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Deep-space CubeSats: thinking inside the box

Roger Walker and colleagues CubeSats: small but perfectly formed consider the potential for sending Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 nanospacecraft into deep space. CubeSats are modular , using multiple standard-sized units of 10 x 10 x 10 cm ince the introduction of the CubeSat (figure 1); the size of the resulting spacecraft standard in the early 2000s, there has is measured by the number of units, e.g. 3U Sbeen a proliferation of nano-/small or 6U. The capabilities of these have microsatellites in low , with increased significantly in recent years, notably 100–300 or more launched annually and at in pointing, propulsion and communications a growing rate (according to reports from as well as the availability of compact optical/ SpaceWorks and Euroconsult). CubeSats radio frequency/environment payloads. have reduced entry-level costs for space CubeSats are launched inside a standard missions in (LEO) by container which simplifies launcher more than an order of magnitude (see box accommodation while ensuring safety for “CubeSats: small but perfectly formed”). the primary passenger, thus facilitating 1 Example of a 1U CubeSat shell. There is a This has brought new players into the widespread low-cost piggyback launch weight standard too – the finished launch unit space sector and launch opportunities for opportunities on many different launch must not exceed 1.33 kg. (ESA/G ) CubeSats have increased significantly in vehicles worldwide. At the same time, the the last decade to address the associated extensive use of miniaturized electronics, entry-level costs mean that the true power of demand. With more piggy­back oppor- sensors/actuators and a stacked integration CubeSat-based nano- or small microsatellites tunities, platform capabilities and small approach has cut spacecraft production and lies in their operation as distributed systems, payloads have rapidly advanced to a level integration costs. The dramatically lower such as constellations or swarm formations. that are suitable for real operational mis- sions. Niche commercial applications are emerging based on operating multiple within the frame of the ESA General Stud- escape or Earth escape . For the CubeSats together in distributed systems, ies Programme on lunar and interplanetary time being, piggyback flight opportunities e.g. constellations and swarms. CubeSat mission concepts. must be found, either on a launcher upper How far can this new paradigm be These studies have involved either stage carrying a primary spacecraft, or on extended from the safety of LEO out to mother–daughter system architectures, the primary spacecraft itself as part of its cis-lunar and deep space, and what unique where the CubeSats are carried to a target mission. However, these opportunities are new missions can be performed? Piggy- destination such as lunar rare, and the constraints on back opportunities for CubeSats to or to a near-Earth object “The first inter­planetary nanospacecraft deployment orbit and interplanetary space are already (NEO) on a larger spacecraft nanospacecraft can significantly influence becoming available, and innovative minia- and deployed at the target in mission was launched the Dv (velocity impulse) and turized technologies are being developed order to fulfil their mission, in May, to ” transfer window needed to overcome the severe technical challenges or a completely stand-alone to reach the final mission des- of deep-space missions. Performance system executing its own mission. The tination – and hence the feasibility of such a has reached a level where the first inter­ mother–daughter architecture alleviates mission. Table 1 shows our assessment of a planetary nanospacecraft mission (NASA’s the technical challenges of propulsion, range of missions that could be performed – MarCO) was launched long-range communication and deep-space beyond Earth orbit for a given set of piggy­ in May 2018 as part of the InSight mission environment survivability, because the host back flight opportunities. (Klesh & Krajewski 2018). And, as was the spacecraft provides resources and accom- case for LEO, it is expected that there will modation during the cruise, as well as Impact Mission be an order of magnitude reduction in the communications to Earth ground stations, Within the ESA General Studies Pro- entry-level cost of interplanetary mis- in conjunction with local inter- gramme, the third edition of the Sysnova sions, thus paving the way to new mission links with the deployed CubeSats. For a technical challenge focused on Cubesat applications and architectures based on stand-alone deep-space CubeSat (i.e. where Payload of Intersatellite distributed systems of deep-space nano- there is no mothercraft), these challenges Networking Sensors (COPINS). The overall spacecraft. This article addresses some of have to be tackled and suitable technology objective of the challenge was to investigate the missions that may be performed with solutions identified and/or developed. new mission concepts involving a number such distributed systems, their system There are presently no dedicated small of CubeSats operating together in inter­ architectures and enabling technologies, launchers that can cost-effectively inject planetary space in support of the objec- based on a number of studies completely nanospacecraft onto specific near-Earth tives of the proposed ESA Asteroid Impact

