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50th International Conference on Environmental Systems ICES-2021-19 12-15 July 2021
Copyright © 2021 Japan Aerospace Exploration Agency (JAXA)
Thermal System Design of Nano Moon Lander
OMOTENASHI with Passive Control
J. Kikuchi 0F
1 and T. Hashimoto.1F
2
Japan Aerospace Exploration Agency, Sagamihara, Kanagawa, 252-5210, Japan
and
T. Osada 2F
3
Shinwa Space Inc., Yokohama, Kanagawa, 235-0045, Japan
OMOTENASHI is a CubeSat that will be launched by a NASA SLS rocket. Its mission is
to demonstrate that a CubeSat can make a semi-hard landing on the Moon. The 6U-size
spacecraft, which weighs 12.6 kg, consists of an orbiting module, a rocket motor for
decelerating toward the Moon, and a surface probe on the landing module. The mission will
prove successful when the signal from the spacecraft is received after landing. In the mission
sequence, the heat input changes greatly in orbit and while approaching the Moon. In
addition, the temperature of the Moon cannot be predicted because it depends on whether
the surface is in the sun or shade. Tight resource constraints make it difficult to mount a
heater and a radiator in the CubeSat, so passive thermal control is a prominent design
feature of OMOTENASHI that dissolves the technical difficulty. In November 2019,
OMOTENASHI underwent a thermal vacuum test. Based on this result, the temperature
distribution of the spacecraft in orbit has been evaluated with numerical simulation. This
paper describes an overview of the design process for the thermal control system of the
OMOTENASHI spacecraft, and also describes the results of the thermal vacuum test and
the numerical simulation.
Nomenclature
α = Solar Absorption
COM = Communication Module
CVCM = Collected Volatile Condensable Materials
⁰C = Degree Celsius
DV1 = Delta-V 1
DV2 = Delta-V 2
𝜀H = Hemispheric Emissivity
G = Gravitational Acceleration
K = Kelvin
NEA = Non-Explosive Actuator
OBC = On Board Computer
OM = Orbiting Module
PCU = Power Control Unit
RCS = Reaction Control System
RM = Rocket Motor
SP = Surface Probe
TML = Total Mass Loss
W = Watt
1 Research Engineer, Research & Development Directorate, and 3-1-1 Yoshinodai, Chuo. 2 Professor, Department of Spacecraft Engineering, and 3-1-1 Yoshinodai, Chuo. 3 Research Engineer, Department Name, and 4-9-17 Yokodai, Isogo.
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I. Introduction
MOTENASHI (Outstanding MOon exploration TEchnologies demonstrated by NAno Semi-Hard Impactor) is
a CubeSat that will be launched by a first NASA SLS rocket. Its mission is to demonstrate that a CubeSat can
survive a semi-hard landing at 50 m/s or more on the Moon. Figure 1 shows the 6U-size spacecraft, which weighs
12.6 kg and consists of an orbiting module (OM), a rocket motor (RM) for deceleration on approach to the Moon,
and a surface probe (SP) on the landing module. The mission will prove successful when the SP sends its
accelerometer information back to Earth.
Figure 1. OMOTENASHI Flight Model
II. System Overview
The purpose of the OM is to transport the RM and the SP to the Moon as shown in Figure 2. Figure 3 shows a
perspective view. Power is supplied to the OM by batteries and a solar cell. The OM contains three lithium-ion
battery cells (INR 18650 MJ1) provided by the NASA/Jonson Space Center. These cells are arranged as 3 in series.
There is a thin-film solar array mounted on the side facing the Sun that generates 24 W. The communication module
(COM) of the OM uses X-band and UHF-band frequencies as the function of transmitting and receiving. The
downlink frequency is synchronized with the uplink while it is established. The COM can relay a special code from
the ground to make a range measurement as well. The three-axis attitude control is realized by XACT, which is
provided by Blue Canyon Technology (BCT). XACT has a star tracker, 3-axis gyros, three reaction wheels, and four
sun sensors attached to the outside of the spacecraft. The reaction control system of the OM uses two Micro
Propulsion System (MiPS) units provided by VACCO Industry. Each module has a propellant tank, a vapor tank,
four thrusters, valves, and pipes. The MiPS uses R-236fa propellant, which is a non-toxic, non-flammable liquid.
