25th Workshop on JAXA Astrodynamics and Flight Mechanics

Design of a Lander for In-Situ Investigation and Sample-Returnfrom a Jupiter Trojan Asteroid on the Solar Power Sail Mission

Ralf C. Boden 1), Osamu Mori 2), Takanao Saiki 2), Jun’ichiro Kawaguchi 2),and the Solar Power Sail Study Group 1) ,2)

1)The University of Tokyo, Tokyo, Japan2)Japan Aerospace Exploration Agency (JAXA), Sagamihara, Japan

Abstract: The paper discusses the design of a Lander, used for the in-situ exploration, and possible sample retrieval from aJupiter Trojan asteroid, as part of the JAXA Solar Power Sail (SPS) mission. Due to the restricted mobility and large size ofthe main spacecraft, use of a separate Lander is being investigated to perform the proximity operations at the targeted Trojanasteroid(diameter 20–30 km). Due to the long signal delay times at Jupiter distance, the Lander must perform most of itsmission autonomously. Furthermore, the distance from the Sun makes it difficult to use photo-electric power, leading to abattery powered design of the Lander. The Lander mission focuses on in-situ science, including collection and investigationof samples taken form both the surface and subsurface of the asteroid. As the SPS mission also considers the optional returnof collected samples to Earth via the main spacecraft, the Lander system includes sample retrieval and transfer mechanisms,and is capable of rendezvous and docking with the main spacecraft to perform the required sample-transfer. The Landermust therefore be capable of both autonomous descent and landing, as well as ascent, rendezvous, and docking. The systemsnecessary to perform these operations are described in this paper, alongside the science instruments and mechanisms used toperform the sample collection for both in-situ analysis and sample-retrieval.

1 THE JAXA SOLAR POWER SAIL MISSION

1.1 SPS Mission Overview

The JAXA Solar Power Sail (SPS) mission, launching in theearly 2020s, aims at performing a rendezvous with a JupiterTrojan asteroid, with the possibly of a sample-return to Earth[1]. The spacecraft uses a 50 m×50 m (2500 m2) membranesurface, covered with thin-film solar cells, to enable oper-ation of its high ISP ion-engines (7000 sec) at distances upto 5.2 AU from the Sun (Figure 1). The SPS utilizes a spin-stabilized solar sail; a concept that was demonstrated in 2010by IKAROS [2].

Fig. 1: Illustration of the 50 m solar power sail, with its thin-filmsolar cell covered membrane.

The current SPS time-line considers a launch in 2021. Oneor multiple Earth Swing-bys using the EDVEGA manoeuvreare performed before the trajectory targets Jupiter for a finalswing-by, putting the SPS on an intercept trajectory with aTrojan asteroid at either the L4 or L5 point. Figure 2 illus-trates the SPS mission, including an outline of the missiontime-line.

Next to remote science observation of the Trojan asteroidand in-situ investigation of samples on the asteroid surface,the SPS mission plans to conduct a number of scientific ob-servations during its outbound cruise phase, making use ofthe expended mission duration and wide range of solar dis-

Fig. 2: Overview of the SPS mission, including the current missiontime-line.

tances traversed on its way from Earth to Jupiter. Plannedobservations include cosmic infra-red background mappingand in-situ measurement of the solar system dust distribu-tion. Due to the long cruise and overall mission durationthese investigations will be concluded before reaching theTrojan asteroid [1].

1.2 Lander Mission Scenario

Two mission scenarios are being investigated for the Lander,carried along on the SPS for landing on the Trojan asteroid;Plan A and B. Plan A considers only in-situ investigationof samples by the Lander on the asteroid surface. Plan Bexpands this scenario to include the retrieval of samples andtransfer to the main spacecraft (MSC) for Earth-return.