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1 Mission scenarios with respect to different piggyback flight opportunities

piggyback launcher primary CubeSat CubeSat final propulsion feasibility (≤12U) on spacecraft deployment destination type spacecraft PSLV/ Lunar Pathfinder highly eccentric low lunar orbit/ chemical mono- yes lunar orbit Earth– L2 propellant spacecraft Soyuz/Ariane 6 HERA Didymos binary NEO Didymos binary NEO cold gas yes launcher JUICE -bound Venus high-altitude chemical mono- yes (4 × 3U CubeSat atmosphere swing-by propellant entry probes with microcarrier) launcher Soyuz/Ariane 6/ various telecom geostationary lunar orbit/ electric no (excessive Δv, time

Falcon 9/ transfer orbit NEO rendezvous and radiation dose) Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 launcher Ariane 6 various transfer to NEO rendezvous/ electric yes (use of L2 astronomy –Earth L2 Sun–Earth L5/ for waiting until transfer heliocentric (<1 au) to target) launcher Space Launch System lunar transfer lunar orbit/ electric yes trajectory NEO rendezvous launcher Orion lunar transfer lunar orbit/ chemical no (excessive Δv) trajectory NEO rendezvous launcher Proton/Atlas V ExoMars/ Mars transfer Mars orbit electric yes (but excessive trajectory duration for spiral down) launcher Proton/Atlas V ExoMars/ Mars transfer Mars orbit chemical no (excessive Δv) Mars 2020 trajectory

2 Illustration of the AIM spacecraft and deployed CubeSats at asteroid Didymos during the DART hypervelocity impact. (ESA)

Mission (AIM). AIM had three different the smaller component of the Didymos the spacecraft would then retreat to objectives relating to technology flight binary asteroid (nicknamed Didymoon) at a distance of 100 km, ready for the impact demonstration in interplanetary space, high velocity; AIM was envisaged as a up to a month later. The CubeSats would investigation of NEO mitigation techniques means to rendezvous with the target aster- perform their mission up until two months and acquiring new scientific knowledge for oid in advance, in April 2022, and charac- after the impact date, i.e. for three months understanding system evolution. The terize the binary system before, during and in total, using a 1 Mbps S-band inter- proposed COPINS payload consisted of two after the . The launch of the satellite (ISL) with the main space- or more CubeSats carried on the main AIM AIM spacecraft was proposed for 2020 on craft’s communications system (and each spacecraft in their deployment systems (two the Soyuz/Fregat launcher from Kourou other) for telecommand, housekeeping and 3U deployers), and released in the vicinity into a direct escape trajectory, with arrival payload telemetry transfer. Five different of the target asteroid Didymos, where the at the asteroid after about 22 months. Dur- CubeSat concepts were selected by ESA main AIM spacecraft acted as a commu- ing the rendezvous phase, the distance to for parallel studies via an open competi- nications data relay between the CubeSats the Sun and Earth is close to 1 au and from tive announcement of opportunity (AO) and Earth ground stations (figure 2). 0.5 to 0.1 au respectively. process, as described in table 2. The ESA AIM mission (ESA AIM Team After an initial measurement phase, the The AIM mission did not receive the 2015) was proposed as part of the AIDA AIM spacecraft would be manoeuvred to required funding for implementation, but international cooperation, together with a within 10 km of the binary asteroid. Before the objectives and concept remain valid. US spacecraft, Double Asteroid Redirection DART’s impact, the COPINS CubeSats Therefore a new mission proposal called Test (DART). DART is planned to impact (Walker et al. 2016) would be released and HERA has since been formulated based