The purpose of the RM is to decelerate the SP as it approaches the Moon, as shown in Figure 2. The RM consists
of an aluminum motor case and HTPB/Al /APl fuel and can provide a total ΔV of 2500 m/s. The RM is ignited via
an optical fiber by a laser diode mounted on the OM.
The purpose of the SP is sending the accelerometer information back to Earth, as shown in Figure 2. Figure 3
shows a perspective view. The SP contains two primary battery cells (LM17500) provided by SAFT. These cells are
connected as 2 in series. The communication module of the SP uses the UHF-band. The SP has a shock absorption
system consisting of an airbag and crushable material. Despite the airbag and crushable materials to absorb some of
the impact, the electronics of the SP will be subjected to more than 8000 G. To improve impact tolerance, all
components are encapsulated in an epoxy resin. The details of the system design are described in references.[1]
O
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Figure 2. Appearance of each module
Figure 3. Perspective view (Left : Orbiting Module, Right : Surface Probe)
III. Mission Sequence
The mission sequence is as follows: After separating from the SLS rocket, the OMOTENASHI spacecraft
consisting of the OM, the RM, and the SP will utilize the attitude control module to initiate an attitude acquisition
sequence to direct its solar cells to face toward the Sun. Then after a 24-hour health check and tracking for orbit
determination, the spacecraft will perform a maneuver to enter lunar impact orbit (DV1) using the two modules
utilizing cold-gas jet propulsion. DV1 will be achieved at about 20 m/s. The trajectory design of DV1 ensures
landing at a site where the SP can communicate with the Earth. Several minutes before lunar impact, the spacecraft
will initiate the landing preparation sequence, an attitude-to-deceleration maneuver (DV2) and spinning in
preparation for the RM firing. To reduce the attitude error and stabilize the spin, the RM is ignited by a laser while
spinning at 5 Hz. DV2 is performed using the RM and its deceleration will be 2500 m/s. Non-Explosive Actuator
(NEA) which connects the SP+RM to the OM is activated a few tens of milliseconds before RM ignition. After this,
there is no mechanical connection between the OM and the SP+RM. Then, the SP and RM are separated from the
OM at ignition. After DV2 is completed, the SP and RM achieve a semi-hard impact on the lunar surface. The SP is
designed to survive for at least a few minutes on the lunar surface, despite its harsh thermal environment. The
mission will prove successful when the signal from the spacecraft received after the Moon landing.
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Figure 4. Mission Sequence
IV. Thermal Design with Passive Control
Tight resource constraints on CubeSat’s mass and size make it difficult to include both a heater and a radiator in
its design. Passive thermal control, a prominent feature of OMOTENASHI’s design, is the solution. Because this
method does not require power consumption or much space, it is suitable for the CubeSat mission under the tight
resource constraints. In addition, it can be adapted for use in environments with large temperature differences, such
as those that face Moon landing missions. This technique uses white paint and a graphite sheet, described in the
following section.
