ICES-2020-282

Flight Model Thermal Design and Verification for the 6U Deep Space CubeSat EQUULEUS

Toshihiro Shibukawa1, Shuhei Matsushita1, Keidai Iiyama1, Akihiro Ishikawa2, Keita Nishii1, and Ryu Funase3. Intelligent Space Systems Laboratory (ISSL), The University of Tokyo, Bunkyo-ku, Tokyo, 113-8656, Japan

EQUULEUS is a 6U CubeSat developed by the University of Tokyo and JAXA, which will fly to a libration orbit around the second - (EML2) as a of the (SLS). At the destination and on the way there, EQUULEUS will conduct three scientific missions such as the detection of lunar impact flashes. To realize these missions, the thermal design of EQUULEUS faces constraints coupled with other subsystems. The Engineering Model (EM) was designed to clear these constraints, but after verification by a thermal vacuum test, one mission component did not meet the temperature requirements. To clear the problems in the EM, changes in operation strategies, temperature range management, and thermal interface materials were made. These design changes were verified by another thermal vacuum test, and we proved that EQUULEUS can meet the constraints to reach EML2 and conduct its scientific missions.

Nomenclature 퐶 = heat conductance 푄푉퐻 = heater power at the vaporization chamber Abbreviations of component names 푄ℎ푒푎푡푒푟 = wattage of the test heater 퐴퐷퐶푆 = the attitude determination and control 푄 = generated heat at a component system 푆 = area 퐵퐴푇 = the battery 푇 = temperature 퐶푂푀푀 = the communication module 휖 = surface absorptance 푂퐵퐶 = the on-board computer 푃퐶푈 = the power control unit Subscripts 푃푅푂푃 퐼퐹 퐵퐷 = the propulsion interface board 푃 = panels 푆퐴퐷퐴 = the solar array distribution unit 퐵푈푆 = bus components 푇퐴푁퐾 = the water tank

I. Introduction QUULEUS (EQUilibriUm Lunar-Earth point 6U Spacecraft) is one of the 13 secondary payloads delivered into E deep space by NASA’s Space Launch System (SLS) as a part of Artemis 1 mission1. The spacecraft is being developed by the University of Tokyo and JAXA. After released by SLS, the spacecraft will fly to a libration orbit around the second Earth-Moon Lagrange point 2 (EML2) and demonstrate the trajectory control techniques such as low-energy transfers utilizing weak stability regions within the -Earth-Moon region for the first time by a nano- spacecraft. EQUULEUS also has three observation instruments for scientific missions. The first instrument is named PHOENIX (Plasmaspheric Helium ion Observation by Enhanced New Imager in eXtreme ultraviolet), which is an imager to visualize the Earth’s in extreme UV wavelength. By observation from the EML2 libration orbit, PHOENIX will be able to take an overall image of the Earth’s plasmasphere. The second is named DELPHINUS (DEtection camera for Lunar impact PHenomena IN 6U Spacecraft), which is a camera to detect lunar impact flashes. Also by observation from the libration orbit, DELPHINUS will observe the impact flashes at the far side of the moon for the first time. The third is named CLOTH (Cis-Lunar Object detector within Thermal insulation), which is a dust

1 Graduate Student, Department of Aeronautics and Astronautics, 7-3-1 Hongo, Bunkyo-ku, Tokyo 2 Researcher, Department of Aeronautics and Astronautics, 7-3-1 Hongo, Bunkyo-ku, Tokyo 3 Associate Professor, Department of Aeronautics and Astronautics, 7-3-1 Hongo, Bunkyo-ku, Tokyo

