Acta Astronautica

Contents lists available at ScienceDirectActa Astronautica

Acta Astronautica 106 (2015) 6389http://d0094-57

AbbreAstronoYear; GHAntennAntennMain EnNationaROM, RPropulsReadine

Thisn CorrE-mjournal homepage: www.elsevier.com/locate/actaastroA lander mission to probe subglacial water on Saturn's moonEnceladus for life$

Konstantinos Konstantinidis a,n, Claudio L. Flores Martinez b, Bernd Dachwald c,Andreas Ohndorf d, Paul Dykta a, Pascal Bowitz a, Martin Rudolph a, Ilya Digel c,Julia Kowalski c, Konstantin Voigt a, Roger Frstner a

a Institute for Space Technology and Space Applications, Bundeswehr University Munich, Werner-Heisenberg-Weg 39, 85579 Neubiberg,Bavaria, Germanyb Centre for Organismal Studies,University of Heidelberg, Im Neuenheimer Feld 234, 69120 Heidelberg, Baden-Wrttemberg, Germanyc Faculty of Aerospace Engineering, FH Aachen University of Applied Sciences, Hohenstaufenallee 6, 52064 Aachen, Germanyd Deutsches Zentrum fr Luft- und Raumfahrt e.V. (DLR), Oberpfaffenhoffen, Bavaria, Germanya r t i c l e i n f o

Article history:Received 20 December 2013Received in revised form16 September 2014Accepted 20 September 2014Available online 6 October 2014

Keywords:EnceladusLanderMelting probeAstrobiologyIcy moonsx.doi.org/10.1016/j.actaastro.2014.09.01265/& 2014 IAA. Published by Elsevier Ltd. A

viations: (N)EP, (Nuclear) Electric Propulsiomical Unit; C&DH, Command and Data Hande, Helium gas; GN&C, Guidance, Navigation

a; IMU, Inertial Measurement Unit; IPR, Ice Pa; LIDAR, Light Detection and Ranging (LIghgine Assembly; MGA, Medium Gain Antennl Science Foundation; PCA, Pressure Control Aough Order of Magnitude; RPS, Radioisotopeion; SIS, Site Imaging System; SOI, Saturn Orss Levelpaper was presented during the 64th IAC iesponding author. Tel.:49 89 6004 3592.ail address: [email protected] (K. Ka b s t r a c t

The plumes discovered by the Cassini mission emanating from the south pole of Saturn'smoon Enceladus and the unique chemistry found in them have fueled speculations thatEnceladus may harbor life. The presumed aquiferous fractures from which the plumesemanate would make a prime target in the search for extraterrestrial life and would bemore easily accessible than the moon's subglacial ocean.

A lander mission that is equipped with a subsurface maneuverable ice melting probewill be most suitable to assess the existence of life on Enceladus. A lander would have toland at a safe distance away from a plume source and melt its way to the inner wall of thefracture to analyze the plume subsurface liquids before potential biosignatures aredegraded or destroyed by exposure to the vacuum of space. A possible approach for thein situ detection of biosignatures in such samples can be based on the hypothesis ofuniversal evolutionary convergence, meaning that the independent and repeated emer-gence of life and certain adaptive traits is wide-spread throughout the cosmos. We thuspresent a hypothetical evolutionary trajectory leading towards the emergence of metha-nogenic chemoautotrophic microorganisms as the baseline for putative biological com-plexity on Enceladus. To detect their presence, several instruments are proposed that maybe taken aboard a future subglacial melting probe.