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2 Overview of the five AIM COPINS CubeSat mission concepts

name organization system mission payload ASPECT: VTT, single 3U CubeSat with spectral imaging of Didymoon before/ imaging spectrometer in VIS/ Asteroid Spectral Uni. Helsinki, <1° pointing, cold gas after impact from 4 km orbit NIR: 1–2 m GSD; Imaging Aalto Uni. propulsion 1 m s–1 Δv SWIR spectrometer DustCube Uni. Vigo, single 3U CubeSat with characterize ejected dust plume after in situ nephelometer, Uni. Bologna, cold gas propulsion impact from 3–5 km orbit, transfer to remote nephelometer MICOS 2 m s–1 Δv, optical IR rel. L4/L5 orbit pre-impact, DRO 280 m alt. navigation post-impact

CUBATA GMV, two 3U CubeSats with gravity field determination of Didymos radio science: Cube-Cube LoS

Uni. La Sapienza, cold gas propulsion Δv system before and after impact Doppler tracking with S-band Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 INTA 1.5 m s–1, <1° pointing, transponder and ultrastable optical rel. navigation oscillator PALS: Swedish Institute of two 3U CubeSats with magnetization, bulk chemical narrow angle camera, volatile Payload of Advanced Space Physics, cold gas propulsion composition, presence of volatiles, super- composition analyser, fluxgate Little Satellites KTH, DLR, IEEC, 12.5 m s–1 Δv, <1° resolution surface imaging of Didymos magnetometer, video emission AAC Microtec pointing, optical rel. components impact ejecta via tour of spectrometer navigation Didymos system AGEX: ROB, two 3U CubeSats: one determination of dynamical state, lander: three-axis seismometer, Asteroid Geophysical ISAE Supaero, lander, one orbiter geophysical surface properties, accelerometers, three-axis Explorer Emxys, subsurface structure of Didymoon gravimeter; orbiter: 30 chipsats Antwerp Space before/after impact deployed to surface

3 Illustration of CubeSats for Exploration (LUCE)”. The over- the Surrey Satellite all rationale for the challenge is to prepare Technology Ltd/ European teams for flight opportunities to Goonhilly Earth Station the Moon that may arise, while addressing Lunar Pathfinder and supporting ESA’s lunar exploration and deployed lunar objectives. The activity was run through CubeSats. (SSTL) an open competitive AO process. Propos- als responding to the AO were able to address one or more themes in their mission concepts and were required to comply with a common set of mission and system con- straints. Single spacecraft as well as multiple distributed spacecraft were permitted, pro- vided they complied with the constraints. It was assumed that a larger lunar orbiter spacecraft would provide transportation for up to 60 kg of CubeSat payload (a maximum of 12U per spacecraft) to a >500 km circular, >50° inclination lunar orbit, and data relay services to/from Earth ground stations via a predefined ISL operating at UHF frequen- cies with Proximity-1 protocol, providing maximum data rate of 512 kbps with one hour per day link availability over one year. The SSTL/GES Lunar Pathfinder (Saunders et al. 2016) mission (see figure 3), proposed to be undertaken in partnership with ESA, is aimed at providing these capabilities for on a consolidation of the extensive techni- and time synchronization, a low-velocity lunar CubeSats as a commercial service, and cal work performed for AIM, focused on (<5 cm s–1 ± 20%) CubeSat deployer with is planned for launch in the 2023 timeframe. the asteroid deflection objectives. This has power/data interfaces, and optical relative Hence, the associated propulsion and com- led to a change in the CubeSat payload to navigation techniques such as centre of munications challenges were alleviated for a single 6U CubeSat with a payload to be brightness determination. the deployed CubeSats. refined during 2018. During the COPINS More than 30 innovative CubeSat mis- studies, a number of key technologies LUnar CubeSats for Exploration sion concept proposals were submitted have been identified that would require Within the ESA General Studies Pro- to the AO, covering a diverse range of development, including short-range S-band gramme, the fourth edition of the Sysnova operations concepts, platform designs and ISLs with ranging (1 Mbps, 1 m accuracy) technical challenge focused on “LUnar instruments, demonstrating a significant