Copyright © 2020 The University of Tokyo detector within Multi-Layer Insulations (MLIs). CLOTH will detect and evaluate the meteoroid impact flux at the libration orbit and on the way there, to understand the size and spatial distribution within the cis-lunar region. Figure 1 shows one example of the designed libration orbit around EML2. The diameter of this orbit around the moon is about 80,000km2, and the altitude from the moon surface is around 40,000km. From this orbit, the spacecraft can observe both the moon surface and the Earth’s plasmasphere with only a slight attitude control. However, as the moon and EML2 orbit around the Earth, the sun direction towards EQUULEUS changes gradually. Therefore, the temperature of the instruments must be kept in operational range considering the gradual change of the thermal environment. Another feature of EQUULEUS is that it adopts a cold gas propulsion system, named AQUARIUS (AQUA ResIstojet Figure 1. A Close-up of one libration orbit for propUlsion System)3. AQUARIUS utilizes water as its observation phase of EQUULEUS2. Illustrates the propellant, which is environmentally friendly and safe to use. approach, the libration orbit, and the position of the However, the large latent heat required for evaporation places moon in the Earth-Moon rotating frame a burden on the small power budget and thermal design of the nano-spacecraft. Also, the temperature of the propellant tank must be controlled so that the water inside does not freeze. From the viewpoints of operation and thermal design mentioned in the later part of this paper, we adopted the configuration shown in Figure 2 and Figure 3 for EQUULEUS. The size of the spacecraft is 6U (10cm x 20cm x 30cm), to meet the requirements of the SLS secondary payload. The observation direction of the two instuments PHOENIX and DELPHINUS are fixed to the +Y direction of the spacecraft, and the solar array paddles (SAP) are arranged to be rotatable around the X axis using a gimbal mechanism. This configuration makes it possible for the spacecraft to make observations and generate adequate power simultaneously. Also, to reuse the waste heat of the communication module (COMM) to evaporate water for the propulsion system, COMM and the vaporization chamber (VAP) are placed next to each other. In this paper, the concept of the thermal design of EQUULEUS by Koshiro et al. 4, which considers the explained constraints, is first introduced in detail. Then, the thermal design of the Engineering Model (EM) of EQUULEUS and the results of the EM thermal vacuum tests by Matsushita et al. 5 is explained. The results of this thermal vacuum test displayed some problems, which were reflected to the Flight Model (FM) thermal design. A thermal vacuum test for the Flight Model verified that the final thermal design of EQUULEUS can meet the complicated requirements to reach and conduct scientific missions at EML2. Finally, conclusions and lessons learned from this complicated CubeSat thermal design are summarized.

Figure 2. External view of EQUULEUS and axis Figure 3. Internal Configuration of EQUULEUS. The definition. The observation directions of DELPHINUS vaporization chamber is placed between the two and PHOENIX are both +Y, and the Solar Array Paddles communication modules. are rotatable around the X axis.

2 International Conference on Environmental Systems II. Concept of EQUULEUS Thermal Design

A. Thermal Environment of EQUULEUS Mission As already mentioned, the thermal design of EQUULEUS faces several constraints coupled with other subsystems. The first constraint is coupled with the orbit design and the uncertainty of the launch date. Since EQUULEUS is a secondary payload of Artemis 1 mission, we cannot decide the launch date. In fact, Artemis 1 mission has already rescheduled the launch date several times. The change in launch date affects the transfer orbit design to EML2 and the timing of thrusting phases. This also changes the sun direction at these thrusting phases, which is critical since large power is needed. To successfully reach EML2, the thermal design must let all of the thrusuting phases successful under any sun direction condition. The second constraint is coupled with the orbital design of the libration orbit around EML2 during observation phase. EQUULEUS will stay at this libration orbit for about six months, and the sun direction periodically changes along with the moon orbital motion during this phase. Figure 4 describes this configuration, and shows that four panels out of six (+Y, -Y, +Z, -Z) experiences sunlight during observation phase, and the temperature of every component must be kept within operational range regardless of this sunlight direction. The final constraint is coupled with AQUARIUS, the water propulsion system for trajectory control and angular momentum unloading. We must mainly consider two factors for using water as propellant: freezing and large latent heat. To keep the water from freezing, the water tank must be kept over 0°C at all time, including storage and right after rocket release. Moreover, since water requires large amount of latent heat for vaporization, EQUULEUS must continuously provide adequate heat to the vaporization chamber during thrusting phases. The small power budget and the complexity of thermal design for makes this requirement very difficult. Since the first delta-V phase (DV-1) requires the largest thrust throughout the entire mission, the required heater power for evaporation is the largest, and therefore we must make the largest consideration.