The Enceladus Explorer (EnEx) project funded by the German Space Administration(DLR), aims to develop a terrestrial navigation system for a subglacial research probe andeventually test it under realistic conditions in Antarctica using the EnEx-IceMole, a novelll rights reserved.

n; ACTC, Attitude Control Thruster Cluster; ASRG, Advanced Stirling Radioisotope Generator; AU,ling; EnEx, Enceladus Explorer project; EOI, Enceladus Orbit Insertion; EOM, End of Mission; FY, Fiscaland Control; GPR, Ground Penetrating Radar; HDA, Hazard Detection and Avoidance; HGA, High Gainenetrating Radar; Isp, Specific Impulse; kWe, kilowatt electric; kWt, kilowatt thermal; LGA, Low Gaint raDAR); LV, Launch Vehicle; MAG-L, Magnetometer-Lander; MCT, Minimal Convergent Trait; MEA,a; MLI, Multi-layer Insulation; MMRTG, Multi-Mission Radioisotope Thermoelectric Generator; NSF,ssembly; PFCA, Propellant Flow Control Assembly; RC, Reconnaissance Camera; RENC, Enceladus Radii;Power Source; RS, Saturn Radii; RTG, Radioisotope Thermoelectric Generator; SEP, Solar Electricbit Insertion; SPT, South-Polar Terrain; TBR, to be refined; TM, Thermal Mapper; TRL, Technology

n Beijing.

onstantinidis).

www.sciencedirect.com/science/journal/00945765www.elsevier.com/locate/actaastrohttp://dx.doi.org/10.1016/j.actaastro.2014.09.012http://dx.doi.org/10.1016/j.actaastro.2014.09.012http://dx.doi.org/10.1016/j.actaastro.2014.09.012http://crossmark.crossref.org/dialog/?doi=10.1016/j.actaastro.2014.09.012&domain=pdfhttp://crossmark.crossref.org/dialog/?doi=10.1016/j.actaastro.2014.09.012&domain=pdfhttp://crossmark.crossref.org/dialog/?doi=10.1016/j.actaastro.2014.09.012&domain=pdfmailto:[email protected]://dx.doi.org/10.1016/j.actaastro.2014.09.012

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638964maneuverable subsurface ice melting probe for clean sampling and in situ analysis of iceand subglacial liquids. As part of the EnEx project, an initial concept study is foreseen for alander mission to Enceladus to deploy the IceMole near one of the active water plumes onthe moon's South-Polar Terrain, where it will search for signatures of life.

The general mission concept is to place the Lander at a safe distance from an activeplume. The IceMole would then be deployed to melt its way through the ice crust to anaquiferous fracture at a depth of 100 m or more for an in situ examination for the presenceof microorganisms.

The driving requirement for the mission is the high energy demand by the IceMole tomelt through the cold Enceladan ices. This requirement is met by a nuclear reactorproviding 5 kW of electrical power. The nuclear reactor and the IceMole are placed on apallet lander platform. An Orbiter element is also foreseen, with the main function ofacting as a communications relay between Lander and Earth.

After launch, the Lander and Orbiter will perform the interplanetary transfer to Saturntogether, using the on-board nuclear reactor to power electric thrusters. After Saturn orbitinsertion, the Combined Spacecraft will continue using Nuclear Electric Propulsion toreach the orbit of Enceladus. After orbit insertion at Enceladus, the Orbiter will perform adetailed reconnaissance of the South-Polar Terrain. At the end of the reconnaissancephase, the Lander will separate from the Orbiter and an autonomously guided landingsequence will place it near one of the active vapor plumes. Once landed, the IceMole willbe deployed and start melting through the ice, while navigating around hazards andtowards a target subglacial aquiferous fracture.

An initial estimation of the mission's cost is given, as well as recommendations on thefurther development of enabling technologies. The planetary protection challenges posedby such a mission are also addressed.

& 2014 IAA. Published by Elsevier Ltd. All rights reserved.Fig. 1. Plumes spray water ice out from many locations along the TigerStripes near the south pole of Enceladus. Image credit: NASA/JPL/SpaceScience Institute.1. Enceladus

1.1. General description

Analyses of Cassini measurements imply a subsurfacesalt-water reservoir on Enceladus, where ice grains con-taining organic compounds escape via cryvolcanism fromwarm fractures in the ice, known as Tiger Stripes, atthe moon's south pole.

