thermal system design of nano moon lander omotenashi with

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.

International Conference on Environmental Systems

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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

thermal system design of nano moon lander omotenashi with

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