B. Thermal Design Concepts Firstly, considering the changes in sun direction mentioned, we designed the product of the surface absorptance 훼 and the panel area 푆 to be identical for the four panels that will face the sun. This concept makes the heat input from solar radiation identical for each panel, which makes the thermal design simple, and also declines the thermal sensitivity to the uncertainty of launch date. To decide this value of the producnt of 훼 and 푆, an one-node simulation was conducted. The result of this simulation is shown in Figure 5. The red line indicates that lower the 훼 value is, the lower power is required at the DV-1 phase. Secondly, considering the large latent heat required at the thrusting phase, the heat conductances between components around the vaporization chamber should be optimized to reuse waste heat of other components and reduce

Figure 4. The sun direction and the attitude of Figure 5. The result of the one-node simulation. The EQUULEUS during observation phase. The sun blue curve illustrates the temperature at the coldest direction seen from the spacecraft gradually changes condition and the red line illustrates the required power due to the orbital motion of the moon. at the DV-1 phase.

3 International Conference on Environmental Systems Table 1. Results of optimization using the three-node model simulation. Heat conductance between VAP and COMM is large. Parameter Value [W/K]

퐶1 5.3 퐶2 0

퐶3 0.24

퐶4 0

퐶5 10

퐶6 0

Figure 6. Three-node thermal model of EQUULEUS. Illustrates BUS, COMM, and VAP nodes, and the panel node treated as the boundary condition. heater wattage. To evaluate this concept, we introduced a three-node thermal model shown in Figure 6, consisted of the vaporization chamber node (VAP), the communication module node (COMM), and a node for all other components (BUS). Also, a panel node was introduced as a boundary. As mentioned, VAP requires a large amount ® of heat during thrusting phases. On the other hand, COMM Figure 7. Thermal Desktop model of EQUULEUS. has the largest internal power generation, and has a large All components in EQUULEUS are modeled as solid operational temperature range. The heat conductances blocks, surfaces or cylinders. between each node was optimized using the total heater wattage as the objective function. The result of this optimization is shown in Table 1, which shows that VAP and COMM should be strongly coupled to each other thermally, while both of them insulated from the BUS and the panel. This result lead to the configuration shown in Figure 3, where VAP and COMM are placed next to each other. Finally, a Thermal Desktop® model of EQUULEUS was created as a sophisticated thermal model, based on the results of the previous two simulations. This Thermal Desktop® model is shown in Figure 7, which is consisted of over 250 nodes. Simulations were conducted in the hottest and the coldest condition the spacecraft will experience, shown in Table 2. The simulation results are shown in Table 3, and demonstrate that all components are within operational range. However, mission components have a margin below 10°C, due to their narrow operational range. Moreover, simulations were conducted in the DV-1 phase to estimate the required heater wattage at the vaporization chamber and the total required power of the spacecraft. These simulations showed that the maximum total required power is 46.8W, which is below the least generable power by the SAP (48.5W). For more details of the thermal design concepts and simualtions explained in this section, we refer the reader to Koshiro et al. 4 and Matsushita et al. 5. These results proved our thermal design to be effective, and we started the development of the Engineering Model based on this design.

Table 2. Thermal Environment of EQUULEUS during the observation phase5. The most thermally severe environments at EML2 to conduct scientific observations The distance b/w EQUULEUS The number of panels Internal heat generation 푄 and the Sun 푑 𝑖푛 facing the sun 퐻 퐻 푄 = 31.3 W The hottest 푑 = 0.975 AU 𝑖푛 (DELPHINUS & CLOTH Two panels environment ⋅퐻 퐻2 2 (푃푠/푑 = 1437 W/m ) Observation mode)

퐶 퐶 The coldest 푑 = 1.025 AU 푄𝑖푛 = 26.4 W 퐶 퐶2 2 One panel environment ⋅ (푃푠/푑 = 1300 W/m ) (3-axis attitude control mode)

4 International Conference on Environmental Systems Table 3. Results of Thermal Desktop® Engineering Model simulations in hottest and coldest conditions before thermal vacuum test modification. All components meet the operation temperature requirements. We aim to have a margin of over 10°C for both boundaries of all components. The boundary of DELPHINUS in the coldest condition is -30°C because DELPHINUS is not in operation in the coldest condition. Component Temp. Boundary [°C] Simulation Results[°C] Margin [°C] Coldest Hottest Coldest Hottest Coldest Hottest SADA -30 70 1 33 31 37 BAT 0 45 10 31 10 14 PCU -20 60 8 42 28 18 COMM -20 60 16 46 36 14 OBC -20 60 7 40 27 20 PROP IF BD -40 85 8 38 48 47 ADCS -30 50 6 38 36 12 PHOENIX -30 50 0 41 30 9 DELPHINUS -30 40 0 36 30 4 CLOTH -50 90 -46 51 4 39 TANK 5 60 15 30 10 30