A. White Paint
In the mission sequence, the heat input changes greatly in
orbit and while approaching the Moon. This is because the
attitude of DV1 and DV2 varies depending on the launch date
and the separation orbit from the SLS rocket. In addition, the
surface temperature cannot be predicted because it depends on
whether it is in the sun or in shade. Regardless of the surface
conditions, the mounted components of the spacecraft must be
maintained within the operating temperature range. Therefore,
the spacecraft is coated with APTEK2711, a white paint that
has high emissivity and reflectivity. This paint can compensate
for small changes in temperature and is not much affected by
solar reflection from the Moon’s surface. In OMOTENASHI,
there is heat input only to the side facing the Sun where the
solar cell is mounted. Therefore, the other five sides are
designed only to radiate heat. The exposed side of the
spacecraft is coated with white APTEK 2711 paint as shown
in Figure 6. As a measurement result, a coating of this white
paint has solar absorption α of 0.141, and a hemispheric
emissivity εH of 0.909. The appearance of APTEK2711
provided by Aptek Laboratories Inc. are summarized in Figure 6. Figure 5. Emission Design
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B. GraphiteTIM
The COM & PCU module generates the most heat (more than 10 W) but are small (less than 0.5 kg) and have a
little heat capacity, so it is necessary to maximize the thermal conductivity in a limited space. Therefore, the thermal
conduction is increased by sandwiching a graphite sheet, GraphiteTIM provided by Panasonic Inc, between the two
radiating sides of the module and panels. The appearance of GraphiteTIM are summarized in Figure 8. This
graphite sheet has a characteristic of the compressibility and thermal resistance being different depending on the
magnitude of pressure from the out-of-plane direction. On the other hand, bonding of components and panels by
RTV S-691 was a candidate because it does not require surface pressure. However, it was not adopted because the
components could not be removed from the space craft during the development test period. Conversely, thermal
contact with the solar cell is minimized by sandwiching a titanium washer at its points of contact, as shown in Figure
7. On the other hand, the materials of a stainless steel with high stiffness and an glass epoxy resin with low thermal
conductivity were also candidates. However, as a result of trade-offs in terms of stiffness and thermal conductivity,
titanium was adopted as the sandwiching washer.
Figure 6. Appearance of APTEK2711
Figure 8. GraphiteTIM Figure 7. Radiation and Insulation Design of
COM & PCU Module
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V. Thermal Vacuum Test
A. Test Configuration
The thermal vacuum test of OMOTENASHI was conducted in November 2019 at JAXA, as shown in the inner
planetary chamber (Figure 9). This chamber has a liquid nitrogen shroud and can simulate the low-temperature
environment of space. Its temperature can reach -180⁰ C and the vacuum can drop to 10-5 Torr or less. Figure 10
shows the inside of the chamber. There is a mounting plate (A5052) for attaching the spacecraft inside the chamber.
The temperature of this plate can be adjusted from -100 ⁰C to +100 ⁰C. As external ports, 50 T-type thermocouples
and 5 Dsub connectors provide power to the spacecraft. This chamber has two V150A flanges that can be removed,
and two RF cables for spacecraft communication are connected in this test.
The test configuration is
shown in Figure 11. Six IR panels
combining a MINCO heater and
an aluminum plate were used to
simulate the heat input to the
spacecraft from the Sun. The
spacecraft was fixed to a jig via
glass epoxy blocks to insulate the
spacecraft from the mounting
plate. It should be noted that the
use of a real RM has not been
approved for handling explosives,
so a RM dummy without
explosives, but with the same
mass and shape, was used. The
temperatures of the spacecraft
were measured with thermistor
telemetry and thermocouples
attached to the spacecraft. In this
thermal vacuum test, each test
mode was set for three purposes.
The details of each mode are
summarized in Table 3.
Figure 9. Vacuum Chamber Figure 10. Inside of Chamber
Figure 11. OMOTENASHI Test Configuration
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Bake-out Test Cycle Test Correlation Test
Purpose Bake-out of non-metallic
materials
Confirmation of system
function and performance in
thermal vacuum cycle
Obtain results for correlation
of thermal simulation model
Temperature 50.56 Within operating temperature
range of each hardware
Predicted temperatures
(High & Low) in orbit
Pressure 1.33x10-3 Pa (1x10-5 Torr)
Cycle 1 : Hot 3 : Hot & Cold 1 : Hot & Cold
Duration 24 hour 8 hour Δ0.3⁰C/h or less
Shroud Heater ON 100K or less 100K or less
Spacecraft OFF ON ON
B. Test Result
The thermal vacuum test of OMOTENASHI was performed over 10 days. Figure 12 shows the temperature
measurement points of each component. The bake-out test was completed without exceeding 1.33 10-3 Pa (1 10-
5 Torr) for 24 hours. In the cycle test, an electrical check was performed at each exposure to high and low
temperatures; no malfunctions were detected. The correlation test simulated the expected high and low temperatures
in orbit. The convergence criterion that the temperature transition is Δ0.3 ⁰C/h or less was achieved in the correlation
test.
