Architecture Trades for Accessing Small Bodies with an Autonomous Small Spacecraft

Home , 99942 Apophis, Areostationary orbit, Artemis 1, Artemis 2, Artemis 3, ASTER (spacecraft)

Sandro Papais Benjamin Hockman Dept. of Mechanical Engineering Jet Propulsion Laboratory McGill University California Institute of Technology Montreal, QC H3A 0G4, Canada Pasadena, CA, 91109 +1-514-979-2593 +1-626-639-5505 [email protected] benjamin.j.hockman@jpl.nasa.gov Saptarshi Bandyopadhyay Reza Karimi Jet Propulsion Laboratory Jet Propulsion Laboratory California Institute of Technology California Institute of Technology Pasadena, CA, 91109 Pasadena, CA, 91109 +1-626-318-4174 +1818-354-6611 [email protected] [email protected] Shyam Bhaskaran Issa Nesnas Jet Propulsion Laboratory Jet Propulsion Laboratory California Institute of Technology California Institute of Technology Pasadena, CA, 91109 Pasadena, CA, 91109 +1818-354-3152 +1-818-354-9709 [email protected] [email protected]

Abstract—Characterizing the composition, properties, and en- 8. CONCLUSION ...... 17 vironments of Small Bodies is key to understanding the origins ACKNOWLEDGMENTS ...... 17 and processes of the Solar System. Traditionally our knowledge has been limited to ground observation and select few missions REFERENCES ...... 17 which cannot fully characterize the diversity of Small Bodies. BIOGRAPHY ...... 19 Advances in miniaturized spacecraft technologies have recently enabled small spacecraft to perform missions in deep space, as demonstrated by Mars Cube One in 2018. Additional missions 1. INTRODUCTION are being developed to further mature these technologies and Small Bodies, including Near-Earth Objects (NEOs), aster- expand their capabilities. We investigate a new approach to oids, and comets, are numerous and diverse in their com- exploring Small Bodies, where standalone small spacecrafts can position and origin. Small Bodies are high-priority targets be used as a more affordable approach to autonomously navigate, rendezvous, and characterize them. We review relevant for origins science, human exploration, in situ resource uti- mission concept studies, required technologies, available targets, lization, and planetary defense. To date, observations of architecture trade offs, and baseline mission design options. Small Bodies have consisted of broad surveys from ground We show that using near-term technologies, available in less telescopes and a few missions to rendezvous and characterize than 3 years, it is possible for a standalone small spacecraft select ones in detail such as NEAR Shoemaker, Hayabusa, to rendezvous with several Small Bodies. It was found that Rosetta, OSIRIS-REx (shown in Figure 1), and Hayabusa a 24 kg and 180 kg spacecraft would be capable of delivering 2. From these observations and missions, the diversity of payloads of 1.5 kg and 10 kg respectively to several interesting Small-Body populations has become apparent. However, the near-Earth asteroids candidates. Critical enabling technologies diversity cannot be fully characterized from ground surveys were identified as highly-capable (delta-V>3 km/s) miniature electric propulsion system, high-efficiency (power density>100 or the limited number of space missions that have flown. W/kg) deployable solar arrays, and improved onboard autonomy algorithms. Advances in miniaturized instruments, high-performance radiation-tolerant avionics, and interplanetary communications systems can also be leveraged. In the long term, this architecture could enable a fleet of standardized autonomous small spacecraft to perform a cursory exploration of a representative sample of the Small Body population.

TABLEOF CONTENTS 1. INTRODUCTION ...... 1 2. INTERPLANETARY SMALL SPACECRAFT STATE OF PRACTICE ...... 3 3. PRELIMINARY MISSIONAND TARGETS ANALYSIS .6 4. LAUNCH OPPORTUNITIESAND EARTH ESCAPE ....8 Figure 1. Artist’s representation of the OSIRIS-REx 5. ENABLING SUBSYSTEMSAND TECHNOLOGIES .....9 spacecraft at near-Earth asteroid Bennu [1]. 6. SPACECRAFT ARCHITECTURE TRADE SPACE .... 13 7. REFERENCE MISSION DESIGN ...... 15 Only five missions have currently attempted to operate for extended periods of time in close proximity to such Small 978-1-7281-2734-7/20/$31.00 c 2020 IEEE Bodies. The difficulties encountered by these missions high- 1 light the challenges involved in operating in this resource- diverse population. constrained, largely uncertain, and dynamic environment. Autonomy has been shown to be essential for current mis- In order to address gaps in autonomy required for sustained sions operating in this environment, in particular for fast fly- operation with limited ground interaction, several new ca- bys and touch and go (TAG), but has only been demonstrated pabilities are needed. One particularly difficult challenge is in limited cases. More capable autonomy will enable more during spacecraft approach. Traditionally, when a spacecraft missions to reach and explore a wider range of diverse bodies. approaches a Small Body it periodically images the target to aid in target-relative navigation. A number of algorithms Small spacecrafts, defined as 180 kg wet mass or less, typ- and techniques, like Stereo-Photo-Clinometry (SPC), have ically have shorter development cycles, smaller teams, and been developed to navigate and characterize the body once consequently lower costs. CubeSats are a subset of small the body is 1000s of pixels in area. However, there is also spacecrafts made of units (U) where 1U is a 10 cm × 10 valuable information even when the Small Body is 10s-100s cm × 11.35 cm cuboid. CubeSats have the added advantage of pixels in area. of standardized form-factor and deployment containers which simplify integration. By incorporating recent improvements To address these gaps, we have developed a multi-phased ap- in software and hardware it has been possible for many proach to estimate the physical (size and shape) and dynam- groups to achieve high capability in small packages, driving ical (rotation rate and pole) properties of an unknown Small interest and growth in small spacecrafts. Whereas 700 small Body in previous work [4] (shown in Figure 2). In Phase 1, spacecraft were launch from 2006-2015 it is expected 3,600 when the body is one to a few pixels in span we can estimate will be launched in the subsequent decade (2016-2025) [2]. the Small Body’s periodicity of rotation from variations in the pixel intensity light curve. In Phase 2, when the Small Body Almost the entirety of the small spacecraft missions have is 5 to 50 pixels in span we can use the its silhouette to resolve been earth orbiting. However, the advances in miniaturized an initial shape, given estimates of periodicity from Phase 1. technologies and science instruments have also enabled the During Phase 3 the surface features become more resolved possibility for small spacecraft to perform challenging in- and we can use landmark-based mapping techniques to refine terplanetary missions. Small spacecraft have potential to the geometry and rotation. The information in each phase is reach numerous destinations, such as Small Bodies, for novel also used to refine the spacecraft’s trajectory. mission concepts and targeted planetary science investigations. Even deep-space CubeSats are becoming capable In this paper, we focus on assessing the feasibility of a space- to make unique planetary science contribution by comple- craft architecture to facilitate access and characterization menting larger missions through unique vantage points and of Small Bodies. Previous interplanetary small spacecraft multipoint measurements, performing high-risk science, or missions flown and in development are reviewed in Section 2 exploring new and unknown destinations [3]. along with relevant concepts studies. A method for preliminary mission and target analysis are presented in Section 3 By combining autonomy and small-spacecraft capabilities, and Section 4 to limit architecture options, guide initial a diverse population could be more affordably explored in requirements, and reduce the design space. In Section 5, key larger numbers. With numerous spacecraft and destinations subsystem options to facilitate the mission are reviewed and and limited ground communications capacity, such assets the overall performance envelopes are outlined. A cohesive would have to rely on on-board decision-making. This set of baseline architectures are developed and key trade- would allow the ground operations to focus on higher-level offs are assessed across the design space in Section 6. Two management of parallel missions. As a result, this can enable architecture point designs are presented in Section 7 and a full more in-depth investigations of the Small Body population’s trajectory optimization and mission design is discussed for heterogeneous compositions, developing a better understand- one NEA target. Finally, the paper is concluded in Section 8. ing of their origin. Ultimately this could lead to a game- changer in our ability to conduct precursor missions, in situ science investigations, and in our understanding of this

Figure 2. A multi-phased sequence of algorithms to estimate the physical and dynamical parameters of a Small Body.

2 2. INTERPLANETARY SMALL SPACECRAFT (ESPA) ring. At this time only two small spacecraft missions, STATE OF PRACTICE PROCYON and MarCO, have operated in deep space. There are 13 CubeSats scheduled for the Artemis-1 launch in 2020 This section reviews relevant small spacecraft missions be- and one CubeSat planned for the DART launch in 2021. yond Earth orbit that are in various project life cycle phases, Several future rideshare opportunities exist through Artemis- from concept studies (pre-phase A) to mission completion. 2, Psyche, Hera, Lucy, and other interplanetary missions. The focus of this section is based on standalone small spacecraft missions with moderate ∆V capability or a Small-Body IMPEL target destination. As a results small spacecraft missions with WATER NEACO low propulsion capability (∆V < 500 m/s) were typically not Athena Trailblazer included, with a few exceptions with other relevant aspects. APEX ASTER ARROWS A summary of the missions reviewed along with the lead VAMOS institution and target destination is provided in Table 1. MiLuV

MISEN EscaPADE Table 1. Summary of 31 standalone small spacecraft PrOVE JANUS PROCYON missions beyond Earth orbit. MAT CUVE DISCUS Mission Name Lead Institution Destination Chariot MMO PRISM M-ARGO WATER JHU APL Moon NEACORE VMMO MiLuV NASA GSFC Moon BOLAS Omotenashi BOLAS NASA GSFC Moon LunaH-Map IMPEL NASA JSC Moon DAVID NEA Scout MarCO

6U 12U Mid-ESPA ESPA Lunar IC LunaH-Map ASU Moon Lunar IceCube MSU Moon Pre-Phase A Phase A/B Phase C/D Post-Phase D Trailblazer NASA JPL Moon Figure 3. Categorization of 31 standalone small VMMO ESA Moon spacecraft mission concepts based on size and maturity. OMOTENASHI JAXA Moon NEA Scout NASA JPL SB Flyby The architectures can be classified by whether the launch is JANUS CU Boulder SB Flyby direct, delivering the spacecraft close to its final orbit or destination, or indirect, requiring large change in velocity, ∆V , APEX JHU APL SB Flyby maneuvers to reach the necessary destination. Regardless of PrOVE U Maryland SB Flyby the destination, these high ∆V maneuvers usually require low Athena ASU SB Flyby thrust electric propulsion systems due to their higher specific impulse. Whereas the direct, low ∆V , missions are typically ARROWS SETI SB Flyby able to complete their objectives with chemical or even cold NEACORE U Strathclyde SB Flyby gas propulsion systems. To highlight this distinction for DAVID NASA GRC SB Flyby the discussed missions, a comparison of propulsion system against destination for select interplanetary small spacecraft PROCYON JAXA SB Flyby missions is shown in Figure 4. M-ARGO ESA SB Rendezvous NEACO CU Boulder SB Rendezvous DISCUS Max Planck SB Rendezvous ASTER AEB SB Rendezvous Trailblazer PrOVE MMO CUVE U Maryland Venus IMPEL NEA Scout MarCO Chariot VAMOS NASA JPL Venus OmotenashiJANUS MMO MSS Mars MISEN UC Berkeley Mars EscaPADE UC Berkeley Mars MAT SSI Mars BOLAS ARROWS ASTER MarCO NASA JPL Mars WATER DAVID NEACO VAMOS MAT PRISM NASA GSFC Phobos VMMO APEX M-ARGO CUVE MISEN PRISM Chariot Purdue U Phobos MiLuV Athena DISCUS EscaPADE