III. Engineering Model Thermal Design Implementation and Verification

A. Implementation After the thermal design was created and verified, the Engineering Model of EQUULEUS was developed. In order to accurately implement the thermal design into the Engineering Model, three thermal control devices were mainly used: Multi-Layer Insulation (MLI), thermal tapes, and thermal interface materials called Thermal Fillers. MLIs and thermal tapes are used to control the optical properties of the panels, and Thermal Fillers are used to control the thermal conductances between components. In EQUULEUS, MLIs are adopted in two panels , +Y and –Y. Both MLIs include PVDF sensors for the CLOTH science mission, which couuts dust particles that hit the MLI. To conduct tests easily for CLOTH, the MLI can be easily detached from and reattached to the spacecraft by hook-and-loop fasteners. For areas of the panels without MLI, three types of thermal tapes were used to control the thermal properties: Black Kapton tapes, Aluminized Kapton tapes, and Ag Teflon tapes. The ratio of these tapes were decided for each area of the panels to meet the thermal model created by Thermal Desktop®. Considering heat conductance, in conventional large scale spacecraft, the temperature of components are controlled by either thermally coupling or insulating the components roughly. However, in CubeSats such as EQUULEUS, thermal design is coupled over multiple components and even other subsystems due to its small size. To realize this thermal design, heat conductance must be controlled precisely. In EQUULEUS, we utilized thermal interface materials that we call Thermal Fillers in critical places, such as COMM to VAP and SADA to +Z panel. In the Engineering Model, a product called CHO-THERM® 1 manufactured by Taiyo Wire Cloth Co., Ltd. was used.

B. Verification of Engineering Model After we completed the development of EQUULEUS EM, a thermal vacuum test was conducted to verify the thermal design and its implementation. We aimed to simulate the hottest and coldest environments using infrared heating panels, and also conducted additional thermal correlation modes to estimate the heat conductances around critical components, such as the battery, the water tank, and the vaporization chamber. The configuration of this thermal vacuum test is shown in Figure 8. Using a Thermal Desktop® model shown in Figure 9, which represents the thermal vacuum test configuration, we modified the Thermal Desktop® model to keep the temperature difference between the thermal vacuum test results and the Thermal Desktop® simulation results within ±5°C for all components. Using this modified Thermal Desktop® model, we again conducted simulations in the hottest and coldest environments. The results of these simulations are shown in Table 4. This table demonstrates that the temperature of almost all components is within operational range. However, some problems emerged in this verification process. First, one component, DELPHINUS, exceeded the higher temperature boundary. This was because the thermal conductances

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Figure 8. Actual configuration of the thermal vacuum Figure 9. Configuration model of the thermal vacuum test of EQUULEUS EM. Shows the EQUULEUS EM, test of EQUULEUS EM. Shows the Thermal Desktop® infrared heating panels and electric cable harnesses. model for simulation and correlation

Table 4. Results of Thermal Desktop® Engineering Model simulations after thermal vacuum test modification. DELPHINUS exceeds the temperature boundary in the hottest condition. Component Temp. Boundary [°C] Simulation Results[°C] Margin [°C] Coldest Hottest Coldest Hottest Coldest Hottest SADA -30 70 -6 36 24 34 BAT 0 45 10 31 10 14 PCU -20 60 6 44 26 16 COMM -20 60 12 47 32 13 OBC -20 60 4 41 24 19 PROP IF BD -40 85 8 44 48 41 ADCS -30 50 0 39 30 11 PHOENIX -30 50 -3 43 27 7 DELPHINUS -30 40 -3 43 27 -3 CLOTH -50 90 -46 52 4 38 TANK 5 60 15 38 10 22