Enceladus, only 504 km in diameter and once believedtoo small to be active, has been found as one of the mostgeologically dynamic objects in the Solar System [1]. Calcu-lated from Enceladus' measured mass and density, modelsassume a rocky core with a radius of E169 km (density3 g/cm3) and a volatile crust with a thickness of E83 km(density 1.01 g/cm3) [1]. Despite its small size, it has a widerange of terrains, including old and young surfaces. The highalbedo of Enceladus results in a colder surface than most ofthe other Saturnian satellites, with a calculated subsolartemperature of 7573 K and an average temperature ofE51 K [2]. The spectrum of Enceladus shows that its surfaceis almost completely dominated by H2O ice, with smallamounts of NH3 and tholins [3], except near its south pole.At the south pole, the Cassini spacecraft has identified ageologically active province, circumscribed by a chain offolded ridges and troughs [4]. The terrain southward of thisboundary is distinguished by its albedo and color contrasts,elevated temperatures, extreme geologic youth (possibly asyoung as 5105 years), and narrow tectonic rifts thatcoincide with the hottest temperatures (145 K and more)measured in the region [4,5]. Much ongoing research workdeals with the origin and explanation of this puzzlingthermal anomaly. The most prominent feature at the southpole are four linear depressions, dubbed Tiger Stripesbecause of their appearance in the infrared. Particularly inthis region, simple organic compounds and CO2 are found(not as free CO2 ice, but rather complexed, probably withwater ice). This high abundance of complexed CO2 suggestsactive replenishment, probably from ongoing geophysicalactivity [6]. Multiple flybys of Cassini at Enceladus haveshown that plumes of H2O, including simple organic com-pounds [7], emanate via cryovolcanism from those warmfractures at the Tiger Stripes (Fig. 1). Salt-rich ice particles arefound to dominate the total mass flux of ejected solids (morethan 99%) [8]. Analysis of the plume material stronglyimplies that it originates from a body of liquid saltwater [8]or even a global ocean [9] below its icy crust.

The unique chemistry found in the plume has fueledspeculations that Enceladus may harbor life [10,11], butthis question can probably not be resolved by the Cassinimission. A lander mission that is equipped with a subsur-face ice melting probe, however, might be able to answerthis question. Because it is considered too risky to landclose to the cracks from which the plumes emanate (seeFig. 2), a Lander would have to land at a safe distance away

Fig. 2. Rough terrain at Enceladus' south pole with boulders resting alongthe tops of high frozen ridges (edited from the original raw image toenhance detail). Image credits: NASA/JPL/Space Science Institute,Universe Today.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 65from a crack and melt its way to the inner wall of a crack toanalyze the plume material in situ. This way, subsurfaceliquids may be much easier to access. Nevertheless, such amission is still considered very challenging, given thecurrent state-of-the-art in space technology.