The numerical correlations were performed using these test results. Most of the correlations were the heat
exchanges between the chamber, the jig and the spacecraft, and the thermal antennas and solar sensors, which have a
low thermal coupling and are exposed to the outside. In particular, the thermal resistivity of the graphite sheet was
40 K-cm2/W, which is higher than the published value. Since the four corners of the COM & PCU module (80mm2)
are fastened with bolts, it is inferred that the contact pressure against the graphite sheet is not uniform. By finding
these correlations, the difference between the measured and the predicted temperature was kept to ± 5⁰ C for all parts.
The analysis result in orbit is given in the next section.
Table 3. Summary of Test Condition
Figure 12. Temperature Profile of Thermal Vacuum Test
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VI. Thermal Analysis in Orbit
A. Analysis Condition of Cruising Phase
Figure 13 shows the spacecraft and the layout of its internal components during the cruising phase. This is a
thermal simulation model composed by the Thermal Desktop, and each parameter such as thermal conductivity is
correlated with the results of the thermal vacuum test. The temperature in four analyses in the cruising phase were
calculated using this model as shown in Table 4.
Cases 1 and 2 were the cold and hot cases of the steady-state (pointing toward the Sun) analysis in the trans-
lunar orbit, respectively. Cases 3 and 4 were the cold and hot cases of high-speed spinning, respectively. The
transient analyses for 30 minutes were conducted in the Moon-approach orbit. In case 3, it was assumed that the Sun
was hidden by the Moon. The spacecraft only received the radiation from the dark side of the Moon. The spacecraft
can survive within the temperature range even if the eclipse continues for more than 30 minutes . However, in terms
of the power consumption, trajectories with solar eclipse of 30 minutes or more before landing will be avoided. In
case 4, it was assumed that the spacecraft had heat input from the Sun, the lunar radiation, and sunshine reflected
from the Moon. The power consumption of each component (COM & PCU, XACT, OBC, RCS, Battery) of the OM
in each case is also shown in Table 4. Considering the thermal dissipation, a factor of 0.9 and 1.2 is applied to the
power consumption of electronics for cold and hot cases, respectively, against the nominal power consumption.
After SLS rocket deployment, the sun pointing is expected to be completed nominally in 10 minutes. Analysis
has confirmed that no components will exceed the operating temperature range, even in case of the lowest initial
temperature at SLS rocket deployment. Moreover, DV1 will be executed during the cruising phase. Due to a power
constraint on the OM battery module, RCS can be activated for a short time and returned to the sun-pointing attitude
for battery charging. Several sets of this RCS activating and charging cycle must be repeated. Analysis has
confirmed that no components will exceed the operating temperature range in these operations. Therefore, analysis
results are not described after the deployment and the DV1 sequence.
Analysis Case 1 2 3 4
Phase Cruising between Earth and Moon Moon Approach
Condition Cold Hot Cold Hot
Calculation Steady State Transient 1800[sec]
Initial Temperature - - Analysis Case 1 Analysis Case 2
Solar Radiation[W/m^2] 1322.0 1412.0 - 1412.0
Moon IR[W/m^2] - - 5.2 1265.0
Moon Albedo - - - 0.11
Spacecraft Attitude Sun Pointing Sun Pointing Spinning Spinning
Power
Consumption
[W]
COM & PCU(OM) 12.83 17.11 14.64 19.52
XACT(OM) 2.25 3.00 0.00 0.00
OBC(OM) 1.63 2.17 3.39 4.52
RCS(OM) 0.00 0.00 1.08 1.44
Battery(OM) 0.00(Charge) 0.00(Charge) 1.77(Discharge) 2.36(Discharge)
Figure 13. Thermal Analysis Model in Cruising Phase
Table 4. Analysis Condition in Cruising Phase
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B. Analysis Result of Cruising Phase
In this mission, the temperature margin on the hottest side is more constrained, so the results of this are given as
a representative example of the analysis.
Figure 14 shows the temperature distribution over the spacecraft in case 2, which is the steady-state analysis. In
addition, Table 5 shows the operating temperature range and analysis results of each component in cases 1 and 2.
The analysis confirms that, although the COM & PCU reach the highest temperature, the heat in each component
can be radiated without exceeding the maximum operating temperature of 60⁰ C. Besides, the temperature of the RM
+ SP tends to be low because the RM+SP are connected to the OM module by four bolts, with minimal surface
contact. Therefore, the thermal coupling between the OM and the RM+SP is small.