Electric Propulsion Other LunaH-Map NEACORE A comparison of the spacecraft size against concept phase is shown in Figure 3. The architecture sizes are defined based on Lunar IC PROCYON form factors as 6U, 12U, mid-ESPA, and ESPA and typically Moon SB Flyby SB Rndz Venus Mars Phobos corresponds to spacecraft wet masses of 8-14 kg, 16-28 kg, 30-100 kg, and 100-180 kg respectively. An ESPA-class Figure 4. Categorization of 31 standalone small spacecraft is limited to 180 kg and can be launched from a spacecraft mission concepts based on destination and standardized payload adapter ring, known as the evolved ex- propulsion system. Other propulsion types include pendable launch vehicle (EELV) secondary payload adapter chemical, cold gas, and solar sail propulsion. 3 An additional classification of the architecture is standalone versus mother-daughter. For mother-daughter architectures, the small spacecraft is deployed by a larger spacecraft, known as the “mothership,” which also serves as a communication relay (typically using a small UHF radio and antenna). In the standalone architecture the data is instead returned direct to earth from the small spacecraft using a more powerful transponder and a high-gain antenna (HGA). However, data rates associated with small HGAs are typically much lower than the HGAs that can be accommodated on the larger mothership. Technology Demonstration Heritage The PROCYON (Proximate Object Close flyby with Optical Navigation) spacecraft was a NEA flyby missions developed by JAXA and University of Tokyo (see Figure 5). It launched with Hayabusa 2 in 2014, becoming the first interplanetary Figure 6. Top side cutaway of MarCO showing the micro-spacecraft (wet mass < 100 kg). It demonstrated layout of the flight hardware [10]. micro-spacecraft interplanetary communication, navigation, and electric propulsion. The 67 kg spacecraft successfully performed three months of operation before the failure of the propulsion system. PROCYON was developed with relatively There have also been a number of small spacecrafts landed low cost (a few million dollars) and a very short period (about on Small Bodies by larger spacecrafts in a mother-daughter 1 year), taking advantage of the heritage from Japanese Earth- architecture. In 2005 JAXA’s Hayabusa mission attempted orbiting micro-satellite missions [5]. to land MINERVA on 25142 Itokawa but failed due to a commanding error which sent the lander into heliocentric orbit instead. In 2014 ESA’s Rosetta mission deployed a 100kg lander Philae which became the first spacecraft to soft land on a comet after a partial failure of its landing system, resulting in several bounces into a non-optimal location. It performed a limited science analysis of the surface of comet 67P before losing battery and being unable to recharge. In 2018 JAXA’s Hayabusa 2 successfully deployed three landers (HIBOU, OWL, and MASCOT) ranging from 1-10 kg and capable of limited surface mobility. They performed a variety of science investigations on 162173 Ryugu using cameras, thermometers, infrared spectrometer, magnetometer, and ra- diometer [11]. NASA Small Innovative Missions for Planetary Exploration In 2014 NASA solicited CubeSat mission concepts to conduct Figure 5. Conceptual image of the PROCYON planetary science for the first time. The program was called micro-spacecraft and comet 67P [6]. SIMPLEx (Small Innovative Missions for Planetary Explo- ration). The proposals were limited to a $5.6 M cost cap The twin INSPIRE (Interplanetary NanoSpacecraft Pathfinder and a size of of 6 U or less. The aim was to be launched In Relevant Environment) 3 U spacecrafts were developed on the Artemis 1 mission. Two missions were selected: the by NASA JPL. They were designed for direct Earth-escape Lunar Polar Hydrogen Mapper (LunaH-Map) and the Cube- launch into heliocentric orbit to demonstrate they can operate, Sat Particle Aggregation and Collision Experiment (Q-Pace) communicate, and navigate far from earth. They contained a which are now awaiting launch. Three proposals were also compact magnetometer for making synchronous solar-wind selected for technology development: the Mars Micro Orbiter structure measurements [7]. The spacecrafts architecture (MMO), the Hydrogen Albedo Lunar Orbiter (HALO), and combined commercially available hardware with custom de- the Diminutive Asteroid Visitor using Ion Drive (DAVID). veloped subsystems, such as the IRIS radio. They were delivered to launch readiness in June 2014 but have not been Technology development for the 6 U DAVID mission worked scheduled for launch [8]. to miniaturize a visible-infrared spectrometer to determine the surface color and spectral properties of a small near- Building on lessons learned from INSPIRE, JPL further Earth asteroid (NEA). Several architecture trade studies were refined the approach to interplanetary CubeSat hardware, performed in 2014 with emphasis on propulsion systems to software, and operations for MarCO (Mars Cube One). The enable the flyby or rendezvous with a NEA [12]. The target two 13.5 kg, 6 U CubeSats were designed to accompany the asteroid selected was 2001 GP2 for its high accessibility, InSight Mars mission lander and act as a real-time commu- requiring approximately 400 m/s for flyby with a 10 W nication relay during the entry, descent and landing phase. electrospray thruster and 2 km/s for rendezvous with a 50 In order to meet the mission requirements a new deployable W ion propulsion system. From the SLS lunar launch a 15 high-gain folded-panel reflectarray antenna was developed m/s burn was planned for a lunar flyby providing a C3 of 0.2 for use with the IRIS radio. The CubeSats were completed km2/s2. in less than two years, from conception to completion of assembly, and were launched in May 2018 [9]. The spacecraft In 2017 a second SIMPLEx call was solicited, and the cost architecture is shown in Figure 6. cap was raised by an order of magnitude to $55M and 180 4 kg based on lessons from previous studies. The objectives of the new program was to perform compelling and targeted science investigations that exploit unique aspects of small spacecrafts, use available launch capacity to reduce missions costs, provide a means to mature technologies, and provide new flight experience opportunities [13].

Three missions were selected in June 2019 for further devel- Figure 8. Lunar IceCube preliminary spacecraft design opment. Janus, shown in Figure 7, is a twin small space- using a miniature electric propulsion system [17]. craft reconnaissance missions to binary rubble pile asteroids. EscaPADE (Escape, Plasma, and Acceleration Dynamics Explorers) is also a twin small spacecraft to characterize the acceleration processes driving ion escape from Mars’ atmo- NEA up to 1 AU distance from Earth [18]. NEA Scout has a sphere. Lunar Trailblazer will directly detect and map water compact camera for characterizing the asteroid and relative on the lunar surface to determine how its form, abundance, navigation to it. Several algorithms were developed for and location relate to geology. At least one of these missions onboard image processing and data analysis. They are used is expected to move to final selection and flight. EscaPADE to enhance target acquisition and move first order science and Janus would launch with the Pysche mission to the main analysis, which are traditionally completed on the ground, to asteroid belt, whereas Lunar Trailblazer would launch on a onboard the spacecraft [19]. OMTENASHI is a 6 U CubeSat commercial Lunar launch. and semi-hard lander that will demonstrate low-cost technology to land and explore the lunar surface. It is made up of an orbiting module, 6 kg solid rocket retro-motor to slow down, an inflatable airbag to cushion the landing, and a nano-lander surface probe. It will take measurements of the radiation environment and soil mechanics using accelerometers. ESA Deep Space CubeSat Studies Within the ESA General Studies Programme a set of studies were performed called LUnar CubeSats for Exploration (LUCE) to support European lunar exploration. Four concept studies were selected in 2017: LUMIO, MoonCARE, CLE, and VMMO [20]. All the mission concepts were 12 U cubesats, with VMMO using electric propulsion and the oth- Figure 7. NASA rendering of one of the twin Janus ers chemical. The Lunar Volatile and Mineralogy Mapping spacecraft flyby of a binary asteroid [14]. Orbiter (VMMO) is a very low lunar orbit mission with a 6 kg multi-wavelength chemical LIDAR system to study south Artemis 1 CubeSats pole lunar volatiles. The 24 kg spacecraft uses cold gas thrusters and two IFM Nano Thrusters capable of a ∆V of Thirteen deep space 6 U CubeSat missions were competi- tively selected as secondary payloads on Artemis 1, the first 750 m/s [21]. SLS test flight. The focus of the missions is on testing key Another separate internal study in 2016, focused on beyond technologies beyond low-Earth orbit and performing science investigations in lunar orbit and beyond. More CubeSat lunar orbit, was also carried out at ESA’s concurrent de- opportunities can be expected on future Artemis missions sign facility called M-ARGO (Miniaturised Asteroid Remote and 12 U secondary payload opportunities have already been Geophysical Observer). It is a 12 U standalone mission which announced for Artemis 2. The Lunar IceCube, LunaH-Map, would launch as a secondary payload to near Earth escape NEA Scout, and OMOTENASHI missions are particularly and use electric propulsion to rendezvous with a NEA. It relevant due to their substantial ∆V capability. is a 22.3 kg spacecraft with a mini-RIT gridded ion engine capable of 3.7 km/s of ∆V . The design, shown in Figure 9, includes a reflectarray antenna and articulated solar arrays The Lunar IceCube mission led by Morehead State University (MSU) to search for water ice and other resources at a low which generate 120 W. The mission has identified at least 4 feasible NEO targets less than 50 m in diameter for a orbit of 62 miles above the moon. The 14 kg, 6 U spacecraft mission in 2021 [22]. It is now in Phase A studies for further bus (shown in Figure 8) is high capable with 120 W power generation, ∆V of 2.4 km/s, and 1σ pointing accuracy of technology development. 0.14 arcseconds. It is designed to house a 3.5 kg, 2 U, 17.8 Additional Mission Concepts W compact broadband infrared spectrometer and cryocooler payload [15]. The LunaH-Map mission is led by Arizon State The NASA Planetary Science Division released the Planetary University (ASU) and will map hydrogen enrichments within Science Deep Space SmallSat Studies (PSDS3) solicitation craters and other permanently shadowed regions at the lunar in 2016. The objective was to determine the types of south pole. It houses a 3.4 kg and 9.6 W miniature neutron planetary science that could be done, required technology spectrometer. Similar to Lunar IceCube, LunaH-Map has developments, and expected mission costs. The missions a 14 kg wet mass and BIT-3 miniature electric propulsion concepts were limited to $100M and 180 kg. 19 proposals system, with slightly less power generation at 90 W LunaH- were selected for further study at various Solar System targets Map [16]. including five for the Moon, four for Venus, three for Small Bodies, five for Mars, and two for the outer planets. Ten NEA Scout, led by NASA MSFC, will image and charac- of the concepts are high ∆V missions that require electric terize a NEA during a slow flyby to demonstrate low cost propulsion for required trajectory maneuvers. The other nine asteroid reconnaissance capability. It is a 6 U, 14 kg CubeSat concepts conducted less demanding missions such as a single with a solar sail for propulsion and is capable of flyby of small fly-by or assumed the primary mission would deliver them to 5 goal is to investigate water and organics in the inner and outer main asteroid belt, test models for the origin and evolution of primitive objects, and provide context for previous sample return missions data with the larger asteroid population [28]. The Athena mission would be the first encounter of 2 Pallas, the last unexplored protoplanet and largest unexplored object in the inner solar system. The proposed design, shown in Figure 10, is an ESPA-class spacecraft with six Halo hall thrusters, a 1-m deployable mesh high gain antenna, and miniature color (RGB) camera. The Athena fly-by encounter is planned to happen when Pallas crosses the ecliptic plane and will comprise of 2 days of radio science and 8 hours of imaging [29].