1.2. Emergence and evolution of putative life on Enceladus

1.2.1. Evolutionary astrobiologyOne of the most intensely debated problems in evolu-

tionary biology, aside from the enigma posed by the origin oflife itself, is the existence of universal laws governing theemergence and subsequent evolution of biological complex-ity through natural selection [12,13]. While the force ofnatural selection in driving adaptation of species to changingenvironments is undisputed, most evolutionary biologistsfreely admit that they can hardly give a comprehensiveaccount of the major transitions that occurred during thehistory of life on Earth, such as, for instance, the bridging ofprebiotic, protocellular and cellular evolution, multiple endo-symbiosis events in ancient microbes, the emergence ofeukaryotic and multicellular organisms and the suddenappearance of complex body plans in the Cambrian. Yet,astrobiologists are especially interested in understandingthese macroevolutionary shifts in biological complexity as amore detailed picture thereof would shed light on thepossible existence or absence of analogous evolutionarytrajectories in extraterrestrial life. One influential school ofthought in evolutionary theorizing has stressed the role ofcontingency in shaping extant phylogenetic diversity andform [14,15]. This view of evolution contends that molecularand morphological traits are essentially frozen accidents ofmomentary ruling selective pressures and it is therebydenying the presence of laws of nature which might struc-ture and necessitate the emergence of life,macroevolutionary change and an overall increase in biolo-gical complexity. The alternate perspective, which opposesthe dominating influence of contingency, is adopted by manytheorists of astrobiology on the basis of the pervasivephenomenon of convergent evolution, the independent andrepeated emergence of similar adaptive traits in distantlyrelated lineages [1622]. This new paradigm proposes thatevolutionary processes can imaginably be extended beyondEarth and it connects the onset of biological complexity tothe vast timescales of cosmological history, the differentia-tion of astrophysical objects and ultimately redefines it as asubset within a vast phase space of universal complexity[2325]. A new interdisciplinary branch of astrobiology isexploring the hypothesis that genetically (as well as spatio-temporally) separated biological systems, for example onEarth and Enceladus, could evolve functionally equivalentadaptations by means of convergent evolution. In contrast tothe orthodox notion of contingency-driven evolution, thenovel hypothesis of cosmologically extended convergentprocesses lends predictability to the astrobiological endeavor[26]. Within the astrobiological landscape [27] or hyper-space [28] of biological complexity certain solutions of lifein dealing with environmental pressures are favored andthus emerge in a convergent manner across planetary bio-spheres, much like dynamical systems are tending to evolvetowards physical attractors. Therefore, astrobiologists shouldbe able to constrain the nature of extraterrestrial life bystudying convergent evolution on Earth. Eventually theinsight gained through such an approach could be translatedinto the evolutionarily informed design of biosignature-detection instruments bound for in situ exploration ofastrobiological Solar System targets.

1.2.2. Origin, evolution and baseline complexity of life onEnceladus

There is currently no consensus among evolutionary-and astrobiologists in regard to the exact mechanism thatled to the origin of life on Earth. Here we adopt ahydrothermal system-centered view for the emergenceof life on Enceladus because such an environment iscommonly suggested as the most likely habitat for nascentand extant biological activity on Enceladus [10,11,29].Although there is a small chance that ejecta from earlyEarth and Mars could have been transported to the outerSolar System, potentially carrying with them biologicalmaterial capable of proliferating after arriving at Encela-dus, an independent beginning of life appears more likelyin the light of the proposed concept of universal conver-gent evolution. Minimal Convergent Traits (MCTs) poten-tially found in life across different planetary habitats couldbe represented by 1) molecular replication and inheri-tance, 2) cellularization via membrane systems and 3)metabolic networks coupling energetically favorable bio-chemical reactions. At some point during the history ofEnceladus' putative biosphere, these supposed universallyconvergent features of life are expected to coalesce into acoherent biological entity, an organism, capable of under-going vertical (Darwinian) evolution rather than remain-ing in a primordial and poorly understood evolutionarymode that is dominated by horizontal gene transfer (HGT).Early molecular replication systems on Earth were most

Table 1Science-traceability matrix for in situ biosignature detection on Enceladus. Traceability is from science goal to science objectives to measurements todetection methods.

Science goal Scienceobjectives

Measurements Detection methods

In situ probing of an aquiferousfracture to search forbiosignatures

Molecularreplication

Analyze water sample for presence of replicatory polynucleotides Antibody microarray/sequencingNanopore-basedinstruments

Cellularization Analyze water sample for presence of cellular structures andmembrane material

MicroscopyFlow cytometryFatty acid markersMass spectrometry

Metabolicnetworks

Analyze water sample for presence of expected metabolicproducts (CH4)

