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 o
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 (60 Co) radiation source for radiation hardness assurance
of the P-HRM in a harsh space environment. The dose rate of 60 Co 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.
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Figure 13. An example of launch vibration test setup of VMLSA on a vibration shaker.
Figure 13.
Figure An example
13. An example of
of launch
launch vibration
vibration test
test setup
setup of
of VMLSA
VMLSA on
on aa vibration
vibration shaker.
shaker.

Input profile (20~500 Hz)

Input profile (20~500
st Hz)
Typical PCB (1st Eigenfreq. 48.9 Hz)
Typical PCB st (1 Eigenfreq. 48.9 Hz)
VMLSA (1st Eigenfreq. 75.0 Hz)
10 VMLSA (1 Eigenfreq. 75.0 Hz)
10

1
Acceleration (g)

1
Acceleration (g)

0.1
0.1

0.01
20 100 500
0.01 Frequecncy (Hz)
20 100 500
Figure 14. Low-level sine sweep test results in z-axis excitation.
Frequecncy (Hz)
Figure
Figure 14.
14. Low-level
Low-levelsine
sine sweep
sweep test
test results
results in
in z-axis
z-axis excitation.
excitation.
Aerospace 2021,
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24

The
The VMLSA’s
VMLSA’s corresponding
correspondingaxisaxissine
sinevibration
vibrationtest testresults
resultsduring
during the
thex-,x-,
y-,y-,
and z-
and
axis
z-axisexcitations
excitationsareareshown
shownin in
Figure
Figure15.15.
TheThequalification
qualification level of full-level
level sinusoidal
of full-level vi-
sinusoidal
bration
vibration loads
loadsforforsmall satellites,
small as preferred
satellites, as preferred by the
by theQB50 System
QB50 requirements
System requirements and and
rec-
ommendations
recommendations [34], was
[34], applied
was appliedalong
alongthe axis
the axisofofthe
thesolar
solarpanel,
panel,but
butthe
thez-axis
z-axis was
was the
most important
most importantaxis axisbecause
becauseititproduced
producedthe thehighest
highest dynamic
dynamic deflection
deflection in in comparison
comparison to
to the
the others.
others. TheThesolarsolar panel’s
panel’s corresponding
corresponding x-, and
x-, y-, y-, and z-axes
z-axes maximum
maximum acceleration
acceleration re-
responses
sponses under
under a sinusoidal
a sinusoidal vibrationinput
vibration inputloadloadofof2.52.5gginineach
eachexcitation
excitationaxis
axiswere
were 2.5,
2.5,
2.5, and
2.5, and 10.26
10.26 g,
g, respectively.
respectively. A maximum 10.26 g resonance response was observed at 75
Hz on
Hz on the z-axis of the VMLSA during the same axis excitation.
the z-axis

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

10
Acceleration (g)

0.1
10 100
Frequency (Hz)
Figure
Figure 15.
15. The
Theresults
resultsofofVMLSA’s
VMLSA’scorresponding
correspondingaxis
axissine
sinevibration
vibrationtests inin
tests the x-,x-,
the y-,y-,
and z-axis
and z-axis
excitations.
excitations.

The VMLSA’s corresponding

The corresponding axis random
random vibration
vibration test results for each excitation
axis are shown in Figure 16. The Grms
axis values in
rms values in each
each corresponding
corresponding axis calculated from
the PSD acceleration profiles
the PSD acceleration profiles under the x-, y-, and z-axis excitations were 16.57, 16.36, and
respectively. Compared
13.51, respectively. Compared to to the
the input
input level
level random
random vibration
vibrationprofile
profileof
of14.1
14.1GGrms
rms, the
Grms
G rmsofofVMLSA
VMLSAwas waslower
lowerbybyaafactor
factorofof 1.04
1.04 in
in the z-axis owing
the z-axis owing to
to the high damping
achieved by
achieved by the
the attached
attached stiffeners.
Aerospace 2021,
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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
PSD Acceleration (g 2/Hz)

0.1

0.01

0.001
100 1000
Frequency (Hz)
Figure 16. The
Figure 16. The results of VMLSA’s
results of VMLSA’scorresponding
corresponding axis
axis random
random vibration
vibration teststests in x-,
in the they-,
x-,and
y-, and z-
axis excitations.
z-axis excitations.