Prior to reaching the target asteroid the Lander is in hiberna-tion. It is switched on for regular heath checks, and option-ally to use some of its science instruments during the mis-sions cruise phase. The Lander main mission at the asteroidcan be divided into the following operations:

• MSC descent from home position (HP)l• Lander deployment at 1 km altitude (HP2)• MSC ascent to 50 km altitude for standby• Lander descent to asteroid surface (powered)• elimination of descent velocity (hovering)• touch-down on asteroid surface• sampling operations in-situ science (incl. context)

ASTRO-WS ISAS/JAXA 2015 Page 1 of 7

• powered take-off and rendezvous with MSC ∗• Lander docking/berthing with MSC ∗• sample transfer from Lander to MSC ∗• Lander decommissioning ∗

The marked items (∗) indicated steps associated only withthe Plan B option. The Lander operations are also depictedin Figure 3.

Fig. 3: Lander mission sequence, showing both the Plan A and Bscenarios.

2 MISSION OBJECTIVES AND SYSTEM REQUIREMENTS

The main mission objectives are divided into those for PlanA, and the additional Plan B objective:

• collection of context science data from the Landingsite

• collection and in-situ analysis of surface and sub-surface (up to 1 m) samples from a Trojan asteroid(diameter: 20–30 km, P/D-type)

• collection of samples for transfer to the MSC forEarth-return (extra success / Plan B)

The targeted asteroid type and size has been defined in [1].

For the study of the Lander systems, a number of Point ofDeparture (POD) parameters are formulated as starting point(Table 1). These are validated and/or further developed dur-ing the progression of the Lander study.Tab. 1: List of Point of Departure (POD) parameters for the Trojan

Lander study.

Based on these POD parameters and the mission objectives,a number of system requirements can be formulated, includ-ing the extra requirements for Plan B:

• Lander operational after 15 year cruise phase (incl.Jupiter swing-by)

• operation Trojan asteroid environment at 5.2 AU Sun-distance (52 W/m2, Tasteroid <−99◦C)

• autonomous descent and safe landing on asteroid withsurface gravity up to 8.36 mm/s2

• operation for at least 20h on asteroid surface (+ de-scent, ascent and rendezvous operations)

• surface and subsurface (extra) collection of samplesand in-situ analysis

• autonomous ascent, rendezvous and docking withMSC, incl. sample-transfer (extra)

3 LANDER CONFIGURATION

The Lander configuration consists of an octagonal shapedmain body (width: 650 mm, height 400 mm), equipped withlanding legs for clearance (300 mm). Devices for sampling(surface & sub-surface), sample-transfer and berthing pro-trude from the top and bottom plates of the Lander (Figure4). The individual components, belonging to the Lander’ssubsystems and instruments are described in the followingSections 4 and 5.

Fig. 4: Schematic showing the external dimensions of the TrojanLander.

The Lander is stored inside the central cylinder of the MSC,as shown in Figure 5. This limits the lander size, includingrequired clearances for the deployment and docking/berthingmechanisms (Section 5.3).

Fig. 5: Location of the Lander inside the MSC.

4 LANDER SUBSYSTEMS

In the following, the individual Lander subsystems, as seenin Figures 4 and 6 are described. The internal layout of theLander in Figure 6 is currently based only on accommodat-ing the required component volumes. However, some con-siderations to component placement have already been takeninto account.

Fig. 6: Internal view of the Trojan Lander. The individual compo-nents are coloured to show individual subsystem groups.

4.1 Propulsion – RCS

The design of the propulsion system is based on a 12 thrusterset-up, similar to the one used for the Hayabusa spacecraft(Figure 7) [3]. This set-up provides redundancy in case ofsingle-point failures. The RCS is used for both propulsivemanoeuvres and attitude control 1 N thrusters are hereby suf-ficient for hovering and take-off in the asteroid’s low gravity

ASTRO-WS ISAS/JAXA 2015 Page 2 of 7

environment. In addition they provide a sufficiently smallminimum impulse bit for attitude control and docking oper-ations.

Fig. 7: 12 thruster layout used fro the Lander.