inside and around DELPHINUS were smaller than expected in the thermal design before modification. Also, some places that utilized CHO-THERM® 1 displayed smaller heat conductance than the designed value. Finally, since the SAP was not included in the thermal vacuum test configuration, we were not able to evaluate the radiation heat exchange between the SAP and the spacecraft. These results required us to refine the thermal design in the Flight Model (FM). Also, we conducted a decompression test and a vibration test to check that the spacecraft does not exceed the dynamic envelope of the CubeSat deployment system of SLS during launch sequence. The results of these tests showed that the inflation of the MLI is quite large during launch sequence, and has a possibility of violating the dynamic envelope requirement. Therefore, we were required to change the manufacturing method of the MLI to keep it from inflating while keeping its thermal insulation ability and its function as a scientific instrument.

IV. Flight Model Thermal Design Refinement and Verification

A. Design Changes in Flight Model To tackle the problems that emerged in the Engineering Model verifiacation, we made some design changes in the Flight Model. First, for the high temperature problem in DELPHINUS, we took methods based on temperature range management and spacecraft operation, rather than changing the actual thermal design. As shown in Figure 10,

6 International Conference on Environmental Systems Table 5. Operation and storage temperature of the elements of DELPHINUS. In the Engineering Model, DELPHINUS was treated as one component altogether with the operation temperature of -10°C to 40°C and storage temperature of -30°C to 70°C.

Operation Storage Element Temp. [°C] Temp. [°C] Low High Low High FPGA -30 70 -30 70 Board Sensor Board + -10 40 -30 70 Mirror Figure 10. Appearance of DELPHINUS. It can be Camera -10 50 -30 80 decomposed into three elements: the FPGA board, the Cover camera cover, and the mirror with the sensor board. DELPHINUS can be decomposed into three elements: the FPGA board, the cover of the camera, and the mirror with the sensor board. In the Engineering Model, the temperature ranges for these three elements were set the same, treating them as one component. However, as shown in Table 5, the FPGA board had a lighter temperature restriction than the other two elements. Therefore, the temperature ranges were reviewed and set individually for the three elements in the Flight Model. This eased the thermal design requirements for DELPHINUS. Also, the operation strategy of the DELPHINUS observation mode, which is the power mode that DELPHINUS has a possibility of exceeding its operational temperature range, was reviewed. In particular, the power mode of the communication module was reconsidered. The communication module of EQUULEUS has two power modes: HIGH mode with power consumption of 12W, for command uplink, housekeeping data downlink and mission data downlink, and LOW mode with power consumption of 8.1W, for housekeeping data downlink only. In the Engineering Model, the power mode of the communication module was always set to HIGH mode during DELPHINUS observation. However, when the spacecraft is not visible from Earth, mission data can not be downlinked, and therefore the power mode can be set to LOW mode. When the spacecraft is visible, the temperature of DELPHINUS can be monitored by the housekeeping data, and we can change the power mode by command uplink. Based on this operation strategy, the power mode of the communication module in the hottest condition can be changed to LOW mode, and the temperature of DELPHINUS can be lowered. For CHO-THERM® 1 displaying lower heat conducatance than expected, we found that we must properly manage the area, thickness, and contact pressure when using Thermal Fillers. By calculating from the contact surface area and screw torque, we identified that the contact pressure applied on CHO-THERM® 1 was not within the recommended range for the places that showed low heat conductance. Therefore, we first managed the pressure and thickness for each place with Thermal Fillers, using glass epoxy spacers. Second, we searched for other Thermal Filler candidates that have different pressure requirements, such as TpliTM 210 manufactured by Laird Technologies, Inc. and λGEL® COH-4000LVC manufactured by Taica Corporation. Both of these Thermal Fillers have flight heritage, and the outgas rate fulfilled requirements. The appearance of the three Thermal Fillers are shown in Figure 11, and their characteristics are shown in Table 6. Before actually using the new Thermal Fillers in the Flight Model, we conducted a test to check the performance of each Thermal Filler candidate. The configuration of this test is shown in Figure 12 and 13. We used each Thermal Filler candidate between two test pieces, and replicated the contact pressure of the Table 6. Characteristics of the Thermal Fillers. Each Thermal Filler has different pressure ranges. Pressure Material Product Name Range [MPa] Type CHO- 2.07 - 3.45 Sheet THERM®1 λGEL® COH- TM 4000LVC Tpli 210 0.1 - 0.7 Fiber λGEL® COH- - 0.42 Gel Tpli™ 210 CHO-THERM® 1 4000LVC Gap Filler