Figure 9. Concept drawing of ESA’s standalone interplanetary CubeSat, M-ARGO, and key technologies [22]. the final destination [13]. There are two PSDS3 Small Body fly-by mission concepts, ROSS and PrOVE, and one Small Body rendezvous mission, APEX. The Primitive Object Volatile Explorer (PrOVE) is a 50 kg chemical propulsion mission concept to study the surface structure and volatile inventory of comets during flyby in their perihelion passage phase. The Asteroid Probe Experiment (APEX) is a 191 kg dry mass spacecraft concept to 99942 Apophis [23]. The spacecraft (see Figure ??) carries a 48 kg instrument suite including a deployable seismometer. The required ∆V for the mission is expected to be around 3 km/s. Figure 10. Proposed small spacecraft mission Athena, to flyby and characterize the protoplanet 2 Pallas [29]. There were two moderate ∆V missions to Mars and one to Venus. The Phobos/Deimos Regolith Ion Sample Mission The Near-Earth Asteroid Characterization and Observation (PRISM) mission concept is a 12 U CubeSat using electric (NEACO) mission is a concept study proposed to explore the propulsion for a 2 year journey to Mars, where it would orbit fast-rotating asteroid 2016 HO3, one of Earth’s few quasi- the Martian moons. The focus of the mission is to resolve satellites. It is a 166 kg standalone spacecraft using electric the question of origins of Phobos and Deimos using an ion propulsion to rendezvous with H03 within 22 months from mass spectrometer of less than 5 kg and 10 W [24]. The Mars launch [30]. The Deep Interior Scanning CubeSat (DIS- Aerosol Tracker (MAT) is a mission concept to put a 12 U, CUS) is a tandem 6 U CubeSat carrying a bistatic radar as 24 kg CubeSat in an areostationary orbit around Mars. The the main payload. It is designed to flyby and determine architecture is designed by ExoTerra Resource based on the the internal macroporosity of a 260–600 m diameter NEA Halo Hall thrusters and it contains one visible camera and and can be operated as a standalone mission [31]. The two infrared cameras [25]. Both mission assumes launch on Nanospacecraft Exploration of Asteroids by Collision and a Mars-bound trajectory, similar to Mars 2020. The Venus flyby Reconnaissance (NEACORE) mission is intended to be Airglow Measurements and Orbiter for Seismicity (VAMOS) a framework to allow for reconnaissance of a large number of conducts measurements of seismic vibrations from the air- NEAs while minimizing costs. It consist of a pair of 12 U, glow layer of the ionosphere. Architecture trade studies were 24 kg CubeSats travelling together on a multi-target flyby conducted and a 300kg SEP architecture, with the MaSMi trajectories with a flexible suit of instruments, including a thruster, was chosen for a ride share using the ESPA Grande miniaturizes time of flight LIDAR [32]. In order to better launch adapter [26]. understand the requirements and targets which are available for our mission objective, a method to analyze the available Several concept studies focusing on Small Bodies have also targets was needed. been published. Aster is a planned space mission under development by the Brazilian Space Agency (BSA) to launch a small spacecraft to rendezvous with 2001 SN263, a triple 3. PRELIMINARY MISSIONAND TARGETS asteroid system. The basic design for the Aster spacecraft is based on the Russian Pilgrin Space Platform. It has a wet ANALYSIS mass of 160 kg, payload mass of 30 kg, 71 kg of Xenon, and As of August 21, 2019 over 20,705 known NEA exist. 2 kW of power. The Aster mission spacecraft is going to be Typical mission design scenarios have a specific target or propelled by four Permanent Magnet Hall Thruster (PMHT) set of targets in mind for a science investigation based on which are currently in development [27]. their properties such as size and spectral type. For this study the inverse problem was posed—we wished to assess what The Asteroid Reconnaissance for Researching Organics and type of targets are accessible for a given mission architecture. Water in the Solar System (ARROWS2) mission is a 12 U The goal was to estimate the range of requirements of the with electric propulsion performing an asteroid fly-by. The spacecraft, starting with the most accessible NEAs. In order 6 to select a preliminary baseline set of targets and facilitate The plane change angle is ν, the change in longitude of the mission design a method of filtering out inaccessible NEAs ascending node is ∆Ω = Ωf − Ωi, the subscript i denotes by ∆V is needed. The ∆V approximations considered here the initial heliocentric orbit, the Earth, and the subscript f are to rendezvous with NEAs from Earth-escape with no denotes the final heliocentric orbit, the asteroid. For the case excess energy. where the target orbit aphelion is smaller than the initial orbit aphelion the departure and arrival burns can be switched such Rendezvous Targets Search—Calculating ∆V for purely im- that the plane change would happen at departure. pulsive trajectories can be approximated using a patched conic method with Hohmann or bi-elliptic transfers for co- The second approximation is to solve Lambert’s problem planar circular orbits. A similar method can be used to from Earth position at an initial time, re(ti), to the target adjust plane and radius for non-coplanar circular orbits. For asteroid after a given time of flight, rt(ti + tof), under only the more general case of eccentric non-coplanar orbits the the Sun’s gravitational influence. This is also a two burn Lambert’s problem can be solved. For low-thrust trajectories impulsive transfer which was solved based on the method the solution is much more challenging, requiring solving a in [34]. In order to find the optimal transfer ∆Vlm for a continuous optimal control problem. For a large population specified range of launch dates, ti, and time of flight, the it will not be possible to solve for the optimal low thrust Lambert’s problem must be solved over a discrete set of the trajectory. Instead a series of increasing fidelity approxima- possible launch times and time of flight values within the tions can be used to reduce the target pool and assess feasible launch window. The minimum ∆V transfer among all the mission options. solutions was taken as the ∆Vlm. Specifically, we wish to The first approximation is a two-burn impulsive approximation to match the aphelion, ra, perihelion, rp, inclination, i, min ∆Vlamb = f(re(ti), rt(ti + tof), tof) and longitude of the ascending node, Ω. This approximation ti, tof (9) does not take into account the phasing of the target along the s.t. tof ≤ tofmax, orbit or its eccentricity but can be calculated very quickly t ≤ t ≤ t . for a large amount of targets. The two burns are used for a i,min i i,max Hohmann-like change in orbital radius and the plane change maneuver can be split amongst the burns. It is more efficient The first approximation was used to find the total number for the majority of the plane change to take place at the larger of NEA with ∆Vapprox < 5km/s in order to reduce the orbital distance [33], however for simplicity we will assume pool of targets to a smaller number. The number of feasible all of the plane change takes place at that point. When the targets, not accounting for phasing, which met the criteria target orbit aphelion is larger than the initial orbit aphelion was 1,468. Next the minimum Lambert’s solutions were the total change in velocity can be approximated by calculated for these targets up to a time of flight of 2 years and launch window from 2022 to 2023. Afterwards, an ∆Vapprox = ∆Va + ∆Vb. (1) additional filter of ∆Vlm < 5 km/s was applied to remove any outliers that were not available in the required time of The first burn, ∆Va, takes place at the departure orbit peri- flight and launch window. The cumulative distribution of the helion in the direction tangential to velocity to increase the approximations for the entire population of known NEA is aphelion and is given by shown in Figure 11.

∆Va = Vtrans,a − Vi, (2) 104 2 where s 1.5 2 2 Vtrans,a = − µ (3) 1 rpi rpi + raf 0.5 and s 0 2 1 0 5 10 15 20 25 30 Vi = − µ. (4) rpi ai 1500 The second burn, ∆Vb, is the combined plane change and perihelion increase, deﬁned as 1000

q 2 2 ∆Vb = Vf + Vtrans,b − 2Vf Vtrans,b cos(ν), (5) 500 where 0 s 0 1 2 3 4 5 6 7 8 2 2 V = − µ, (6) trans,a r r + r af pi af Figure 11. Cumulative distribution of ∆V approximation for all known NEA (top). Cumulative s 2 1 distribution of minimum ∆Vlambert solution for all NEA Vf = − µ, (7) with ∆V approximation of less than 5 km/s (bottom). raf af and As expected there was a correlation between the ∆V and high cos(ν) = cos(ii) cos(if ) + sin(ii) sin(if ) cos(∆Ω). (8) eccentricity, semi-major axis, or inclination. The ﬁltering 7 5 process for the target population compared to their semi- U>5 U=5 major axis and eccentricity is shown in Figure 12. 4.5 U=4 U=3 U<2 1 4 V , All NEA, N=20761 3i V , V <5, N=1468 0.9 lm 3i 3.5 V , V <5 & U 3, N=227 lm lm 0.8 3

0.7 2.5 0.6 2 0.5 1.5 0.4 1 0.3 0.5 0.2 10-3 10-2 10-1 100

0.1 Figure 13. Distribution of most accessible asteroids 0 0 1 2 3 4 5 according to their diameter, uncertainty code, and ∆V .