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3 Overview of LUCE mission concept studies

name organization system mission payload LUMIO: Politecnico Di Milano, single 12U CubeSat with impact flash optical camera with Lunar Meteoroid TU Delft, EPFL, S[&]T chemical propulsion for L2 monitoring of the far side from 3.4 km/pixel spatial resolution, Impact Observer Norway, Leonardo S.p.A, transfer, lunar disc optical an Earth–Moon L2 halo orbit 3 Hz frame rate, FOV 3.5° Uni. Arizona navigation MoonCARE: Von Karman Institute, three 12U CubeSats with lunar radiation environment radiation detector Moon CubeSat DLR, chemical propulsion for orbit characterization, study (0.06–200 MeV), astrobiology for the Analysis Tyvak International, transfer radiation effects on specific experiment (VIS/UV of the Radiation Politecnico di Torino microorganisms spectrometer, four Environment microorganisms)

CLE: ISIS bv, three 12U CubeSats with inter­ technology demonstrator for <30 MHz software-defined Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 CubeSat Low , satellite link for ranging/time distributed low-frequency radio radio receiver, three frequency Explorer Radboud Uni. Nijmegen, synchronization/data transfer interferometric telescopes in deployable monopole Uni. Twente, for interferometry, propulsion lunar orbit, operated in radio antennae per spacecraft TU Delft for loose formation flying quiet zone around far side VMMO: MPB Communications, single 12U CubeSat with electric content of water ice deposits in active fibre laser at 1560 Lunar Volatile Uni. Surrey, propulsion for transfer to very permanently shadowed craters and 530 nm with 10 m spot and Mineralogy Lens R&D, low lunar orbit at pole, other volatiles size, VIS/NIR spectrometer, Mapping Orbiter Uni. Winnipeg (e.g. ilmenite) on the day side, radiation environment sensor lunar radiation environment

4 Deep-space CubeSat configuration and key technologies. (ESA)

interest in lunar CubeSats. As a result, light rejection. Additionally, a near-infrared navigation (e.g. full lunar disc at high four mission concepts (Walker et al. 2017a (NIR) channel was added to the LUMIO altitude, or feature tracking at low altitude), and summarized in table 3) were awarded payload for improved detection of impact X-band transponder with ranging directly parallel six-month study contracts. flashes and determination of impact energy. from/to Earth ground stations, or GNSS Based on an evaluation of the parallel The CDF studies identified several key receiver with medium gain antenna. study results, an ESA panel selected the technologies for lunar CubeSats operating LUMIO and VMMO mission in mother–daughter system M-ARGO concepts as joint winners of “Four mission concepts architectures. These include For deep-space missions, assuming a the Sysnova challenge. The were awarded parallel chemical propulsion with a piggy­back launch opportunity to near- winners received a prize six-month study Dv capability of >200 m s–1 in Earth escape as a starting point, truly stand- consisting of a further study contracts” order to either transfer from alone nanospacecraft – those with no larger on their concept in the ESA the injection orbit provided mothercraft – have to be independently Concurrent Design Facility (CDF) sup- by, e.g. Lunar Pathfinder, to final opera- capable of reaching their target destination ported by an ESA engineering team. tion orbit (LUMIO) or maintain the orbit using on-board propulsion with several The main refinements made were on the to ensure perilune at low altitude over the km s–1 of Dv, and communicating directly mission design to ensure compatibility with south pole region (VMMO). Additionally, back to Earth over distances of up to 1–2 au. feasible injection from Lunar Pathfinder, long-range (4000–8000 km) UHF or S-band A deep-space CubeSat system called the resulting propulsion subsystem selec- ISLs with the mothercraft are required for M-ARGO (Miniaturised Asteroid Remote tion (chemical propulsion >200 m s–1 for communication back to Earth. For naviga- Geophysical Observer; Walker et al. 2017b) both), increase in power, and increase of the tion, there are different options depending has been designed, using the requirements baffle length for the optical payload (both on the accuracy requirements including: for a near-Earth object (NEO) rendezvous LUMIO and VMMO) for improved stray ISL with ranging/Doppler, optical relative reference mission as a design driver.