Based on the delta-v requirements (Table 2) and the need toavoid sample contamination, a cold gas (CG) system usingN2 is selected due to its relatively low mass, simplicity, andreliability. An experimental propulsion system using super-critical CO2 was also considered, but did not provide sig-nificant mass or volume reduction to justify the added com-plexity. Total propellant mass is 4.43 kg, including individ-ual margins for the Plan A (50%) and B (20%) operations.The N2 CG-system system utilizes high pressure (HP) on the

Tab. 2: Breakdown of the Lander delta-v requirements.

storage side, and a redundant regulator system to supply thethrusters on the low pressure (LP) side with propellant. Asystem of redundant isolation valves minimizes leakage dur-ing the 15 year cruise (≈40 g), and ensures supply of fuel tothe LP side after activation.

A schematic of the RCS, showing the HP storage side andthe LP thruster side is provided in Figure 8. A connectionfrom the LP side to operate the pneumatic sampling drill(Section 5) is considered as an option.

4.2 Guidance, Navigation and Control – GNC

The GNC system of the Lander consists of a number of in-struments, needed to ensure safe landing on the asteroid dur-ing the descent, as well as rendezvous and docking capabili-ties for the rendezvous phase of the Plan B scenario (Figure3). The detailed descent operation is shown in Figure 9, withTab. 3: Overview of the Lander’s GNC components, including the

phased array antenna of communication system (Section 4.3),used as a RF sensor during rendezvous with the MSC.

Component obtained measurements

LIDAR range, range-rateIRU angular velocitiesFlash-LIDAR relative attitude to asteroid/MSCONC asteroid features, MSC docking markersAccelerometer accelerationsSTR/SAS absolute orientation to Sun/StarsRF sensor heading to MSC

an expected residual velocity of 1 m/s (vert.) and 0.05 m/s(lat.). Both descent and the optional ascent/rendezvous op-erations need to be performed autonomously, as the long sig-nal times to and from Earth prevent any direct control formthe ground.

Fig. 8: Schematic of the RCS.

Fig. 9: Overview of the Lander descent phase.

Fig. 10: Communication architecture used for the Lander, with theMSC acting as a relay for communications between Earth andthe Lander.

4.3 Communication – TT&C

The TT&C system of the Lander is based on the architectureshown in Figure 10, using the MSC as a relay. The bottle-neck is the transfer rate between Earth and the MSC, withexpected bit-rates of around 1 kbps at 5.2 AU. Different sys-tems are under investigation, such as X-, S- and UHF-band(Table 4). While UHF seems as the most promising systemregarding data-rate, mass, and power, there is an ongoing

ASTRO-WS ISAS/JAXA 2015 Page 3 of 7

trade between the different systems, due to additional con-siderations. These include use of the X-band system hard-ware as an RF-sensor, interference of the UHF system withthe sail of the MSC, and the availability of newer S-bandhardware.Tab. 4: Overview of mass, power and data-rate of the considered

TT&C systems.

4.4 Electrical Power – EPS

Due to the low solar constant at 5.2 AU (≈50 W/m2) the Lan-der relies on a battery as its power source. This results in alimited operational time after separation from the MSC. Tosuccessfully conduct the required sampling and science op-eration, a period of 20 h on the surface of the asteroid is re-quired. The Lander battery is sized to provide this surfacetime, as well as the time needed for the initial descent andlanding, and in case of Plan B the ascent and rendezvousphase. In addition, a fixed amount of 600 Wh is reservedfor payload operations during the surface phase (Section 5).Figure 11 shows the Lander power profile during its opera-tion time of 22 h (Plan A) or 33 h (Plan B).

Fig. 11: Power profile of the Lander mission scenario.

A trade between primary and secondary battery options ismade to identify the lowest mass system. The results of bothoptions are listed in Table 5, showing that the primary batteryhas higher mass requirements despite the higher cell energydensity of 468 Wh/kg, compared to the 190 Wh/kg of sec-ondary battery cells. The reason for this discrepancy is thelow efficiency of the primary battery at high currents, due tothe high internal resistance of the Saft LSH-20 cells.Tab. 5: Trade of primary and secondary battery options for both the

Plan A and B scenario.