Figure 12. Distribution of NEA populsion according to 4. LAUNCH OPPORTUNITIESAND EARTH semi-major axis, eccentricity, and approximate impulsive ∆V for rendezvous. ESCAPE Launch opportunities were assessed and are summarized in Two additional parameters of interest were the orbit condition Table 2. A rideshare as a secondary payload to Geosyn- code, U, a measure of the uncertainty of the target’s orbit, chronous Transfer Orbit (GTO) or Geosynchronous Equato- and the known or expected diameter of the asteroid. Very few rial Orbit (GEO) was the most frequent opportunity of those asteroids have known size. Assuming a spherical object with considered. Over 30 launches to GTO are performed annually no albedo variation their diameter in km can be estimated by for telecommunication satellites, many of which have excess payload capacity. Rideshares to GEO can be facilitated 3.1236−0.5 log (a)−0.2H through the SSL Payload Orbit Delivery System (PODS) d = 10 10 , (10) with up to 8 opportunities annually. Rideshare to GTO or GEO would be the most cost effective and most available. where a is the geometric albedo and H is the absolute However, it increases the overall mission ∆V significantly. It magnitude [35]. For known asteroids an estimate for H exists would require an additional 3.2 km/s or 1.4 km/s to escape from ground measurements but albedo is typically unknown. Earth orbit with low-thrust using a lunar flyby or chemical However, we can simply use an assumed mean albedo of propulsion system respectively. 14% to estimate all unknown NEA sizes. It can be seen that the majority of the most accessible targets have a very Table 2. Overview of different launch opportunities small diameter (less than 100 m) and vary widely in condition available to small spacecraft including typical size limits. code. The corresponding distribution is shown in Figure 13. The most interesting of these targets tend to be those with diameters > 50 m and it is typically desirable to focus on Orbit Frequency Cost Size Limit targets with U < 4, which was the subsequent focus of this Secondary Payload study. GTO/GEO Monthly Low ESPA Grande While these estimates are for impulsive maneuvers they give Lunar Annually Moderate CubeSat insight into the low-thrust case. The low-thrust ∆V is HCO Indirect Annually Moderate ESPA more complicated to approximate as it can vary depending on the target, thruster performance, and active constraints. HCO Direct Rarely Moderate ESPA From similar studies found in literature the accuracy of these Primary Payload estimates can be estimated. A study of N = 143 NEA targets HCO Direct On Demand High None for a small spacecraft design calculated the mp required to rendezvous based on different approximation methods [36]. From this data it is found that the correlation coefficient is Several launch opportunities as a secondary payload on a 0.59 and 0.81, mean error is -13% and -1%, and standard lunar mission will become available in the coming years deviation is 19% and 15% for mp using ∆Vapprox and ∆Vlm under the Artemis program and Commercial Lunar Payload methods compared to the “true” optimized low thrust solu- Services (CLPS) launches. These secondary payloads are tion. These results are mission specific, but they indicate that currently restricted to CubeSat form factors. From trans-lunar the impulsive approximations usually undershoot the required injection (TLI) it will be possible to perform an Earth-escape propellant and have a 3-σ of 45% and above compared to the maneuver for very little ∆V . For both geosynchronous and true ∆V budget. Based on these results it was expected that lunar launches it is possible to use a lunar flyby to reduce a ∆V of at least 3 km/s per second would be required from escape ∆V and achieve a hyperbolic escape energy C3 of up Earth escape to reach larger bodies and provide robustness to to 2 km2/s2 [37]. Another strategy is to use a geosynchronous other mission parameters. launch and receive a “perigee kick” to enter a highly eccentric 8 orbit with apogee at near-lunar distances (similar to a TLI). in red indicating their lack of feasibility. Given the current The perigee kick could come from an additional launch technologies in development it is not expected to possible vehicle kick stage motor, propulsive ESPA rideshare adapter, to achieve the required mass fractions for these in a 180 kg or chemical propulsion system onboard the spacecraft. or smaller spacecraft. Of the three options which seemed technically feasible the best option was to launch on a ride Opportunities also exist on launches to indirect heliocentric share to near Earth-escape (C3 ≈ 0) since there is much orbits (HCO). This refers to any Earth-escape launch such higher availability for this type of rideshares. This is only as transfer orbits on missions to other planets, Small Bodies, achievable with electric propulsion, therefore that was chosen and Sun-Earth Lagrange points. These opportunities are less as the emphasis of the remainder of the trade study. Beyond frequent and vary in direction of the outgoing hyperbolic feasibility, it is also found that electric propulsion will expand asymptote which may or may not be beneficial to the sec- the number of accessible NEA compared to chemical propul- ondary payload. sion for a given spacecraft mass. High-level Mission Architectures Based on the assessment of launch opportunities and target 5. ENABLING SUBSYSTEMSAND accessibility, six architectures were assessed and are shown TECHNOLOGIES in Table 4. The architectures are combinations of chemical or low-thrust electric propulsion system and launching from This section reviews the selection and performance of technologies that can enable an autonomous small spacecraft either GTO, C3 ≈ 0 (Lunar or HCO indirect), and HCO direct. In order to assess the architecture feasibility, the ∆V mission to rendezvous with Small Bodies. Several small for each was estimated based on the pool of nearby NEA spacecraft technology reviews exist [2], [38], however they targets that were assessed. Based on the expected specific vary in scope are relevance to interplanetary missions. Em- impulse and ∆V , the minimum required propulsion mass phasis is on technologies with a Technology Readiness Level fraction, m /m , was calculated. Finally, the corresponding (TRL) of 5 or more which can enable the mission objectives p 0 for ∆V of 3 km/s or more, time of flight of 2 years or more, required maximum dry mass, md, was calculated assuming a 180 kg wet mass spacecraft. communication to Earth up to 1 AU away. A summary of the overall performance envelope expected, based on these technologies, is provided in Table 3 and will be discussed Table 4. Overview of architectures available for a small throughout this section. spacecraft to rendezvous with accessible NEAs. Required md is shown for a 180 kg wet mass. Propulsion Propulsion system options can be divided into three cate- Launch ∆V (km/s) Req. mp/m0 Req. md (kg) gories: cold gas systems, chemical propulsion, and electric SEP, Isp = 2000 s propulsion. For small maneuvers (∆V < 50 m/s) and attitude control, cold gas has been well established due to it’s sim- GTO 6 0.26 133 plicity and heritage on a range of small spacecraft including C3 ≈ 0 3 0.14 154 CubeSats. For moderate changes in velocity, several chemical Direct 1.5 0.07 167 propulsion options exist. The most mature include solid rocket engines and liquid monopropellant systems. Chemical, Isp = 250 s GTO 3 0.71 53 For large changes in velocity (∆V > 2 km/s), electric propul- C ≈ 0 2 0.56 80 sion is required due to the higher specific impulse. Types of 3 electric propulsion include electrospray, Hall Effect thrusters Direct 0.5 0.18 147 (HET), ion propulsion systems (IPS), pulsed plasma, resistojet, and vacuum arc thrusters. All these systems are active Of the six architectures, three architectures are highlighted in development, testing, and technology demonstration for a

Table 3. Summary of key performance envelopes in near-term based on subsystems in development.

Size 6U 12U Mid-ESPA ESPA Wet Mass (kg) 12-14 20-28 40-100 150-180 Prop Mass (kg) 1-2 1-4 7-15 10-40 Speciﬁc Impulse (s) 1000-2300 1000-3000 1200-3200 1500-3200 Thrust (mN) 0.5-1.5 0.7-6.5 10-33 20-54 ∆V (km/s) 1-2.5 1-3 2-4 3-6 Power BOL 1AU (W) 60-120 100-250 200-1000 500-2000 Array Power Density (W/kg) 60-100 60-100 80-120 80-120 Payload Mass (kg) 0.5-2 0.5-3.5 1-15 2-30 Payload Power (W) 1-6 1-20 5-50 10-150 Downlink 1AU (kbps) 1-8 1-8 1-15 1-45 90-day Data Volume (GB) 0.3-2.6 0.3-2.6 0.3-4.9 0.3-15 1-σ Pointing Accuracy (◦) 0.004-0.1 0.004-0.1 0.003-0.01 0.003-0.01

9 range of small spacecraft sizes. Hall thrusters and ion engines Table 5. Specifications of electric propulsion systems are the most mature options for high ∆V . Resistojets, applicable to a range of interplanetary small spacecraft electrospray, and pulsed plasma thrusters can be applicable sizes for the design space considered. for more efficient alternatives to cold gas systems. Component T (mN) Isp (s) P (W) Bus To assess the performance of technologies available and in development they can be compared by thrust and specific BIT-3 1 2150 80 6U-12U impulse, shown in Figure 14 [38], [39], [40], [41]. It is BHT-200 13 1390 200 12U-ESPA preferred to have an engine with high specific impulse in Halo 33 1500 450 12U-ESPA order to maximize ∆V for a given propellant mass ratio. It is also important to have sufficient thrust to ensure the XIPS-13 18 2350 450 ESPA spacecraft can perform the maneuvers in a reasonable amount BHT-600 25 1560 600 ESPA of time. T5 20 3000 700 ESPA MaSMi-60 54 1940 1280 ESPA 104 Green Monoprop PPS-1350 89 1650 1500 ESPA Solid Rocket Cold Gas 2 Electrospray 10 Hall Thruster Ion Thruster Pulsed Plasma/Vacuum Arc and xenon, it is planned for use on a range of small spacecraft Resistojet 0 [42]. The XIPS-13 xenon IPS using which has flown on 10 17 or more Boeing 601 HP satellites [43]. The T5 xenon IPS has previously flown one the GOCE mission and is being developed further by the European GIESEPP program 10-2 [44]. The MaSMI-60 xenon HET and ASTRAEUS system are being developed for high throughput interplanetary small spacecraft missions [45]. The PPS-1350 xenon HET, based 10-4 on the SPT-100, was used on the SMART-1 mission from GTO to the moon and two geostationary satellites, Inmarsat- 4A F4 and Hispasat AG1 [46].

10-6 2 3 4 These propulsion systems are meant to be representative 10 10 10 options for a variety of interplanetary small spacecraft sizes based on their performance. As a spacecraft increases in Figure 14. Comparison of small spacecraft propulsion mass a higher thrust is needed, as a result thruster power technologies according to thrust and specific impulse. input increases which is accommodated by larger solar arrays. The specifications shown are typically for the maximum or nominal operating point and many of these systems Preliminary selection of electric propulsion systems can be filtered based on total system efficiency, power input, lifetime can accommodate additional throttle points. Finally, newer technologies such as the microfluidic electrospray propulsion propellant throughput, system dry mass, system volume, sys- (MEP) by JPL and field-emission electric propulsion (FEEP) tem TRL, and availability. In order to have a fair comparison ∆V of different options, the systems are considered to contain by Enpulsion also have significant potential for high if the thruster, power processing unit (PPU), and flow control they can be scaled up to the required throughput. To select unit (FCU). The propellant tank is scalable and sized to the a single propulsion system from these options for a given mission, low-thrust trajectory optimizations is required to application, and gimbal is optional but likely necessary. compare overall mission impact. The propellant mass throughput is particularly important as Communications many electric propulsion systems have limited lifetimes due to degradation. For a 3 km/s mission requirement using a The communications system is required for command uplink, 1200 s to 3000 s propulsion system, a required propellant telemetry downlink, and spacecraft navigation. The main throughput would be needed of of 1.4 kg to 3.1 kg for a components are the antennas and transponder or radio. The 14 kg spacecraft and 17 kg to 40 kg for a 180 kg spacecraft. two radios flown on interplanetary small spacecraft include Maximum power available to the thruster based on solar array the 15 W output PROCYON X-Band radio and 4 W output technology would be limited to around 120 W, 250 W, and X-band IRIS V2 radio flown on MarCO. Small radios flown 2000 W for 6 U, 12 U, and 180 kg spacecraft respectively. on larger missions, such as the Frontier radio may also be From a catalog of propulsion systems, 8 promising options suitable. Radios in development which are applicable to small were selected which cover a range of spacecraft sizes, shown spacecraft include Universal Space Transponder (UST) lite, in Table 5. Thrusters which were not selected was due to IRIS V2.1, and Frontier lite. The specifications of three of uncertainty in their current status, low TRL, low throughput, these options are shown in Table 7. lack of availability, concerns with reliability, or simply lower overall performance relative to similar options. Multiple antennas will be needed onboard the spacecraft. A low-gain antenna (LGA) is required for near-Earth communi- The BIT-3 is being developed for the Lunar IceCube and cations, a medium-gain antenna (MGA) for safe-mode com- LunaH-Map missions expected to launch in 2020 [15]. The munications far from Earth, and a high-gain antenna (HGA) BHT-200 using xenon has flown on the TacSat-2, FalconSat- for downlink back to Earth over large distances. A patch 5, and FalconSat-6, an iodine variant is expected to fly on antenna and small patch array can used for the LGA/MGA iSat, and a larger version (BHT-600) is under development structure, as was the case for MarCO. The accommodation [38]. The Halo HET is being developed for use with iodine of a HGA is more difficult since the larger aperture requires 10 Table 7. Three interplanetary small spacecraft radios by using a higher frequency band or more powerful radio options and their specifications [50] [51] [52]. and traded for other spacecraft capabilities. For example, by incorporating Ka-band transmission capability the data rate Specification Iris V2.1 Frontier UST lite can be roughly tripled (including additional losses incurred for pointing, atmospheric, receiver performance). Frequencies X S/X/Ka X/Ka Mass (kg) 1 2.1 3 Power Volume (cc) 600 2050 2700 The electrical power subsystems (EPS) can be divided into Rx Power (W) 12 5 30 three main functions: power generation, energy storage, and power management and distribution (PMAD). Power gener- Rx Thres. (dBm) -151 -160 -160 ation is limited by the solar cell efficiency and deployable Uplink (sps) 63-8k 1-1M 7.8-37.5M array structures. State of the art in deployable rigid solar Downlink (sps) 63-6.1M 10-150M 10-300M arrays for small spacecraft is currently between 70 W/kg to TID (krads) 23 100 300 120 W/kg. The MMA eHaWK which will fly on several 6 U lunar CubeSats can generate up to 112 W and 121 W/kg [53]. The MMA rHaWK is also in development for ESPA-size spacecraft and can generate up to 2kW and 130 more space and mass onboard. W/kg. Deployable flexible solar arrays are lower TRL but over 150 W/kg, such as SolAero Composite Beam Roll-Out The HGA antenna size is restricted to a body-mounted patch Array (COBRA) for small spacecraft which can generate array or deployed antenna. Four options are shown in Table 6, 1 kW at 170 W/kg [54]. Further major advances in solar the aperture size and area listed for the patch array and the array performance are envisioned in the near-term leading reflectarray are the maximum allowable for those designs to 150–200 W/kg and in the mid- to far-term up to 200–250 given a 12 U CubeSat size. The patch array and reflectarray W/kg [55]. both have heritage in deep space, the 0.5 m mesh antenna has flown in LEO, and the 1 m mesh reflector is under Current energy storage systems use lithium ion (Li-ion) or development. lithium polymer (LiPo) rechargeable cells. Different models qualified for space vary in specific energy from 140 to 250 Wh/kg [38]. The cylindrical 18650 Li-ion cells, have a Trade offs for communications design are typically between > data rate to earth, power required, and heat generation. high specific energy ( 200 Wh/kg) and energy density (600 Achieving higher data rates requires more power (larger solar Wh/L) along with excellent cycle life, standard form factor, panels) and more complex thermal management solutions. and offers internal safety feature options [56]. Notable 18650 The data rate will be dependant on several factors such as the cell designs are the VTC4 and LG HG2 by Sony, and the ground antenna used, the spacecraft antenna, the spacecraft NCR18650A and NCR18650G from Panasonic with specific radio, and distance from Earth. For example, a small space- energy of nearly 250 Wh/kg. PMAD systems control the flow craft on a mission to Mars or NEA could be at a distance of power to the spacecraft which are often custom designed to of around 1 AU on arrival. Using the 34m DSN antenna the mission but several COTS options have become available with the IRIS V2.1 X-band radio at 1 AU for the options for CubeSats. Several manufactures offer standard products in Table 6 would give approximate data rates of 1 kb/s, 2 for small spacecraft which include Pumpkin, GomSpace, kb/s, 4 kb/s, and 14 kb/s. Assuming a notional mission with Stras Space and AAC-Clyde. 90 days of downlink at 8 hours per day this corresponds to 0.3, 0.6, 1.2, and, 4.2 GB respectively. This estimate represents a minimal baseline data rate and could be increased