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A bottom-up approach was used to investigate how much propulsive Dv could 4 Overview of M-ARGO deep-space CubeSat design be achieved within a 12U CubeSat, while still accommodating a science payload of attribute specification 1–2U and downlinking the science data volume 12U form factor (226 x 226 x 340 mm) back to Earth at a reasonable data rate when propulsion mini-RIT gridded ion engine, gimbal, PPU, neutralizer, two Xenon the spacecraft is operating in close proxim- propellant tanks and feed system (max. 2.8 kg) ity to the NEO. Figure 4 shows the space- communications X-band deep space transponder with ranging/Doppler (two receive, craft design and key enabling technologies, three transmit channels), four patch antennas for omni-directional and the design specifications are provided TT&C, deployable high gain antenna reflect-array for payload data in table 4. Achieving very high Dv requires electric propulsion. Because the duration of power single body-mounted panel (6U face) + Li-ion battery the interplanetary transfer phase using an two-wing deployable solar array with drive mechanism electric propulsion system is proportional power to EP at 1 au 93 W (six panels) 120 W (eight panels) Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 to the spacecraft thrust-to-mass ratio, a thrust at 1 au 1.7 mN 2.4 mN great deal of effort was made to maximize I uat 1 a 3050 s 3180 s the thrust within the mass constraints sp Δv (low thrust) ~3750 km s–1 while achieving a high specific impulse I( sp) to minimize propellant storage mass. This wet mass (w/margins) 21.6 kg 22.3 kg involved: maximizing power generation; altitude and orbit sensors: visnav camera, tracker, six Sun sensors, IMU selecting an available propulsion technol- control system actuators: three reaction wheels, eight Xe cold gas RCS thrusters ogy with a moderate power-to-thrust ratio data handling modular avionics with payload data processing for high Isp; and minimizing electrical losses from sunbeam to ion beam. This thermal control passive with radiator panels and heaters trade-off and optimization process led to Key: RIT = radio frequency ion thruster. PPU = power processing unit. TT&C = telemetry, tracking and the engine model shown in figure 5 for command. EP = electric propulsion. IMU = inertial measurement unit. RCS = reaction control system. two different solar array configurations (six-panel or eight-panel array with extra body-mounted panel) for the selected min- iaturized gridded ion engine. This engine model was used as a key input to the low-thrust trajectory optimiza- tion during the analysis of specific mission scenarios. For direct-to-Earth communica- tions, the downlink performance in figure 6 was investigated for different antenna diameters of the ESTRACK deep-space ground station network (15 and 35 m), as well as the Sardinia Radio Telescope (64 m), and for different RF transmit power levels (5 and 15 W) as a function of spacecraft– 5 Sunbeam to ion beam optimization for the M-ARGO CubeSat using gridded ion engine. Earth distance. With 15 W radio frequency power, it is possible to achieve rates of Toutatis and (162173) Ryugu, all up to a The M-ARGO spacecraft design would 25 kbps and 7 kbps for the 64 m and 35 m kilometre in diameter. Almost no objects give a unique opportunity to rendezvous antennas respectively at 1 au from Earth. larger than about 100 m in diameter rotate with one of these small . A fleet Initial thermal analysis suggests that the faster than a period of a few hours. It is of M-ARGO nanospacecraft deployed heat dissipation from the electric propul- assumed that this is because the centrifu- from the same piggyback launch oppor- sion could be marginally sustained with a gal force arising from the rotation is larger tunity (each targeting a different NEO), passive thermal control using a large area than the tensile strength of the object. Some could achieve a wide survey of near- of the spacecraft body as radiator panels. models suggest that the larger objects are Earth asteroids very cost-effectively. This Apart from the electric propulsion and typically built up from small “monolithic” would reveal any fundamental differ- X-band communications system, other key objects, which can have larger spin rates. It ences between these small objects and enabling technologies identified during the can be expected that these smaller objects the previously visited larger asteroids, study include a cold gas reaction control do not have regolith on their surface and as well as potentially the compositional system for detumbling/wheel offloading, are more homogenous than larger aster- diversity among different types of asteroid. and a solar array drive mechanism to - oids. Just confirming the absence of rego- Additionally, with commercial interest mize power generation. With the system lith on a smaller asteroid would contribute in space resource exploitation emerging, design presented in this section, a number to refining formation models of asteroids. conducting a wide survey of the closest of different mission scenarios for stand- With a NIR spectral imager one could asteroids to Earth (in Dv terms) would be alone distributed nanospacecraft systems check the mineralogical composition of the the first exploratory step to identifying were investigated. object and determine its level of homoge- resources in situ for later extraction. Apart neity. Tracking the spacecraft with high from minerals, precious metals would be of The near-Earth object population accuracy using radio science during very high interest, and therefore the addition of While quite a number of asteroids have close fly-bys, the mission can constrain the a magnetometer to measure magnetic field been visited by spacecraft, so far only four internal structure of the object, again show- of the asteroid would be relevant. are NEO: (433) Eros, (25143) Itokawa, (4179) ing whether it is monolithic or not. A payload assessment was carried out