Mass spectrometry

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638966likely defined by such a protocellular stage of evolutionbefore the Last Universal Common Ancestor (LUCA) of allterrestrial life made the transition towards a cellular formof organization [30]. How early life achieved this feat is notknown. Nonetheless, the directional shift in the flow ofgenetic material marks the essential evolutionary step thatallowed for the subsequent increase in biological complex-ity of cellular life by means of natural selection. How, then,could life conceivably emerge within hydrothermal sys-tems on Enceladus in the first place? Life on Earth hasevolved in a manner suggestive of a minimum set ofprerequisites, which comprises liquid water, biogenic ele-ments (C, H, N, O, P, S) and biologically usable energy (i.e.transducible into chemical bonds for storage and laterusage in metabolic reactions). In regard to the mostpressing constraint for the onset of a potential biosphereon Enceladus, namely available energy sources, recentstudies including theoretical work and on-going hydro-thermal laboratory experimentation have emphasized theimportance of evolutionarily conserved molecular dise-quilibrium converting engines in the protometabolism ofearly terrestrial life. This work is part of the SubmarineHydrothermal Alkaline Spring Theory for the emergenceof life [31,32]. It posits that off-ridge alkaline hydrothermalsprings reacted with the slightly acidic, and metal andCO2-rich Hadean Ocean. During this reaction, compoundssuch as H2, CH4 and NH3, as well as calcium and traces ofacetate, molybdenum and tungsten were released byprogressing serpentinization of ultramafic rock, therebyoffering a sustained source of chemically transducibleenergy for early biological systems. Employing an ancientmolecular apparatus resembling the ATP synthase com-plex commonly observed in extant life to harvest the freeenergy contained in the resulting proton and redox gra-dients (H2 and CH4 acting as electron donors, CO2 as anelectron acceptor), these ancestral proto-life forms couldhave been able to maintain and gradually increasethe complexity of their metabolic and replicator systems[3033]. Such a scenario for the origin of life is not onlypossible for the Hadean ocean on Earth, but also forsustained liquid water reservoirs in contact with thesilicate core of Enceladus, which could contain portionsof ultramafic rock. Simulations have shown thatserpentinization-driven cracking of Enceladus' core couldoccur to a depth of about 172 km, suggesting that emer-ging and sustained biological activity could be envisionednot only at the water-rock interface but down into the coreitself [34]. Having evolved from ancestral life forms deriv-ing from geochemical-disequilibrium-converting enginesinto cellular organisms displaying MCTs, as describedabove, microbial life on Enceladus would resemble themethanogenic chemoautotrophs found at terrestrialhydrothermal vents and seafloor sediment. These extre-mophilic microorganisms from the domain of the archaeacan perform their primary biosynthetic processes in theabsence of light by feeding off molecular hydrogen andconsuming CO2, thereby releasing CH4 as a metabolicbyproduct. Being evolutionarily ancient, extant species ofmethanogens can be found at a number of terrestrialanalog environments possibly resembling putative habi-tats on Enceladus.

1.2.3. Instrument suite for biosignature detectionA first consideration for a possible suite of in situ instru-

ments is based on the hypothesis of convergent evolution,leading to the emergence of methanogen-like organisms onEnceladus. Given the limited knowledge about the potentialmicrobial habitat in terms of prevailing environmental condi-tions, the assumption of convergence has been made toconstrain the vast landscape of biological complexity. Ascience-traceability matrix for a set of envisioned instrumentscan be found in Table 1. Aside from the goal of in situ probingof an active aquiferous fracture for biosignatures, the resultingobjective is the detection of extant life that potentiallyincorporates the previously mentioned MCTs, namely mole-cular replication, cellularization and metabolic networks.Several of the proposed instruments are routinely used inthe (biological) exploration of extreme habitats on Earth suchas deep-sea hydrothermal vents and sediment, subglaciallakes, and the continental subsurface [3540]. Due to volumeconstraints, however, only a limited number of instrumentswould be able to be integrated into an ice melting probe. Interms of the science goal which is the in situ probing of anaquiferous fracture to search for biosignatures, measure-ments relating to the molecular replication and cellularizationof putative life appear especially promising. Techniques for

Fig. 3. Polar stereographic map of Enceladus' South-Polar Terrain (SPT)showing the location of 100 plume sources. The circles are the 2uncertainties [50]. (For interpretation of the references to color in thisfigure legend, the reader is referred to the web version of this article.)