Table 6. Summary of four potential HGA options applicable to interplanetary small spacecraft [47] [48] [49].

Specification Patch Array Reflectarray Deployable Mesh Reflector Size 20×30 cm 60×34 cm 0.5 m dia 1 m dia Area 680 cm2 2040 cm2 1964 cm2 7854 cm2 Gain 24 dB 29.2 dBi 42.6 dBi 49.2 dBi Stowage 0.2 U 0.8 U 1.5 U 3 U Deployment None Hinge and unfold 30 hinged ribs Boom deployed Mass 2 kg 1 kg 1.4 kg 2 kg Heritage Various MarCO Rain Cube None TRL 9 9 7 5

11 The Guidance, Navigation & Control (GN&C) The GN&C subsystem includes the components used for position estimation and attitude determination and control system (ADCS). Miniaturized technologies available for ADCS have been proven (TRL 9). Several integrated units are available which combine many sensors together, such as the the BCT XACT which includes a star tracker, gyro, coarse sun sensors, and 3-axis reaction wheels. This system was ﬂown on MarCO and is planned for several upcoming lunar CubeSat missions. It provides 0.003◦ to 0.007◦ 1-σ pointing accuracy, 6 to 40 arcsecond 1-σ attitude knowledge, and 10 deg/s slew rate [57]. For reaction wheel desaturation outside of Earth’s magnetic ﬁeld magnetorquers cannot be used, instead a gimbal on the primary propulsion system or additional reaction control thrusters (cold gas, chemical, or electric propulsion) is required.

During interplanetary cruise the position of the spacecraft in the solar system must also be estimated. This is typically done through tracking sessions with deep space network (DSN) antenna using ranging, Delta-DOR, and Doppler. Figure 15. Estimate of 90th percentile σpos for camera Onboard autonomy can be used to reduce DSN tracking options using AutoNav at 1 AU based on IFOV of camera required through autonomous navigation (AutoNav) which and max apparent magnitude, adapted from [59]. can be performed using only optical measurements of distant asteroids which was first demonstrated on Deep Space 1 [58]. Once the target is visible, it is also possible for autonomous target-relative navigation to meet precise relative positioning as it is used to manage complex on board computation and demands required for Small Body proximity operations. operation. A miniaturized, radiation-tolerant avionics such as the JPL Sphinx LEON3-FT or radiation-hardened RAD750 is A comparison was done using several baseline camera mod- typically used. However, significantly more of computational els to ensure adequate performance of AutoNav. A kinematic power is needed for intensive autonomy functions, which approximation can be used to estimate position accuracy require modern high performance computing system. This achievable with AutoNav based on the triangulation of sev- is possible using high performance COTS cell-phone and eral visible bodies with known orbits [59]. Two particular automotive-grade processors with fault handling and redun- camera parameters which affect the performance of AutoNav dancy, similar to the Mars Helicopter architecture using the are the instantaneous field of view (IFOV) of one camera Qualcomm SnapDragon and ARM Cortex. A distributed pixel, Θ, and the maximum visible apparent magnitude de- architecture will be needed to separate basic mission critical C&DH tasks from advanced autonomy tasks. tectable, Mmax. While Θ is fixed based on choice of camera, Mmax is influenced by several factors such as pointing sta- bility, the exposure time, and image processing techniques. Building up a fully autonomous capability to access and operate on Small Bodies is a paradigm shift from the cur- Two representative baseline camera systems in development rent approach, which have limited autonomous capability are the IntelliCam are for NEA Scout [19] and the mini- (Figure 16). To achieve this, several autonomy-enabling advanced pointing imaging camera (mAPIC) narrow angle onboard algorithms would be needed. These may include camera [60]. Based on the camera specifications, near 1 interplanetary cruise autonomous navigation (AutoNav), op- AU from the Sun the IntelliCam and mAPIC may be able tical target-relative navigation once the target is visible, and a to achieve a mean σ 6000 km and 900 km respectively multi-phased navigation and mapping algorithm sequence on pos approach as shown in Figure 2 [4]. Agile science algorithms and 90th percentile σpos of 20,000 and 4,000 km, shown in Figure 15. This performance is worse than traditional DSN (ASA) could be used for onboard science event detection and radiometric tracking. However if AutoNav target-relative response [8]. Automated instrument operation to improve position uncertainty is much less than the distance of first image quality and size and onboard algorithms to prioritize target visibility, then a trajectory correction maneuver (TCM) data for downlink will be required. can be made with minimal impact relative to the overall mission ∆V budget. As long as Small Bodies with low To achieve responsiveness in a dynamic Small-Body environ- condition codes are chosen both camera models would be be ment systems for continuous planning and command execu- sufficient. tion across different activities will be needed. For example, MEXEC (Multi-mission Executive) is a flight software for Command and Data Handling (C&DH) activity scheduling and execution which can be scaled to more complex task definition and state models, allowing for The C&DH system performs two primary functions. It more robust autonomous behaviours. receives commands then it validates, decodes, and distributes them across subsystems. It also also gathers, processes, and Structure and Thermal formats data collected for housekeeping, operations, or mission objectives which may be used onboard or downlinked. The current practice for primary structures is typically custom It also includes timekeeping, health monitoring (watchdogs), and mission-specific for interplanetary small spacecraft. Tai- and other interfaces. lored solutions are being developed to meet specific mission requirements by companies such as ExoTerra. These may The C&DH system is critical for an autonomous spacecraft eventually result in attempts to establish standardized plat- forms similar to what has occurred in LEO. Structure mass 12 Figure 16. The limited autonomy framework today compared to an end-to-end autonomy framework. can be estimated empirically based on previous missions. For 6. SPACECRAFT ARCHITECTURE TRADE small spacecraft missions this typically ranges from 15-25% SPACE with 20% being typical [61]. Trade studies and systems engineering analyses were con- Thermal management is used to regulate the spacecraft tem- ducted as part of the architecture trade space exploration. A perature range to within that required of the subsystems. The range of spacecraft sizes were explored from 14 kg to 180 most commons solution is patchwork passive radiator pan- kg. A top-down approach was used initially, starting from an els, spacecraftcoatings, multi-layer insulation, and electrical overall size, mass, and power expectation and selecting sub- heaters. These solutions are highly specific to the spacecraft systems which would fit within it. The specific subsystems layout, payload, and operations required. Technologies such and technologies were selected from catalogs of those with as cryocoolers, sunshades, passive louvers, and others also flight heritage or in development as discussed in Section 5. In also being developed. some cases where specific products were not available or they should be tailored to the mission, reasonable assumptions Thermal challenges are particularly difficult for interplane- were made based on previous mission data and current state tary CubeSat missions. Limited surface area for avionics and of the art. communications transponder heat dissipation and providing replacement heat in low-power modes at large solar distances The architecture design space of 14 kg to 180 kg was dis- are the largest difficulties. In the case of the lunar CubeSat cretized into a select few baseline designs of interest. The missions in development, thermal duty cycling, particularly chosen wet mass values for the designs were 14 kg, 24 kg, for the communications system, was required as an additional 100 kg and 180 kg. The current best estimate (CBE) of measure. subsystems was used to create a mass budget. For contin- gency, mass growth allowances (MGA) were adapted from Science Instruments [62] for different categories of components and mapped from the maturity levels specified to a TRL range. The MGA Miniaturized small spacecraft instruments can achieve a were used to estimate the maximum expected value (MEV) broad range of scientific objectives in a compact low-power of mass for each system. An additional margin was also package. Investigations could be aimed at elemental and incorporated, where the maximum possible value (MPV) was mineralogical composition of the surface along with physi- set to the upper limits of 14 kg, 24 kg, 100 kg and 180 kg. cal characterization of regolith. In addition, they could be The overall dry mass margin from CBE to MPV achieved directed at addressing strategic knowledge gaps for human was between 20% and 30% for all the designs. Power budgets exploration by quantification of the dust dynamics, subsur- were also created and a flat 20% margin was applied to power face water, and surface electrostatic charging generation. A reference payload could include a multispectral camera, In- The propulsion system was a primary driver of the designs telliCam visible imager developed for the NEA Scout Cube- and was typically the first technology to be selected. Other Sat, small alpha particle and X-ray spectrometer (APXS) technologies selected were based on performance for a given which was flown on the Philae comet lander, and the Mi- spacecraft size and heritage. The solar arrays were sized to crOmega microscope developed for the Phobos-Grunt mis- the power budget which was driven by the limit operating sion. Each of these instruments weigh less than 1 kg and mode during thrusting. As a result over 90% of the power require less than 3 W peak power. Additional instruments budget was driven by the propulsion and communications could also include a more capable high-resolution camera, system. A 6 U, 14 kg design was first explored drawing on radio wave sounder, or magnetic, electron and ion detectors. subsystems in development. The BIT-3 electric propulsion systems was selected which is capable of up to 2.3 km/s on a 14 kg spacecraft assuming maximum power of 80 W [15].