In addition to a reduction in mass, the use of a secondarybattery in the Lander provides additional advantages, suchas increased flexibility and added safety due to the abilityto check and charge the battery prior to Lander deployment.A simplified Lander design is also made possible due to re-duced thermal loads on the TCS (Section 4.5).

4.5 Thermal Control – TCS

The thermal control of the Lander is designed as a simplesystem, using a combination of radiator surfaces and multi-layer isolation (MLI). Currently, a simple single node modelis used to determine the radiator size and temperature insidethe Lander during high power steady-state conditions (finaldescent and rendezvous/docking). Heaters are used to keepthe temperatures from dropping below the operational limitsof the subsystem components during low power sequences(surface operation). Their individual minimal and maximaloperating and non-operating temperatures are shown in Fig-ure 12.

Fig. 12: Temperature range of individual lander components usedin the thermal design of the Lander.

Based on the Lander’s operational sequence (Section 2) andpower profile (Section 4.4), the Lander temperature is sim-ulated with varying power profiles (Figure 13). All scenar-ios are identical with exception of the surface phase. Thenominal profile has a total power consumption of 104 W. Inaddition, a 130 W peak power and a 4 W hibernation modeprofile are used.

Fig. 13: Temperature progression based on Lander power profile fordifferent operating cases and initial temperatures.

Due to its high thermal mass (≈100 kg), the Lander can keepits temperature for an extended period of time, even when inlow power / hibernation mode. During high power phases,the temperatures increase significantly, but remain within theoperational range. Heater, radiator, and MLI masses are rel-atively small, resulting in a total TCS mass of ≈2 kg.

5 PAYLOAD SYSTEMS

The Lander payload consists of instruments for the collec-tion and analysis of samples from the asteroid, as well ascontext science. The devices needed for Plan B for the sam-ple transfer (including docking/berthing) are also consideredas part of the payload systems.

ASTRO-WS ISAS/JAXA 2015 Page 4 of 7

5.1 Science Instruments

The entire science payload is restricted to a mass of 20 kgand 600 Wh of energy, including the sampling devices (Sec-tion 5.2). Table 6 shows a list of the science instruments, in-cluding allocated mass, peak power and energy usage. Thecentral instrument is the high resolution mass spectrometer.Two compatible systems are currently being developed inJapan (MULTUM) and Europe (Cosmo-Orbitrap) [], []. Ad-ditional instruments are used to provide additional methodsto analyse the collected samples (GC, Microscope), and toprovide context science (Camera, Magnetometer).Tab. 6: List of science instruments, including their mass, power and

energy requirements.

Figure 14 shows the science sequence that has been devel-oped for the expected 20 hr operation and 600 Wh energyresources. The sequence includes the time needed to beginthe science sequence after successful landing, as well as anadditional sampling sequence for the Plan B sample returnand a blank sample sequence before touchdown.

Fig. 14: Breakdown of the Lander science sequence. Additionalsequences for redundancy and calibration have been included,as well as a landing confirmation period before the science se-quence is initialized.

5.2 Sample Collection, Distribution, and Transfer

Sample collection is performed with two different devices.A sample horn, similar to that of Hayabusa and Hayabusa2 is used to collect surface samples using a projectile (Fig-ure 15) [4]. Sub-surface samples are collected with a newlydeveloped pneumatic drill, which also utilize a projectile toperform the sampling (Figure 16).

In addition to the two sampling devices, a distribution sys-tem is used to transfer the samples to the science instruments(HRMS, GC), as well as to the sample transfer container.Figure 17 shows an overview of the sample system compo-nents.

Fig. 15: Hayabusa sampler-horn [4].

Fig. 16: Image of the sub-surface sampling device currently underdevelopment at JAXA.

Fig. 17: Schematic of the sampling and distribution system.