Figure 11. Thermal Fillers used in EQUULEUS FM. 7 International Conference on Environmental Systems Thermal Filler Test Piece 1 Test Piece 2 푇1 푇2

Heater Spacer 푄ℎ푒푎푡푒푟

Figure 12. Configuration of the Thermal Filler test. The conductance between the two test pieces can be calculated from the heater imput 푸 and the temperature 풉풆풂풕풆풓 Figure 13. Image of the Thermal Filler test. difference (푻ퟐ − 푻ퟏ). places that they would be used. By turning on a heater placed on one of the pieces and measuring the temperature difference of the two pieces after equilibrium, the heat conductance of each Thermal Filler candidate was evaluated. The results of this test demonstrated that CHO-THERM® 1 actually displays lower heat conductance than expected if the pressure is out of its recommended range. It was also demonstrated that these new Thermal Filler candidates such as TpliTM 210 can be used as alternates of CHO-THERM® 1. Following these results, new Thermal Fillers were adopted in the Flight Model. Finally, to reduce the inflation of the MLI, we created 10 test pieces of MLI with different manufacturing methods, and conducted decompression tests and measured the inflation for each test piece. Using the results, we decided the best manufacturing method and manufactured the MLI for the Flight Model. These design changes were reflected to the Thermal Desktop model, and again simulations were conducted in the hottest and the coldest condition, to validate the design change. The results are shown in Table 7, and all components are within the operation temperature. Especially, the change in operation strategy lowered the temperature of the entire spacecraft in the hottest condition, cleared the problem in DELPHINUS, and created large margins for several components.

B. Verification of Flight Model After we finished the development of the Flight Model, we conducted another thermal vaccum test for verification. We included the Engineering Model of the SAP in the configuration of this test, as shown in Figure 14, to evaluate the radiation heat exchange between the SAP and the spacecraft. We placed heaters on one side of the SAP, to simulate

Table 7. Results of Thermal Desktop® Flight Model simulations before thermal vacuum test modification. All components meet the operation temperature requirements due to design changes. Component Temp. Boundary [°C] Simulation Results[°C] Margin [°C] Coldest Hottest Coldest Hottest Coldest Hottest SADA -30 70 -3 28 27 42 BAT 0 45 10 24 10 21 PCU -20 60 8 47 28 13 COMM -20 60 15 30 35 30 OBC -20 60 7 36 27 24 PROP IF BD -40 85 11 31 51 54 ADCS -30 50 5 31 35 19 PHOENIX -30 50 -3 37 27 13 DELPHINUS -30 70 7 37 37 33 FPGA BD DELPHINUS -30 40 -3 37 27 3 CAM+SENSOR CLOTH -50 90 -43 48 7 42 TANK 5 60 15 25 10 35