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5 Summary of payload instrument options for NEO rendezvous missions

name details mass volume maturity (kg) (U) VIS/IR spectral imager VIS (500–900 nm), NIR (900–1600 nm), SWIR (1600–2500 nm) 1 1 Aalto-1 (2017), PICASSO (2019), ASPECT RSE X-band transponder with Doppler tracking 0 0 part of comms subsystem development laser altimeter 1.5 km range (10% albedo), 0.5–1 mm accuracy 0.033 0.04 DLEM 20 (Jenoptik) to be space qualified magnetometer <2 nT sensitivity, deployed on 1 m boom 0.2 0.2 MAGIC on CINEMA (2012), RadCube (2019) low-frequency radar 20 MHz with 7.5 m dipole antenna for interior studies 1 1 DISCUS (MPI) to be developed thermal IR imager 10–14 µm (TIR) 1.5 1.5 some adaptation Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 thermal IR imager 2–25 µm (IR) 1 1.5 MIMA development for ExoMars

6 Communication downlink data rates as a function of radio frequency 7 Results from the NEO target screening process (visual magnitude vs power and Earth station antenna. propellant mass). to better understand the availability and and transfer duration less than three years. operations centre tasks performed in nor- capabilities of existing or projected minia- The complete MPCORB database of more mal working hours. The two-week pattern turized payloads that can fulfil the science than 700 000 objects was taken as a start- is enabled by flying an M-ARGO CubeSat objectives of NEO physical characteriza- ing point, and a three-impulse chemical in a square trajectory relative to the NEO tion. The results are presented in table 5. approximation used as a pre-filter. This target on its sunward side; small manoeu- The spacecraft design can only accom- resulted in 143 objects (Dv < 3.9 km s–1, more vres with the electric propulsion system modate up to 2 kg and 1.5U than 40 observations, magni- are executed offline at the corners. Between of payload, so the baseline “Several potential piggy­ tude H < 26). Ephemerides of manoeuvres, the spacecraft flies on pas- design includes all options back launch opportuni- these objects were input into sively safe hyperbolic arcs. Close approach with exception of the low- ties in the 2020s have a global low-thrust trajectory points for multispectral imager and laser frequency radar and thermal been identified” optimization tool in order to altimeter science observations are centred infrared instruments. If these calculate rendezvous trajec- mid-way through the coast arcs, with can be further miniaturized, or later on the tories for each target, optimizing for mini- navigation and radio science performed in payload resources become available, then it mum propellant mass within the launch between manoeuvres and close approach would be possible to accommodate them. window and transfer duration constraints. points. The navigation scheme includes It is assumed that the M-ARGO fleet This step resulted in 83 different targets ground-based navigation (as for ) is released after the main payload on a using a propellant mass of less than 2.5 kg used as the basis for manoeuvre com- Sun–Earth L2 transfer trajectory and, after (the maximum capacity for xenon propel- mand generation together with on-board reaching the L2 region, they are inserted lant). The distribution of these accessible optical navigation (based on centroiding into a L2 halo orbit where they