The results of the LLSS tests conducted before and after the full-level sinusoidal and
The results of the LLSS tests conducted before and after the full-level sinusoidal and
random vibration tests are summarized in Table 8. 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.
VMLSA
The visual was within 4.85%
inspection throughout
of the solar the test
panel module sequences,
performed afterwhich
all theisvibration
within the 5%
tests criterion.
did
The
not report dissociation and plastic deformation of the stiffeners. Thus, the structural safetytests did
visual inspection of the solar panel module performed after all the vibration
not report
of the solardissociation and plastic
panel under launch deformation
vibration loads wasof the stiffeners.
validated Thus, the structural
at the qualification level. safety
of theFurthermore,
solar panel in under launch
the middle ofvibration
the VMLSA, loads
the was validated
maximum at the
relative qualification
dynamic displace-level.
ment under the random vibration load calculated by the 3-sigma value of acceleration Grms
response
Table was 0.12
8. Results of mm. The dynamic
the VMLSA’s displacement
low-level of the
sine sweep VMLSA
(LLSS) testsisconducted
reduced by a factor
before andofafter full-
4.3 relative
level to atests.
vibration 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
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.
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24

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

6
Voltage (V)

Before vibration After TV tests: 1.5 s

test: 0.72 s
4
Burn wire
2 triggering initiation
(power supply on)

0
0 5 10 15 20
Time (s)

Figure
Figure 17.
17. Summary
Summaryof
of release
release time
time of
of the
the mechanism
mechanism at
at each
each test
test event.
event.

The solar
The solar panel
panel was
was examined
examinedmicroscopically
microscopicallyusingusingoptical
opticalmicrophotographs,
microphotographs,fol- fol-
lowing all the above-mentioned tests. The sidereal edge optical microphotographs
lowing all the above-mentioned tests. The sidereal edge optical microphotographs taken taken
on the
on the solar
solar panel
panel before
before and
and after
after the
the TV
TV tests
tests did
did not
not show
show cracks,
cracks, plastic
plastic deformation,
deformation,
ordissociation
or dissociationininthetheattached
attachedstiffeners.
stiffeners. These
These qualification-level
qualification-level tests
tests andand inspection
inspection re-
results
sults validated
validated thethe structural
structural safety
safety of the
of the VMLSAVMLSA under
under launch
launch andand in-orbit
in-orbit thermal
thermal en-
environments.
vironments. TheThe optimized
optimized design
design of of theHRM,
the HRM,simplified
simplifiedelectrical
electricalcircuit,
circuit,andand solar
solar
panel deployment experiments carried out under different test conditions
panel deployment experiments carried out under different test conditions led to confi- led to confidence
in the in
dence performance of the of
the performance deployment
the deploymentsystem. The highly
system. damped
The highly deployable
damped solar panel
deployable solar
module proposed herein is effective for guaranteeing the structural
panel module proposed herein is effective for guaranteeing the structural safety safety of solar cells
of solar
under launch environment without reducing the area of solar cell attachment
cells under launch environment without reducing the area of solar cell attachment and and assuring
reliable release
assuring reliableaction
releaseof action
the panel in apanel
of the spaceinenvironment.
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,
stiffeners and adhesive tapes. The novel three-pogo pin-based HRM has several ad-
including high loading capability in launch configuration, simpler electrical system, and
vantages, including
guaranteed solar panel high deployment
loading capability in launch
in a space configuration,
environment. simpler electrical
In addition, sys-
the radiation
tem, and guaranteed solar panel deployment in a space environment.
test results of the electrical interface PCB confirmed the radiation hardness of the P-HRM In addition, the ra-
diation test results of the electrical interface PCB confirmed the radiation
to deploy the solar panel. Further TID testing of the mechanism will be performed at a hardness of the
P-HRM
system to deploy
level the solar
of P-HRM panel.
under Further
the TID testing
estimated one-year of the
totalmechanism
ionizing dose will of
be 10
performed
krad for
at
STEP Cube Lab-II to assure radiation hardness over the mission life. Compared to a10typical
a system level of P-HRM under the estimated one-year total ionizing dose of krad
for STEP Cube Lab-II to assure radiation hardness over the mission
PCB solar panel, the relative dynamic displacement at the center of the VMLSA under life. Compared to a
typical PCB solar panel, the relative dynamic displacement at the center
random vibration loads was decreased by a factor of 4.3. The solar panel deployment tests of the VMLSA
under random
conducted aftervibration loadsand
the vibration wasTV decreased by a factor
tests validated of 4.3. Theofsolar
the reliability panel deploy-
the mechanism. In
ment tests conducted after the vibration and TV tests validated
the near future, the solar panel module flight model will be manufactured and the reliability of the mech-
validated
anism.
again by In carrying
the near future,
out launchthe solar paneland
vibration module flight
TV tests at model will be manufactured
the acceptance and
level. This design
validated again by
strategy could alsocarrying out launch vibration
be advantageous for rapid and TV tests of
attenuation at the acceptance
in-orbit level.
vibration onThis
the
design strategy could also be advantageous for rapid attenuation of
panel caused by the satellite during the slew maneuver, which could significantly reduce in-orbit vibration on
the panel caused by the satellite during the slew maneuver, which could
the performance degradation of satellites where the rapid acquisition of the target point significantly re-
duce the performance degradation of satellites where the rapid acquisition of the target
is required.
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.

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