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Figure 14. Actual configuration of the thermal Figure 15. Configuration model of the thermal vacuum test of EQUULEUS FM. Shows the cage, vacuum test of EQUULEUS FM. Shows the Thermal infrared heating panels, and the SAP. EQUULEUS FM Desktop® model for simulation and correlation is behind the heating panels. its temperature in the hottest condition and the coldest condition. We again modified the Thermal Desktop® model to keep the temperature difference between the thermal vacuum test results and the Thermal Desktop® simulation results of the model shown in Figure 15 within ±5°C for all components. However, we had some problems during this modification. Firstly, the panels of the spacecraft had an unexpected temperature distribution. In particular, areas of the panels that were not covered by MLI tended to show lower temperature in the measured value than the analyzed value. Also, temperature differences over 10°C were measured within the same panel, which was not expected since each panel was made of one piece. These problems emerged because in the Engineering Model thermal vacuum test, restrictions of the number of thermocouples allowed us to place zero or one thermocouple per panel, whereas we were able to place two or three per panel in the Flight Model test. One reason for this difference between the measurement and the analysis is because we were not able to implement holes, bumps, and uneven thicknesses of the panels into the Thermal Desktop® model. These uneven factors may have created a large temperature distribution and hot spots within one panel. For modification of the Thermal Desktop® model of the panels, we mainly focused on adjusting the temperature of the flight thermometers which will be used in actual operation of the spacecraft, rather than the thermocouples used only for the thermal vacuum test. Secondly, we had difficulty in modifying the thermal model of the SAP. When we tried to raise the temperature up to the hottest condition, larger heater wattage was required than expected in the analysis before the test, which meant that there was a heat leak that we did not recognize. We assumed this heat leak to be from the harnesses of the heaters and thermocouples attached to the SAP, and implemented this leak into the model by connecting the SAP and the vacuum chamber by a very small heat conductance. However, there was not a appropriate value for this conductance that satisfied both the hottest condition and coldest condition. Alternatively, we set the SAP as a boundary node in the thermal model to modify the rest of the spacecraft. Therefore, we were not able to completely verify the heat exchange between the SAP and the spacecraft. Although we had some problems, the modified Thermal Desktop® model for the Flight Model was created. The results of the simulations in the hottest and coldest environments for this model are shown in Table 8. This table demonstrates that all components except DELPHINUS sensor board has a margin of over 10°C for both high and low boundaries. However, the upper boundary of DELPHINUS sensor board has no margin. The reason why we were not able to lower the temperature of DELPHINUS in the worst hottest condition is because the hardware of DELPHINUS was designed and developed by the mission team in a way that did not consider thermal issues. In particular, DELPHINUS did not have enough paths for the internal heat generated at the sensor board to escape to the panel or other components. When developing in-house instruments, it is important for thermal teams to supervise the hardware design so that the heat generated inside the instrument can easily escape to external components. This is because there must be a heat path both inside and outside the instrument to let the internal heat escape, while insulating the internal heat can be done either inside or outside the instrument. Actually, the thermal design concept of coupling the vaporization chamber and the communication module was realized by cooperating with the propulsion team and the communication team when designing and developing the hardware. This procedure should have been taken in the mission components, too. In the end, we regarded this result to be acceptable for two reasons. Firstly, the possibility of this worst hottest condition to actually occur is low, and in most occasions it is likely for DELPHINUS to be under

9 International Conference on Environmental Systems Table 8. Results of Thermal Desktop® Flight Model simulations after thermal vacuum test modification. All components meet the operation temperature requirements, but DELPHINUS sensor board has no margin in the hottest condition. Component Temp. Boundary [°C] Simulation Results[°C] Margin [°C] Coldest Hottest Coldest Hottest Coldest Hottest SADA -30 70 -19 33 11 37 BAT 0 45 10 31 10 14 PCU -20 60 2 47 22 13 COMM -20 60 14 40 34 20 OBC -20 60 0 46 20 14 PROP IF BD -40 85 8 37 48 48 ADCS -30 50 -1 34 29 16 PHOENIX -30 50 1 38 31 12 DELPHINUS -30 70 1 47 31 23 FPGA BD DELPHINUS -30 40 -8 40 12 0 CAM+SENSOR CLOTH -50 90 -39 54 11 36 TANK 5 60 15 31 10 29 40 °C. Secondly, the operation temperature in Table 5 shows the range for maintaining performance. Even though the operation performance may degrade, we have tested that DELPHINUS can be operated in temperature conditions up to 50°C. Additionally, we conducted simulations for the required power in the DV-1 phase with the modified model. The largest total required power possible resulted to be 47.7W, which is below 48.5W, the least generable power of the SAP. These simulation results demonstrated the mission feasibility of EQUULEUS in aspects of thermal design.