13 The major deviation from other designs, such as the Artemis- Since xenon is stored as super critical pressurized gas the 1 6U CubeSat missions, was the necessity to accommodate mass per volume is not fixed but the optimal design is around an additional antenna for downlink of data at distances up to 1.4 kg/L [63]. 1 AU. A MGA patch array was chosen which could either be body mounted or deployed similar to the solar arrays. To assess the impact of payload mass on ∆V , we can rear- range equation 11 for The second baseline developed was a 12 U, 24 kg spacecraft. Many of the same technologies from the 14 kg could be m0 − mb − mpl mp(mpl) = (12) incorporated, such as the C&DH, transponder, EPS, batteries, 1 + mt and LGA/MGA. However, components such as the reaction mp wheels, structures, thermal, wiring, and solar arrays must be sized up. The propulsion system chosen was two BIT-3 and substitute it into the ideal rocket equation thrusters due to the lack of other higher performance options available at this power level. For the 100 and 180 kg designs   the propulsion design space opens up significantly. The m0 ∆V (mpl) = g0Isp ln   . (13) Halo thruster, XIPS-13, T5, BHT-600, and MaSMi all were m0−mb−mpl m0 − 1+ mt feasible within power generation constraints. The optimal mp choice could not be made without in depth trajectory analysis. Regardless, the MaSMi was selected as a baseline option due This further simplifies to to its higher throughput capacity which may be needed. A summary of these four baseline designs are shown in Table 8. mt ! m0(1 + m ) ∆V (m ) = g I ln p . (14) pl 0 sp mt Table 8 Selected design alternatives for trade studies. m0 + mb + mpl . mp

Alt. Design A B C D and can be plotted from 0 < mpl < m0 − mb for each of the different baseline architectures as shown in Figure 17. Form Factor 6 U 12 U ESPA ESPA

Wet Mass (kg) 14 24 100 180 102 14 kg 100 kg APEX Power BOL 1AU (W) 120 240 1100 2000 24 kg 180 kg Primary ∆V (km/s) 2.39 2.82 3.50 4.78 228kg NEACO Payload (kg) 1 1.5 5 10 165kg Dry Mass (kg) 12.5 21 80 140 Specific Impulse (s) 2150 2150 1800 1940 1 Primary Prop. Mass (kg) 1.5 3 20 40 10 VAMOS 260kg Once the baseline designs were selected, further trades be- NEACORE 24kg PRISM tween subsystems could be made to assess the sensitivity 27kg to certain design changes. For example more payload mass can be traded for less spacecraft capability. Typically that DISCUS can be done by reducing propellant, resulting in less ∆V 0 13kg M-ARGO and fewer reachable targets, reducing communication system 10 22kg size, resulting in less science data returned to Earth, or decreasing the size of the ADCS, resulting in less accurate 0 1 2 3 4 5 6 7 attitude knowledge or pointing. The trades can also be made in other directions, such as less payload mass for more spacecraft capabilities or trading one spacecraft capability for Figure 17. Trade offs between payload mass and another, within physical limits such that the design will still spacecraft ∆V capability for the four baselines (solid close. lines) across the design space and compared to known mission studies (dots). The primary trade identified for the Small Body rendezvous mission considered here is between payload mass and ∆V . By assessing the payload mass against ∆V trade offs we can To explore this trade we can fix all other design variables see the architecture performance across the design space. For and only vary these two with respect to each other. More the design step increment from 6 U to 12 U, capability in- specifically the wet mass of the spacecraft, m0, is fixed and creases from 2-3 km/s to 3-4 km/s with a small payload (0.5- can be represented by 1 kg), while still benefiting from staying within the CubeSat standardized deployer. Considering that several of the closest mt m0(mp, mpl) = mb + mp(1 + ) + mpl, (11) NEA require around 3 km/s for low thrust rendezvous, the mp 12U architecture will enable access to many more targets than 6U. The next step up to 100 and 180 kg moves away from where mb is a fixed baseline mass, mp is the propellant mass, the CubeSat standard but allows for substantially increased mt/mp is the tankage fraction, mpl is the payload mass. The capability. Within the 4-5 km/s range, many more interesting propulsion system can be scaled by increasing the propellant NEAs become available for rendezvous, along with other mass and tank size, up to the maximum propellant throughput Small Bodies and the Martian moons, Phobos and Deimos. of the thruster design. Depending on the propellant used the For this reason, 12U and 180 kg were both assessed to be tankage fraction will change, it is generally between 5% and promising options that represent distinct capabilities, cost, 10% for pressurized xenon and much less for solid iodine. and risk posture. 14 7. REFERENCE MISSION DESIGN The flight systems have all the necessary components to com- In order to further evaluate technology risk for system per- plete an interplanetary small spacecraft mission successfully. Both propulsion systems are equipped with 2-axis gimbals formance, a notional point design was necessary. The ability capable of reaction wheel desaturation. The 180 kg space- to perform the objective of low-thrust rendezvous can only be verified by low-thrust trajectory optimization and mission craft has an additional monopropellant RCS system which is capable of high thrust trajectory correction maneuvers and analysis. Therefore a set of reference designs are established proximity operation burns. An X-band Iris radio is used to confirm initial feasibility and as a baseline for more de- tailed studies with more refined requirements. which is based on a model with previous flight heritage on MarCO. Patch antennas and deployable high-gain antennas Baseline Spacecraft Architectures are used which are capable of downlink to the Deep Space Network in a compact package. The attitude control system From previous analysis the two most promising architectures includes three reaction wheels for fine pointing, an IMU, were the 12 U, 24 kg and ESPA-class, 180 kg spacecraft. A star tracker, and coarse sun sensor. While a specific science mass summary for the entire notional spacecraft systems can instrument is not used in the design, an allocation is made for be seen in Table 9 and Table 10. The mass summaries reflect a a miniaturized instrument suite representative of applicable margined system with contingencies applied to each subsys- instruments currently being developed. tem to account for mass growth from this early stage design. The achieved total margin was higher for the larger spacecraft Trajectory and Mission Analysis due to lower TRL components used and less constrained mass To access a range of available NEAs, the spacecraft will limits. launch on a near Earth-escape trajectory. This may include Table 9. Notional 24 kg flight system mass summary. a ride share to lunar orbit, Sun-Earth Lagrange point, another heliocentric mission, or GTO with a large apogee kick from a high-thrust propulsion system. From the launch orbit a small Subsystem CBE (kg) MGA (%) MEV (kg) maneuver can be used to reach Earth-escape with C3 = 0 or 2 2 if a lunar gravity assist is available up to C3 = 2 km /s is Structure/Mech. 3.4 14% 3.9 possible within two months [37]. This range of options was Thermal 0.3 15% 0.4 considered to remain flexible to different available launch op- Power 4.8 14% 5.5 portunities. From Earth-escape, with little to no characteristic Telecom. 2.4 7% 2.6 energy, the spacecraft will begin an interplanetary low-thrust transfer to rendezvous with a NEA. C&DH 0.3 20% 0.4 GN&C 1.2 5% 1.3 Two routes available for the mission trajectory were focused Propulsion 2.8 10% 3.1 on, either direct from Earth to asteroid (EA) or returning to Earth for gravity assist before rendezvous (EEA). The Spacecraft Bus 15.2 12% 17.0 advantage of the EEA route is that it reduces ∆V and required Payload 1.5 10% 1.7 fuel consumption, however it significantly increases the time Dry Mass 16.7 12% 18.7 of flight of the mission. Several targets were assessed with the spacecraft design requirements specified in Table 11. Margin 26% MPV Dry Mass 21.0 Table 11. Established requirements for trajectory optimization and mission design. Propellant 3.0 Wet Mass 24.0 Requirement 12 U Design ESPA Design Wet mass m = 24 kg m = 180 kg Table 10. Notional 180 kg flight system mass summary. 0 0 Propellant mass mp ≤ 3 kg mp ≤ 30 kg Thruster power (1 AU) P ≤ 180 W P ≤ 1,500 W Subsystem CBE (kg) MGA (%) MEV (kg) t t Time of flight tof ≤ 2 yrs Structure/Mech. 25.0 14% 28.5 2 2 Characteristic energy 0 ≤ C3 ≤ 1 km /s Thermal 3.0 15% 3.5 Engine duty cycle D ≤ 90% Power 29.8 19% 35.5 Telecom. 3.1 7% 3.3 C&DH 0.6 20% 0.7 A broad target search based on patched-conic impulsive GN&C 3.5 5% 3.7 Lambert’s approximation was used to initialize the trajectories. The targets were limited to NEAs with estimated or Propulsion 25 10% 27.5 real diameters of greater than 100 m. Next an impulsive Spacecraft Bus 90.0 14% 102.7 trajectory optimizer was used to converge the solutions from the patched-conic approximation for each target. The time- Payload 10.0 10% 11.0 of-flight and launch C3 constraints were relaxed slightly as Dry Mass 100.0 14% 133.7 needed during several refinement stages. The converged Margin 35% solutions from the impulsive optimization were turned into a low-thrust electric propulsion format and optimized for flyby MPV Dry Mass 135.0 and then again for rendezvous. Prop (RCS&HET) 45.0 The results showed that the number of available NEA which Wet Mass 180.0 met the set criteria was much higher for the 180 kg ar- 15 Table 12. Selected trajectory solutions to 99942 Apophis. chitecture. The C3 constraint was particularly limiting for the spacecraft to reach targets within the time of flight and propellant throughput limits. The time of flight and propellant Baseline 24 kg 180 kg mass constraints, were more likely to be active and limiting for the 24 kg trajectories. This was primarily due to the Route EA EA EEA EA EEA 2 2 limited specific thrust available to the 14 kg design of only C3 (km /s ) 1 2 1 2 0 0.14 mm/s2 whereas the 180 kg design had more than twice tof (yrs) 1.46 1.39 2.42 0.43 1.23 2 as much, 0.3 mm/s . The propellant throughput for the 24 m (kg) 3 2.7 2.7 25.44 26 kg design, based on the thruster limits, was also an issue p and if it could be raised that would increase the performance. de (au) 0.6 0.5 0.6 0.4 0 The distance from earth typically remained under 1 AU and was not an active constraint. After a sample of different targets were assessed, one was picked for a reference mission design to evaluate performance. The chosen target was 99942 Apophis and the parameters of select trajectory cases are shown in Table 12, Figure 18, and Figure 19.

2 2 (a) Trajectory overview with C3 = 0 km /s , tof=1.3 yr, mp=28 kg.

2 2 (a) Trajectory overview with C3 = 1 km /s , tof=1.5 yr, mp=3.0 kg.

(b) Range from Sun, Earth, and Target during the mission.

Figure 19. Trajectory design for 180 kg SEP architecture to reach 99942 Apophis with Earth gravity assist.

(b) Range from Sun, Earth, and Target during the mission.