wait for NEO targets over visual magnitude (hence image processing algorithm and unscented their optimum low-thrust transfer win- size) and propellant mass is given in figure Kalman filter) used to improve pointing dow to their specific target. A number of 7. Most targets are within visual magni- and for collision risk assessment. potential piggyback launch opportunities tude range 22–26 (i.e. 15–250 m diameter The distance to Earth during the science in the 2020s timeframe have been identified depending on albedo). phase and the communication link perfor- related to launch of medium/large-class Low-thrust trajectories for a few targets mance shown in figure 6 has been used to astronomy missions to L2. were then further refined using a low- determine the total payload data volume In order to identify the possible NEOs thrust trajectory local optimization tool that can be transmitted to Earth as a func- accessible by the M-ARGO fleet, a complete with numerical integration. A six-month tion of ground station utilization time per NEO target screening process (Mereta & phase for science is assumed and, in order week. This is shown in figure 8 for object Izzo 2018) was performed within the mis- to minimize the science operations cost, 2012 UV136 (0.71 au from Earth), as well as sion/system design constraints presented the close proximity operations at the the other studied targets (2013 BS45 at 1 au, above, considering a launch in the 2020–23 NEO target have been designed to follow 2016 FU12 at 0.75 au, and 2011 MD at 0.39 au period, starting with an L2 a two-week repeat pattern with mission from Earth).

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of the spacecraft then performs a transfer 6 Payload instrument options for space weather from L2 to L5, while the others perform a heliocentric missions spiral-in transfer to a circular with a semi-major axis of <1 au, one name details mass (kg) volume (U) maturity after the other in order to achieve an equal radiation electron energy 0.3–8 MeV, 0.8 0.8 RadTel on RadCube distribution of spacecraft around the orbit. telescope proton energy 3.1–500 MeV, (2019) The transfer from Sun–Earth L2 to L5 has heavy ions 50–1000 MeV/n been optimized in a low-thrust trajectory magnetometer <2 nT sensitivity, 0.2 0.2 MAGIC on CINEMA local optimization tool with numerical deployed on 1 m boom (2012), RadCube (2019) integration. The transfer takes around 20 months and consumes 1.7 kg of Xe with 342 solar X-ray flux energy range 1.5–25 keV, 0.5 1 XFM-CS under days of total thruster firing D( v of 2.6 km s–1). monitor energy channels 10+, development, demo For the other spacecraft in the fleet, each flux accuracy 1% in 2020 with 2.5 kg of Xe propellant, the lowest Downloaded from https://academic.oup.com/astrogeo/article-abstract/59/5/5.24/5099069 by guest on 22 September 2018 semi-major axis that could be achieved for a circular heliocentric orbit (with no plane change) is 0.8 au ( 0.72 years) for departure from Sun–Earth L2, and 0.75 au (orbital period 0.66 years) for depar- ture from Sun–Earth L1. The transfer takes around 440 days and 395 days respectively. For a fleet of 10 M-ARGO-like spacecraft equally spaced around the final orbit, each spacecraft would need to depart from L2 around 94 days apart, or from L1 approxi- mately 70 days apart, based on the synodic period with respect to Earth.