V. Conclusion EQUULEUS is a 6U deep space CubeSat which will fly to EML2 as a secondary payload of NASA’s SLS. It will demonstrate the ability of trajectory control within the Sun-Earth-Moon region for the first time by a nano-spacecraft, and also conduct scientific observations at EML2 and on the way there. To achieve this mission, the thermal design of EQUULEUS must consider three major constraints. Due to the uncertainty of the launch date and the design of the libration orbit for observation, EQUULEUS will experience sunlight on four panels out of six. Moreover, EQUULEUS adopts a water propulsion system, which requires large latent heat for vaporization and care for propellant freezing. Under these constraints, we developed the thermal design concept of making the surface properties of the four panels homogenious and utilizing the waste power of the communication module to vaporize the water at the vaporization chamber. In order to validate this design concept, we made an one-node analysis to decide the surface properties, a three- node analysis to optimize the heat conductance around the vaporization chamber, and a Thermal Desktop® analysis to make detailed simulations. After the thermal design was created and validated, we developed the Engineering Model of EQUULEUS, and conducted a thermal vacuum test to verify the design and its implementation. We modified the Thermal Desktop® model based on the results of the thermal vacuum test, and made simulations in the hottest and the coldest conditions the spacecraft will experience. The results showed that one mission component called DELPHINUS has a possibility of exceeding the allowable temperature range, and we were required to make a design change in the Flight Model. To solve the problem in DELPHINUS, we revised operation strategies and temperature ranges in the Flight Model. Also, to control heat conductances between components precisely, several types of thermal interface materials were tested and introduced. We verified the effects of these design changes by conducting another thermal vacuum test for the Flight Model. We had some problems in the modification, but created the final Thermal Desktop® model for the Flight Model of EQUULEUS. However, simulations based on this final model showed that DELPHINUS still had no temperature margin. This was because the hardware design of DELPHINUS made it difficult for the internal heat generated in the sensor board to escape to external components. We regarded this result as acceptable, and the thermal

10 International Conference on Environmental Systems design of EQUULEUS was shown to satisfy the requirements to successfully reach EML2 and conduct scientific missions.

Acknowledgments We would like to thank I. Mase and K. Yamaguchi (NESTRA), for their suggestions for the thermal design and great help and operations to conduct the thermal vacuum tests of EQUULEUS. Also, we would like to thank H. Masui (Kyushu Institute of Technology) for his support for the Flight Model thermal vacuum test. Finally, we would like to acknowledge the continuous support of M. Nakano (WEL Research), and the advice of the EQUULEUS orbital design of Y. Kawabata.

References Ryu Funase, Satoshi Ikari, Yosuke Kawabata, Shintaro Nakajima, Shunichiro Nomura, Kota Kakihara, Ryohei Takahashi, Kanta Yanagida, Shuhei Matsushita, Akihiro Ishikawa, Nobuhiro Funabiki, et al, “Flight Model Design and Development Status of the Earth―Moon Lagrange Point Exploration CubeSat EQUULEUS Onboard SLS EM-1,” SSC18-VII-05, 32nd Annual AIAA/USU Conference on Small Satellites, Utah, USA, 2018. 2 Stefano Campagnola, Javier Hernando-ayuso, Naoya Ozaki, Nicola Baresi, Tatsuaki Hashimoto, Yasuhiro Kawakatsu, Kota Kakihara, Yusuke Kawabata, Takuya Chikazawa, Ryu Funase, Toshinori Ikenaga, Kenshiro Oguri, and Kenta Oshima. “Mission analysis for the EM-1 CubeSats EQUULEUS and OMOTENASHI,” IAC-18-B4.8.2x45356, 69th International Astronautical Congress, volume 1, pages 1–7, Bremen, Germany, 2018. 3 Jun Asakawa, Keita Nishii, Hiroyuki Koizumi, Naoki Takeda, Ryu Funase, Kimiya Komurasaki, “Engineering Model Development of the Water Resisojet Propulsion System: AQUARIUS for the SLS EM-1 CubeSat: EQUULEUS,” IEPC-2017-401, 35th International Electric Propulsion Conference, Atlanta, USA, 2017. 4 Yuki Koshiro, Naoya Ozaki, Shuhei Matsushita, Akihiro Ishikawa, and Ryu Funase, “Thermal Design and Analysis for a 6U Deep Space CubeSat EQUULEUS under Tightly-Coupled Spacecraft Resource Constraints,” 2017-f-047, 31st International Symposium on Space Technology and Science, Matsuyama, Japan, 2017. 5 Shuhei Matsushita, Toshihiro Shibukawa, Keidai Iiyama and Ryu Funase, “Thermal Design and Validation for a 6U CubeSat EQUULEUS under Constraints Tightly Coupled with Orbital Design and Water Propulsion System,” ICES-2019-193, 49th International Conference on Environmental Systems, Boston, USA, 2019.

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