Figure 15 shows the transition condition of the Moon altitude of the spacecraft in Cases 3 and 4. The spacecraft
transits the Moon approach phase from an altitude of about 2000 km, and then descends to an altitude of about 23
km in 30 minutes. Figure 16 shows the temperature profile for each component in case 4 (approaching the Moon)
with the 30 minutes of spin. Although the temperature of all components rises, no components are subjected to a
rapid temperature rise or exceed the maximum operating temperature range. This result confirms that there are no
components that will exceed their operating temperatures and each component has a temperature margin of 10°C or
more during the cruising phase.
Analysis Case 1 2
Phase Operating
Temp
Range
Cruising between
Earth and Moon
Condition Cold Hot
Predicted
Temp
[°C]
COM&PCU(OM) -20~60 38.1 49.3
XACT(OM) -22~60 21.9 30.5
OBC(OM) -20~85 24.2 33.2
RCS+X(OM) -22~55 23.9 32.8
RCS-X(OM) -22~55 15.8 23.0
Battery (OM) 0~45 9.5 20.2
Figure 14. Analysis Result in Case 2
Table 5. Analysis Result of Components
Figure 16. Temperature Profile in Case 4 Figure 15. Moon altitude of Spacecraft in Case 3, 4
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C. Analysis Condition of Landing Phase
Figure 17 shows the RM+SP and the SP alone during the landing phase. There is also a thermal simulation
model composed by the Thermal Desktop, and each parameter such as thermal conductivity is correlated with the
results of the thermal vacuum test. The temperatures in four analyses dealing with the landing phase in Table 6 were
calculated using these models.
Cases 5 and 6 were transient analyses of the cold and hot cases, the RM had been ignited under a high spin and
the RM+SP had separated from the OM, respectively. It was assumed that the RM was continuously burning until
landing, for 18.9 seconds. The SP is connected to the other side of the RM nozzle with four bolts. From the RM
combustion test results, the time profile of the temperature of the RM top was used for the analysis as the boundary
condition with the SP.
Cases 7 and 8 were the cold and hot cases of the analysis which assumed that the SP had landed on the Moon, The
RM had been removed, and the SP was in contact with the Moon’s surface with the side having the highest heat
conductivity. The lunar surface does not simulate the regolith, but rather a flat plate. This is the worst case (i.e., a=
large heat conduction) of these analyses because both the heat input for the hot case and the heat outflow for cold are
expected to be at their maximum values. The surface temperature is -170 °C when the Moon is facing away from the
Sun (in the shade) in case 7. On the other hand, the surface temperature of the moon is 113 °C when the Moon is
facing the Sun in case 8. The power consumption of each component (COM, OBC, Accelerometer)of the SP in each
case is also shown in Table 6. Considering the thermal dissipation, a factor of 0.9 and 1.2 is applied to the power
consumption of electronics for cold and hot cases, respectively, against the nominal power consumption.
Analysis Case 5 6 7 8
Phase Separation from Orbiting Module Landing on Moon
Condition Cold Hot Cold Hot
Calculation Transient 18.9[sec] Transient 300[sec]
Initial Temperature Analysis Case 3 Analysis Case 4 Analysis Case 5 Analysis Case 6
Solar Radiation[W/m^2] - 1412.0 - 1412.0
Moon IR[W/m^2] 5.2 1265.0 5.2 1265.0
Moon Albedo - 0.11 - 0.11
Spacecraft Attitude Spinning Spinning Landing Landing
Power
Consumption
[W]
COM(SP) 4.05 5.40 4.05 5.40
OBC(SP) 1.24 1.66 1.24 1.66
Accelerometer(SP) 0.11 0.14 0.11 0.14
Figure 17. Thermal Analysis Model in Landing Phase (Left:RM+SP, Right:SP)
Table 6. Analysis Condition in Landing Phase
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D. Analysis Result of Landing Phase
In this mission, the temperature margin on the hottest side is more constrained, so the results of this are given as
a representative example of the analysis.