The distribution is performed by a revolver device (Figure18) which catches the samples and rotates them to the re-spective instruments. A hole in the revolver plate allows the

Fig. 18: Drawing of the revolver – sample distribution device. Thepath to the HRMS is also shown.

collected samples to pass through to the sample container forthe sample-return part of the mission. Figure 19 shows howthe sample transfer device operates to transfer the sample tothe container and finally to the MSC, using an extendiblecarbon-fibre boom to push the sample container into the re-entry capsule.

ASTRO-WS ISAS/JAXA 2015 Page 5 of 7

Fig. 19: Schematic of the sample transfer process from sampler tocontainer, and from Lander to MSC.

5.3 Docking and Berthing Mechanism

In order to allow the sample to be transferred between Lan-der and MSC a docking/berthing mechanism is used to alignthe Lander with the MSC. To achieve propel alignment, therelative rotation of the Lander to the MSC must be cancelledand the Lander needs to be attached to the MSC. A berthingmast is used to attach the Lander to the MSC,as shown inFigure 20, using an extendible boom on the Lander and amagnetic latch on the MSC. By retracting the boom, the Lan-der is placed in close proximity to the MSC. Proper align-

Fig. 20: Schematic of the docking/berthing concept using a mag-netic latch and a carbon tube boom.

ment of the sample-transfer path is achieved using a pro-trusion/groove mechanism on the Lander and MSC (Figure21). This mechanism also eliminates any residual rotationbetween the two spacecraft.

6 LANDER MASS BREAKDOWN AND BUDGET

Based on the previously described systems, the total mass,power and volume breakdown of the Lander for the PlanB scenario is listed in Table 7. The results shows that theLander fulfils the 100 kg mass restriction, as well as the vol-ume and power requirements. In case of a Plan A option,Lander mass can be reduced to below 80 kg, due to lower

Fig. 21: Alignment mechanism to eliminate relative rotation andensure that the sample can be transferred from Lander to MSC.

fuel and battery masses and a less sophisticated AOCS sys-tem. While the sample-transfer and docking/berthing mech-anisms are also not needed, the reduction in payload massmay be compensated by adding additional science instru-ments to the Plan A option.

Within the 100 kg restriction of the Lander, Plan A wouldalso allow extended surface operations by equipping a largerbattery.Tab. 7: Breakdown of the Lander mass, power, and volume for all

subsystems, on an individual component level.

7 CONCLUSION

In conclusion, the preliminary study of a Trojan Landersystem for in-situ analysis and possible sample-return ofa Jupiter Trojan asteroid shows that the system can beachieved as part of the JAXA SPS Mission, within the givenrestrictions.

A continuation of the system design will be performed aspart of a joint study between JAXA and DLR, including co-operation with the European and Japanese science commu-nity. The results are expected to provide a more detailedanalysis of the Lander systems, and the development of a sci-ence package that can provide the best scientific data withinthe constraints of the SPS Mission.

REFERENCES

[1] O. Mori, T. Saiki, H. K. Y. T. Y. M., “Jovian TrojanAsteroid Exploration by Solar Power Sail-craft,” Inter-

ASTRO-WS ISAS/JAXA 2015 Page 6 of 7

national Symposium on Space Technology and Science(ISTS), 2015.

[2] Tsuda, Y., Mori, O., Funase, R., Sawada, H., Yamamoto,T., Saiki, T., Endo, T., Yonekura, K., Hoshino, H., andKawaguchi, J., “Achievement of IKAROS—Japanesedeep space solar sail demonstration mission,” Acta As-tronautica, Vol. 82, No. 2, 2013, pp. 183–188.

[3] Kuninaka, H., Nishiyama, K., Funaki, I., Yamada,T., Shimizu, Y., and Kawaguchi, J., “Powered flightof HAYABUSA in deep space,” proceedings of 42ndAIAA/ASME/SAE/ASEE Joint Propulsion Conference,2006.

[4] Yano, H., “Sampling Systems for Hayabusa and follow-on missions: Scientific Rationale, Operational Consid-erations, and Technological Challenges,” InternationalMarco Polo Symposium and other Small Body SampleReturn Missions, May 2009.

ASTRO-WS ISAS/JAXA 2015 Page 7 of 7