Figure 18. Trajectory design for 24 kg SEP architecture to propellant mass. Several options are available to increase the reach 99942 Apophis with direct from Earth. pool of targets which include increasing the time of flight, propellant mass, and assessing different propulsion system Based on this mission analysis it was found that trajectories options. It is possible to relax some of the constraints but could be converged to several NEA targets. However, as ex- it is dependant on the specifications of the propulsion tech- pected the trajectories were very constrained and many were nologies as they mature, and reliability of small spacecraft not able to meet the required constraints with the allocated for long duration missions, which are uncertain. 16 8. CONCLUSION mission analysis, and Kieran Carroll for discussing his find- In this paper we proposed a more affordable approach to ings from similar mission studies and the overall approaches characterize Small Bodies in detail using a small spacecraft for small spacecraft design. with increased onboard autonomy. We presented a review of relevant mission concepts and selected necessary subsystem This research was carried out at the Jet Propulsion Labora- technologies which have already been demonstrated in space tory, California Institute of Technology, under a contract with (TRL 7 or higher) or are expected to be in the near term (<3 the National Aeronautics and Space Administration. This years). The architecture trade space analysis and reference work was supported by the Canadian Aeronautics and Space Institute, McGill Engineering Student Centre, and McGill mission design explored the range of mission capabilities Institute for Aerospace Engineering. within the design space. The mission analysis indicated the feasibility of enabling Small Body rendezvous missions with a standalone small spacecraft. The mission places the small spacecraft in high-Earth orbit, Lunar orbit, or low- REFERENCES energy Earth-escape trajectory and is followed by a low-thrust [1] NASA/GSFC. Launch of first u.s. spacecraft interplanetary transfer using electric propulsion. It was found to sample asteroid set for september, nasa that a 12 U, 24 kg architecture would be capable of delivering media accreditation open. [Online]. Available: a 1.5 kg payload to a few interesting nearby NEA candidates. https://www.nasa.gov/press-release/launch-of-first-us- It was also found that an ESPA-class, 180 kg small spacecraft spacecraft-to-sample-asteroid-set-for-september-nasa- could deliver 10 kg or more of payload to a larger population media of NEAs with a more robust and capable architecture. [2] B. Lal, E. de la Rosa Blanco, J. R. Behrens, B. A. The primary critical paths identified for the architecture are Corbin, E. Green, A. J. Picard, and A. Balakdrishnan, in the availability of miniature electric propulsion systems, “Global trends in small satellites,” Science & Technlogy high-efficiency large deployable solar arrays, and autonomy- Policy Institute, Washington, DC, 2017. enabling algorithms. All of these technologies are actively [3] E. National Academies of Sciences, Medicine et al., in development or planned for demonstrations in upcoming Achieving science with CubeSats: Thinking inside the missions. This work leaves several questions open for further box. National Academies Press, 2016. investigation. First and foremost, is performing a broad low- thrust trajectory analysis to quantify the number of NEA that [4] S. Bandyonadhyay, I. Nesnas, S. Bhaskaran, B. Hock- are accessible to the 12U and ESPA-class architectures. This man, and B. Morrell, “Silhouette-based 3d shape recon- can help answer the question of the size of the total accessible struction of a small body from a spacecraft,” in 2019 pool of small bodies with this approach, and what their IEEE Aerospace Conference. IEEE, 2019, pp. 1–13. range of properties are. Secondly, more in-depth propulsion [5] R. Funase, T. Inamori, S. Ikari, N. Ozaki, S. Naka- systems trades are required using different systems, operating jima, H. Koizumi, A. Tomiki, Y. Kobayashi, and points, and configurations in a low-thrust trajectory optimiza- Y. Kawakatsu, “One-year deep space flight result of tion framework to determine the highest performance system. the world’s first full-scale 50kg-class deep space probe The analysis provided here was a preliminary feasibility procyon and its future perspective,” 2016. study, however an increased fidelity mission analysis and spacecraft design would help further address uncertainties in [6] N. Miyazaki. Micro spacecraft investigates the design space. cometary water mystery. [Online]. Available: https://www.nao.ac.jp/en/news/science/2017/20170124- Both the 24 kg and 180 kg architecture could enable lower- procyon.html cost access to Small Bodies to characterize them in detail [7] A. T. Klesh, J. D. Baker, J. Bellardo, J. Castillo-Rogez, from close proximity over extended duration. These architec- J. Cutler, L. Halatek, E. G. Lightsey, N. Murphy, and tures are capable of facilitating varying degrees of autonomy C. Raymond, “Inspire: Interplanetary nanospacecraft to enable long periods of sustained activity without ground pathfinder in relevant environment,” in AIAA SPACE support. This will allow for the opportunity to build up 2013 Conference and Exposition, 2013, p. 5323. heritage in autonomy and demonstrate advancements incre- [8] B. Sherwood, S. Spangelo, A. Frick, J. Castillo-Rogez, mentally. In the longer term, with further development in A. Klesh, E. J. Wyatt, K. Reh, and J. Baker, “Planetary end-to-end autonomy, these architectures can also be used cubesats come of age,” 2015. to create a standardize platform to enable large-scale exploration of Small Bodies with a fleet of small spacecraft. It was [9] J. Schoolcraft, A. Klesh, and T. Werne, “Marco: in- found that a high degree of architectural similarities exists terplanetary mission development on a cubesat scale,” with other interplanetary small spacecraft missions proposed, in Space Operations: Contributions from the Global including main asteroid belt flybys, comet flybys, multi-NEA Community. Springer, 2017, pp. 221–231. flybys, Venus, Mars, and Martian moons missions. Therefore, [10] NASA/JPL-Caltech. Marco makes space the architecture provided here can be useful to provide a range for small explorers. [Online]. Available: of immediate and long-term benefits for science, ultimately https://www.jpl.nasa.gov/news/news.php?feature=7235 leading to an improved understanding of the evolution of the Solar System. [11] Y. Tsuda, M. Yoshikawa, M. Abe, H. Minamino, and S. Nakazawa, “System design of the hayabusa 2—asteroid sample return mission to 1999 ju3,” Acta Astronau- tica, vol. 91, pp. 356–362, 2013. ACKNOWLEDGMENTS [12] G. A. Landis, S. R. Oleson, M. L. McGuire, L. M. The authors would like to thank Alex Austin for providing Burke, M. C. Martini, J. E. Fittje, and T. W. Packard, “A his advise and feedback regarding small spacecraft concept cubesat asteroid mission: Propulsion trade-offs,” in 50th development and best practices, Damon Landau for sharing AIAA/ASME/SAE/ASEE Joint Propulsion Conference, his experience and assistance in trajectory optimization and 2014, p. 3755. 17 [13] C. Mercer, “Small satellite missions for planetary sci- from orbit,” in Space Telescopes and Instrumentation ence,” 2019. 2018: Optical, Infrared, and Millimeter Wave, vol. [14] N. Boulder. Double take: Satellites to get 10698. International Society for Optics and Photonics, an eye on asteroid pairs. [Online]. Available: 2018, p. 106985T. https://www.colorado.edu/today/2019/06/21/double- [27] J. L. Ferreira, A. A. Martins, R. A. Miranda, H. O. take-satellites-get-eye-asteroid-pairs Junior, R. Silva, I. S. Ferreira, A. Sukhanov, and O. C. [15] B. K. Malphrus, K. Z. Brown, J. Garcia, C. Con- Winter, “Development of a solar electric propulsion ner, J. Kruth, M. S. Combs, N. Fite, S. McNeil, system for the first brazilian deep space mission iepc- S. Wilczweski, K. Haught et al., “The lunar icecube em- 2017-166,” 2017. 1 mission: Prospecting the moon for water ice,” IEEE [28] A. D. Takir, J. Mulqueen, and J. Castillo-Rogez, “Aster- Aerospace and Electronic Systems Magazine, vol. 34, oid reconnaissance for researching organics and water no. 4, pp. 6–14, 2019. in the solar system (arrows2): A deep space smallsat [16] C. Hardgrove, J. DuBois, L. Haffern, E. Cisneros, mission (concept) to primitive asteroids,” in Lunar and J. Bell, T. Crain et al., “The lunar polar hydrogen Planetary Science Conference, vol. 49, 2018. mapper (lunah-map) mission,” 2019. [29] J. O’Rourke, J. Castillo-Rogez, L. Elkins-Tanton, [17] N. Bosanac, A. D. Cox, K. C. Howell, and D. C. Folta, R. Fu, T. Harrison, S. Marchi, R. Park, B. Schmidt, “Trajectory design for a cislunar cubesat leveraging D. Williams, C. Seybold et al., “Athena: The first-ever dynamical systems techniques: The lunar icecube mis- encounter of (2) pallas with a smallsat,” in Lunar and sion,” Acta Astronautica, vol. 144, pp. 283–296, 2018. Planetary Science Conference, vol. 50, 2019. [18] L. McNutt, L. Johnson, P. Kahn, J. Castillo-Rogez, and [30] C. Venigalla, N. Baresi, J. D. Aziz, B. Bercovici, D. N. A. Frick, “Near-earth asteroid (nea) scout,” in AIAA Brack, A. Dahir, S. De Smet, J. Fulton, M. M. Pel- Space 2014 Conference and Exposition, 2014, p. 4435. legrino, and S. Van wal, “Near-earth asteroid characterization and observation (neaco) mission to asteroid [19] J. Lightholder, D. R. Thompson, J. Castillo-Rogez, and (469219) 2016 ho 3,” Journal of Spacecraft and Rock- C. Basset, “Near earth asteroid scout cubesat science ets, pp. 1–16, 2019. data retrieval optimization using onboard data analysis,” in 2019 IEEE Aerospace Conference. IEEE, 2019, pp. [31] P. Bambach, J. Deller, E. Vilenius, S. Pursiainen, 1–7. M. Takala, H. M. Braun, H. Lentz, and M. Wittig, “Discus–the deep interior scanning cubesat mission to [20] R. Walker, D. Binns, C. Bramanti, M. Casasco, P. Con- a rubble pile near-earth asteroid,” Advances in Space cari, D. Izzo, D. Feili, P. Fernandez, J. G. Fernandez, Research, vol. 62, no. 12, pp. 3357–3368, 2018. P. Hager et al., “Deep-space cubesats: thinking inside the box,” Astronomy & Geophysics, vol. 59, no. 5, pp. [32] L. Walker, C. Greco, M. Di Carlo, A. Wilson, L. Riccia- 5–24, 2018. rdi, A. Berquand, and M. Vasile, “Nanospacecraft exploration of asteroids by collision and flyby reconnais- [21] R. Kruzelecky, P. Murzionak, J. Lavoie, I. Sinclair, sance,” in Low-Cost Planetary Missions Conference, G. Schinn, C. Underwood, Y. Gao, C. Bridges, 2019. R. Armellin, A. Luccafabris et al., “Vmmo lunar volatile and mineralogy mapping orbiter.” 48th Inter- [33] D. A. Vallado, Fundamentals of astrodynamics and national Conference on Environmental Systems, 2018. applications. Springer Science & Business Media, 2001, vol. 12. [22] R. Walker, D. Koschny, C. Bramanti, I. Carnelli, E. C. S. Team et al., “Miniaturised asteroid remote geophysical [34] N. Arora and R. P. Russell, “A fast and robust multiple observer (m-argo): a stand-alone deep space cubesat revolution lambert algorithm using a cosine transforma- system for low-cost science and exploration missions,” tion,” Paper AAS, vol. 13, p. 728, 2013. in 6th Interplanetary CubeSat Workshop, Cambridge, [35] A. W. Harris, “On the revision of radiometric albedos UK, vol. 30, no. 05, 2017. and diameters of asteroids,” Icarus, vol. 126, no. 2, pp. [23] J. Plescia, O. Barnouin, D. Richardson, N. Schmerr, 450–454, 1997. D. Lawrence, B. Denevi, C. Ernst, and H. Yu, “The [36] A. Mereta and D. Izzo, “Target selection for a small low- asteroid probe experiment (apex) mission,” AIAA/USU thrust mission to near-earth asteroids,” Astrodynamics, Conference on Small Satellites, 2017. vol. 2, no. 3, pp. 249–263, 2018. [24] P. Clark, M. Collier, M. Schaible, W. M. Farrell, [37] T. P. McElrath, G. Lantoine, D. Landau, D. Grebow, D. Folta, K. M. Hughes, J. W. Keller, B. Malphrus, A. S. N. Strange, R. Wilson, and J. Sims, “Using gravity Rivkin, S. Murchie et al., “Overview of phobos/deimos assists in the earth-moon system as a gateway to the regolith ion sample mission (prism) concept,” in Cube- solar system,” 2012. Sats and NanoSats for Remote Sensing II, vol. 10769. International Society for Optics and Photonics, 2018, p. [38] E. Agasid, R. Burton, R. Carlino, G. Defouw, 107690I. A. Dono Perez, A. G. Karacahoglu, B. Klamm, A. Rademacher, J. Schalkwyck, R. Shimmin et al., [25] L. Montabone, M. VanWoerkom, B. Cantor, M. Wolff, “Small spacecraft technology state of the art,” NASA, M. Capderou, F. Forget, and M. Smith, “Mars aerosol Ames Research Center, Small Spacecraft Systems Vir- tracker (mat): an areostationary cubesat to monitor dust tual Institute, NASA/TP-2018-220027, 2018. storms and water ice clouds,” in Lunar and Planetary Science Conference, vol. 49, 2018. [39] D. Lev, R. M. Myers, K. M. Lemmer, J. Kolbeck, [26] B. M. Sutin, J. Cutts, A. M. Didion, M. Drilleau, H. Koizumi, and K. Polzin, “The technological and M. Grawe, J. Helbert, A. Karp, B. Kenda, A. Komjathy, commercial expansion of electric propulsion,” Acta As- S. Krishnamoorthy et al., “Vamos: a smallsat mission tronautica, 2019. concept for remote sensing of venusian seismic activity [40] A. R. Tummala and A. Dutta, “An overview of 18 cube-satellite propulsion technologies and trends,” T. Yi, J. Lyons, M. Piszczor, J. McNatt, C. Taylor et al., Aerospace, vol. 4, no. 4, p. 58, 2017. “Solar power technologies for future planetary science [41] K. Lemmer, “Propulsion for cubesats,” Acta Astronau- missions,” Jet Propulsion Lab, California Institute of tica, vol. 134, pp. 231–243, 2017. Technology, 2017. [42] M. VanWoerkom, V. Gorokhovsky, G. Pulido, A. Sei- [56] K. B. Chin, E. J. Brandon, R. V. Bugga, M. C. Smart, dcheck, J. Williams, and C. Farnell, “Test results of S. C. Jones, F. C. Krause, W. C. West, and G. G. exoterra’s halo micro electric propulsion system for Bolotin, “Energy storage technologies for small satellite microsatellites,” in AIAA Propulsion and Energy 2019 applications,” Proceedings of the IEEE, vol. 106, no. 3, Forum, 2019, p. 4163. pp. 419–428, 2018. [43] W. Tighe, K.-R. Chien, and R. Spears, “Xips ion [57] B. C. Technologies. Our components. [Online]. thrusters for small satellite applications,” 2007. Available: https://bluecanyontech.com/components [44] P. N. Randall, R. A. Lewis, and S. D. Clark, “Qinetiq [58] S. Bhaskaran, J. Riedel, S. Synnott, and T. Wang, “The t5 based electric propulsion system and architectural deep space 1 autonomous navigation system-a post- 35th International flight analysis,” in Astrodynamics Specialist Confer- options for future applications,” in ence Electric Propulsion Conference, 2017. , 2000, p. 3935. [59] S. B. Broschart, N. Bradley, and S. Bhaskaran, “Kine- [45] R. Conversano, A. Barchowsky, R. B. Lobbia, V. H. matic approximation of position accuracy achieved us- Chaplin, A. Lopez-Ortega, J. A. Loveland, A. D. Lui, ing optical observations of distant asteroids,” Journal of G. Becatti, S. W. Reilly, D. M. Goebel, J. S. Snyder Spacecraft and Rockets, pp. 1–10, 2019. et al., “Overview of the ascendant sub-kw transcelestial electric propulsion system (astraeus),” in 36th Interna- [60] R. Park and J. Riedel, “Mini-advanced pointing imag- tional Electric Propulsion Conference, 2019. ing camera (mapic) concept,” in Lunar and Planetary Science Conference, vol. 49, 2018. [46] O. Duchemin, F. Marchandise, N. Cornu, and E. Bourguignon, “Electric propulsion thruster assem- [61] W. J. Larson and J. R. Wertz, “Space mission analysis bly for future small geostationary comsats,” in 44th and design,” Torrance, CA (United States); Microcosm, AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Inc., Tech. Rep., 1992. Exhibit, 2008, p. 5182. [62] Standard: Mass Properties Control for Space Systems [47] Y. Rahmat-Samii, V. Manohar, J. M. Kovitz, R. E. (ANSI/AIAA S-120A-2015). [Online]. Available: Hodges, G. Freebury, and E. Peral, “Development of https://arc.aiaa.org/doi/abs/10.2514/4.103858.001 highly constrained 1m ka-band mesh deployable offset [63] R. P. Welle, “Propellant storage considerations for elec- reflector antenna for next generation cubesat radars,” tric propulsion,” in 22nd International Electric Propul- IEEE Transactions on Antennas and Propagation, 2019. sion Conference, 1991, pp. 91–107. [48] R. E. Hodges, N. Chahat, D. J. Hoppe, and J. D. Vacchione, “A deployable high-gain antenna bound for mars: Developing a new folded-panel reflectarray for BIOGRAPHY[ the first cubesat mission to mars.” IEEE Antennas and Propagation Magazine, vol. 59, no. 2, pp. 39–49, 2017. [49] S. Gao, Y. Rahmat-Samii, R. E. Hodges, and X.-X. Sandro Papais is a visiting researcher Yang, “Advanced antennas for small satellites,” Pro- at the Jet Propulsion Laboratory. He ceedings of the IEEE, vol. 106, no. 3, pp. 391–403, received his B.Eng degree in Mechani- 2018. cal Engineering from McGill University in 2019. He previously worked as a [50] M. M. Kobayashi, S. Holmes, A. Yarlagadda, research assistant at McGill University F. Aguirre, M. Chase, K. Angkasa, B. Burgett, L. Mc- where his work focused on trajectory Nally, T. Dobreva, and E. Satorius, “The iris deep-space modeling and optimal control. Sandro transponder for the sls em-1 secondary payloads,” IEEE has led a student sounding rocket group Aerospace and Electronic Systems Magazine, vol. 34, for two years where they have designed, no. 9, pp. 34–44, 2019. manufacture, and launched 6 sounding rockets up to 30,000 [51] M. B. O’Neill, W. P. Millard, B. M. Bubnash, R. H. feet, static tested a student made hybrid rocket engine, and Mitch, and J. A. Boye, “Frontier radio lite: A single- placed first overall at the Spaceport America Cup in 2018 board software-defined radio for demanding small satel- against 124 international universities. His interests include lite missions,” 2016. design, estimation, and control of aerospace and robotic systems. [52] M. Pugh, I. Kuperman, F. Aguirre, H. Mojaradi, C. Spurgers, M. Kobayashi, E. Satorius, and T. Jedrey, “The universal space transponder: A next generation software defined radio,” in 2017 IEEE Aerospace Con- ference. IEEE, 2017, pp. 1–14. [53] M. Design. Hawk specifications. [Online]. Available: https://mmadesignllc.com/specs-table/ [54] D. Turse, L. Adams, K. A. Medina, K. Steele, and T. Stern, “Design, analysis and testing of a composite beam roll-out array (cobra) for small satellites,” in AIAA Scitech 2019 Forum, 2019, p. 2024. [55] R. Surampudi, J. Blosiu, P. Stella, J. Elliott, J. Castillo, 19 Ben Hockman is a Robotics Technol- Group in the Mission Design and Navigation Section. As ogist at JPL in the Robotic Mobility manager of this group, Dr. Bhaskaran is responsible for Group. He received is PhD from Stan- the technical and administrative oversight of 15 engineers ford University in 2018, where his grad- responsible for navigation analysis and operations of several uate work focused on robotic surface current and future missions, including Cassini, Dawn, Juno, mobility on small Solar System bodies. and Europa Clipper. Dr. Bhaskaran’s honors include two Ben’s research interests include design, NASA Exceptional Achievement Medals, awards for Techni- control, modeling, estimation, and deci- cal Excellence, and a NASA Space Act Award for his work sion making for space robotic systems. on the Deep Space 1 Autonomous Navigation System. He has Ben has worked on extreme-terrain teth- also received over 20 NASA Group Achievement Awards for ered rovers for exploring Lunar pits, internally-actuated hop- his work on the various missions listed above. Dr. Bhaskaran ping robots for asteroids and comets, melt probes for access- received a B.S (1985). and M.S. (1987) from the University ing the oceans of icy moons, and algorithms for spacecraft of Texas at Austin, and a Ph.D (1991) from the University of and rover autonomy. Colorado at Boulder, all in Aerospace Engineering.