Figure 18 shows the transition condition of the Moon altitude of the SP+RM in Case 5 and 6. The SP+RM transits
the Moon approach phase from an altitude of about 23 km, and then descends to the Moon surface in 18.9 seconds.
Figure 19 shows the temperature profile for each component in case 6 when the RM continues to burn and the
RM+SP are separated from the OM. It confirms that the temperature rise of the RM is quite rapid at the start of
ignition. On the other hand, the temperature rise of the SP components is not as sharp as that of the RM because the
SP is connected to the RM by four bolts and crushable material that is installed to absorb impact shock. Therefore, it
is considered that the SP is not greatly affected by the temperature of the RM. These four bolts are not expected to
withstand the impact at landing and will break. Then the SP and RM will be separated. The SP and electronic
devices are designed and verified to withstand the large impact of 8500G by hardening with epoxy resin. Even if the
SP separates and collides with the RM, the SP will not be damaged.
Figure 20 shows the temperature distribution over the SP 5 minutes after the Moon landing in case 8. In addition,
Table 7 shows the operating temperature range and analysis results of each component in cases 7 and 8. In this
mission, the success criterion is the confirmation of a successful landing on the Moon, that is, the signal from the SP
is received just after the landing. Therefore, there are no requirements on the SP's lifetime. After 5 minutes, the
temperature of all SP components rose greatly, affected by the Moon environment, and exceeded the maximum
operating temperature. This tendency is the same as in case 7, which is for a cold environment. It confirms that the
survival time of the SP on the Moon will be very short.
Analysis Case 7 8
Phase Operating
Temp
Range
Landing
on Moon
Condition Cold Hot
Predicted
Temp [°C]
COM(SP) -20~60 -75.7 71.8
OBC(SP) -20~85 -80.8 72.9
Accelemeter(SP) -55~175 -74.8 70.2
Figure 20. Analysis Result in Case 8
Table 7. Analysis Result of Components
Figure 19. Temperature Profile in Case 6 Figure 18. Moon altitude of Spacecraft in Case 5, 6
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Conclusion
This study covers the passive thermal control that is incorporated in the thermal design of CubeSat under tight
constraints. The major features are the white paint that reduces the effect of the external environment and the
graphite sheet that conducts heat with high efficiency. The thermal vacuum test and in-orbit analysis verify this
concept and show that the spacecraft can be kept within the operating temperature range while in orbit.
All the environment testing of the OMOTENASHI using the flight model has been completed, and the mission’s
feasibility has been confirmed. Currently, the software of the flight model for the orbital operation is being evaluated.
Acknowledgments
We would like to express our gratitude to our JAXA colleagues, Shinwa Space Inc., and Mitsubishi Space
Software Co.,Ltd who participated in the experiments for helping in interpreting the significance of the results of
this study.
References
[1] Hashimoto, T.., et al.: Nano Semihard Moon Lander: OMOTENASHI, IEEE Aerospace & Electronics Systems Magazine,
Volume 34, Issue 9, pp20-30, September 2019
[2]Kikuchi, J., et al.: On-orbit Separation and Semi-Hard Landing Mechanism of Nano Moon Lander OMOTENASHI,
SciTech 2020, Florida, USA, 2020.1
[3]Campagnola S, et al., Mission Analysis for EQUULEUS and OMOTENASHI, 31st International Symposium on Space
Technologies and Science, 2017-f-044, Matsuyama, Japan, 2017.
[4]Otsuki, M, et al., Study of Impact Attenuation Device for Planetary Small Lander, Proceedings of the Space Sciences and
Technology Conference, 2018 (in Japanese)
[5]Morishita N, et al., Ways to use of super-small solid rocket motors for deep space exploration missions, Proceedings of the
Space Sciences and Technology Conference, 2018 (in Japanese)
[6]Funase, R., et al.: Flight Model Design and Development Status of the Earth - Moon Lagrange Point Exploration CubeSat
EQUULEUS Onboard SLS EM-1, Small Satellite Conference, Utah, 2018
[7]Funase, R., et al.: 50kg-class Deep Space Exploration Technology Demonstration Micro-spacecraft PROCYON, Small
Satellite Conference, Utah, 2014
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