Conclusions Several opportunities exist to embark CubeSats on larger spacecraft to lunar orbit and to a specific NEO, where they can be deployed locally for deep investigation in a mother–daughter system architecture. 8 Science data return for close proximity operations at 2012 UV136 and other objects. Numerous feasible mission concepts have been studied for this approach. A stand- Space weather measurements of solar longitude, potentially leading to alone deep-space 12U CubeSat nanospace- Space weather and its effects on Earth- new knowledge on solar–terrestrial physics craft has been designed to be capable of orbiting satellites and terrestrial infra­ and even more advance warning of storms. rendezvous and characterization of NEOs structure is driven by particle radiation A payload assessment was carried out or space weather measurements at the from the Sun, emanating from solar flare in order to identify existing or projected Sun–Earth L5 (or other events. Continuous in situ monitoring of the miniaturized payloads that can fulfil the locations inside Earth orbit). Based on energetic particle flux, inter- objectives of space weather exploiting piggy­back launch opportunities planetary magnetic field, and “Such a spacecraft may monitoring. The results are to near-Earth escape, such a spacecraft is other space weather param- reduce the cost of inter- presented in table 6. All of expected to reduce the entry-level cost of eters such as the solar X-ray planetary missions by the payloads can be accom- deep space missions by an order of mag- flux emission at locations an order of magnitude” modated in the M-ARGO nitude, thus enabling stand-alone distrib- other than the Earth, would platform design, replacing uted systems in deep space with a fleet of dramatically increase our understanding of the NEO payload suite. nanospacecraft. New missions, such as a these processes, even with only rather lim- As with the NEO rendezvous mis- cost-effective wide survey of the NEO pop- ited measurement capabilities. A location sion scenario, it is assumed that a fleet of ulation for science/resource identification at Sun–Earth L5 would provide a few hours nanospacecraft like M-ARGO, equipped and simultaneous multi-point in situ space warning of major storm events. Ideally, with the space weather payload suite, weather measurements, would be feasible a constellation of spacecraft in a circular are injected onto a Sun–Earth L2 trans- for the first time, allowing hitchhiking heliocentric orbit inside that of the Earth fer trajectory from a piggyback launch nanospacecraft to write a new chapter in would allow us to gain a complete simulta- opportunity. From there, the spacecraft are the guide to the . ● neous picture of solar activity as a function manoeuvred to an L2 parking orbit. One

AUTHORS REFERENCES Saunders C et al. 2016 Solving communications Exploration (LUCE) mission concept studies, 6th Roger Walker, David Binns, Cristina Bramanti, ESA’s AIM Team 2015 Aiming for asteroids ESA and navigation requirements for small lunar Interplanetary CubeSat Workshop Massimo Casasco, Paolo Concari, Dario Bulletin no. 164 missions, 5th Interplanetary CubeSat Workshop Walker R et al. 2017b Miniaturised Asteroid Izzo, Davar Feili, Pablo Fernandez, Jesus Gil Klesh A & Krajewski J 2018 MarCO: Mars Cube SpaceWorks 2018 Nano/Microsatellite Market Remote Geophysical Observer (M-ARGO): a Fernandez, Philipp Hager, Detlef Koschny, One – lessons learned from readying the first Forecast 8th edition stand-alone deep space CubeSat system for Vasco Pesquita, Neil , European Space interplanetary CubeSats for flight, 49th LPSC, Walker R et al. 2016 CubeSat Opportunity Payload low-cost science and exploration missions, 6th Agency, ESTEC, Noordwijk, The . Ian paper 2923 Inter-satellite Network Sensors (COPINS) on the ESA Interplanetary CubeSat Workshop Carnelli, ESA HQ, Paris, France. Michael Khan, Mereta A & Izzo D 2018 Target selection for a Asteroid Impact Mission (AIM), 5th Interplanetary Mehdi Scoubeau, Daniela Taubert, ESA, ESOC, small low-thrust mission to near-Earth asteroids CubeSat Workshop Darmstadt, Germany. Astrodynamics 2(3) 249 Walker R et al. 2017a LUnar CubeSats for

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