Saptarshi Bandyopadhyay is currently Issa Nesnas is a principal technologist a Robotics Technologist at the Jet and the supervisor of the Robotic Mobil- Propulsion Laboratory, California In- ity group at the Jet Propulsion Labora- stitute of Technology. He received his tory with over two decades of research Ph.D. in Aerospace Engineering in Jan- in space robotics and industrial automa- uary 2016 from the University of Illinois tion. He leads research in autonomy at Urbana-Champaign (UIUC). Sap- and mobility with a focus on extreme tarshi received his Dual Degree (B.Tech terrain access and microgravity mobil- and M.Tech) in Aerospace Engineering ity. He contributed to the development of in 2010 from the Indian Institute of Tech- autonomous rover navigation and visual nology Bombay, India. At IIT Bombay, he co-founded and led target tracking and participated in the development of the the student satellite project. IIT Bombay’s Pratham satellite Curiosity and Mars 2020 rovers. Dr. Nesnas served on was launched into low Earth orbit in September 2016. Sap- NASA’s Capability Leadership Team for Autonomy and was tarshi won the gold medal for India at the 9th International the co-chair for NASA’s Technology Roadmaps for Robotics Astronomy Olympiad held in Ukraine in 2004. Saptarshi’s and Autonomous Systems. He holds several patents and has research interests include aerospace systems, robotics, multi- over fifty publications in this field. He holds a B.E. degree in agent systems and swarms, dynamics and controls, estimation Electrical Engineering from Manhattan College and a M.S. theory, probability theory, and systems engineering. and Ph.D. in Mechanical Engineering with a specialization in robotics from the University of Notre Dame. Reza Karimi joined Jet Propulsion Lab- oratory in June 2014. He’s currently a mission designer and orbit determination analyst. He has worked on a variety of projects including Rosetta, Dawn, ISRO MOM, New Frontiers and Discov- ery proposals, few Strategic and Spon- taneous Research & Technology Devel- opment tasks, Europa Clipper, Europa Lander, CubeSat formation flying design, tens of Team-X studies, and numerous short-term mission design and navigation tasks. Reza received his Ph.D. degree in Aerospace Engineering from Texas A&M University in May 2012 followed by a post Doc position in Aerospace Engineering, Texas A&M University, until May 2014.

Shyam Bhaskaran began his profes- sional career at the Jet Propulsion Lab- oratory in January of 1992 as an orbit determination specialist on the Galileo mission. Since then, he has served as a member and lead of the navigation team for many missions. He is the one of the principal architects of the autonomous navigation system used on the Deep Space 1, STARDUST, and Deep Impact missions. He then moved into management, serving as the Mission Management Ofﬁce Team Chief for Guidance, Navigation and Control, during which time he oversaw the successful navigation efforts of STARDUST’s ﬂyby of the comet Wild 2, the Genesis Earth return, and Deep Impact’s encounter with comet Tempel 1. Since 2005, Dr. Bhaskaran has been the Supervisor for the Outer Planet Navigation