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 67isolating and characterizing replicatory polynucleotides, DNAand RNA in the case of terrestrial biology, are highly specificand often require basic knowledge about certain sequencemotifs. Furthermore, it is far from certain if polynucleotide-based replicatory macromolecules necessarily appear as auniversal link between chemical and biological evolution indifferent planetary environments. However, it has beenrecently suggested that a novel technique employingnanopore-based instruments could circumvent these pro-blems by potentially recognizing a vast set of different linearand water-soluble biopolymers [107,108]. Such moleculeswould most likely contain the genetic information of micro-bial life in the aqueous environment of Enceladus. Althoughthe detection of specifically terrestrial biopolymers is proble-matic during in situ exploration of Enceladus, an approachbased on multiplex antibody microarrays could target a largevariety of potentially biologically-derived molecules and iden-tify these as bona fide signatures of life [42].

On Earth, there was likely a distinct emergence of cellmembranes in archaea, based on glycerol-ether lipids withisoprenoid side-chains rather than glycerol-ester lipids com-posed of fatty acid tails as found in bacteria and eukarya[43,44]. A number of other amphiphilic molecules couldpotentially form spherical compartments, micelles, whichmight function as confined spaces in which metabolic andreplicatory reactions could be coupled. Cellular organizationcould thus be assessed and visualized using methods such asmicroscopy and flow cytometry. Different chemical labelingtechniques, for example via fluorescence, would target avariety of potential biomolecule classes and could even aid inresolving subcellular structures that should be highly indi-cative of the underlying biological complexity responsible forthe origin of a given sample. Lipids have been suggested asuniversal biomarkers for terrestrial and alien life forms andcertain lipid compositions in a sample, measured by a gaschromatograph-mass spectrometer, would be highly indica-tive of a biological origin [41].

As already described above, biologically-producedmethane presents yet another convergent biosignaturethat could be associated with the MCT of metabolic net-works. A gas chromatograph-mass spectrometer would beable to distinguish, via distinct isotope ratios resultingfrom biological processing, between biogenic methane andenvironmental sources such as hydrothermal systems.Since the envisioned methanogens should be endogenousto the subsurface ocean, it is questionable if they aremetabolically active after having been transported in thewater plumes to near-surface aquiferous fractures and ifbiogenic methane could be detected in these locales.

Current biotechnology and routine procedures in labora-tory experimentation employ desktop-scale platforms forsemi- or fully automated microscopy and flow cytometry[4547]. These could imaginably be further miniaturized andtransformed into biosignature instrumentation for a subgla-cial melting probe that is bound for planetary exploration ofextreme environments on Earth and the icy moons of theouter Solar System. Such in situ biosignature detection, forexample in the case of flow cytometry, requires a complexautomated workflow from environmental sampling, pre-paration of the samples, and analysis [48,49].The exact design and calibration for the biologicalexperiments should be accompanied by ongoing theore-tical and experimental work in the field of astrobiologyinto the possible nature of extraterrestrial biomolecu-les, organismal organization and metabolic versatility.An approach founded on convergent evolution is meantto constrain but not to theoretically limit the potentialitiesof evolution on planetary bodies apart from Earth. Wesuggest that the chosen methodology can facilitate thedesign and implementation of instruments aimed atdetecting biosignatures into future space missions. Inaddition, it promotes the development of prototype tech-nology that could be conceivably deployed in explorationefforts on Earth.

1.3. The topography and environment of the south-polarterrain of Enceladus

The South-Polar Terrain (SPT) is host to the plumesources, the target of the mission. The most prominentfeatures characterizing the interior of the South-PolarTerrain as mentioned above are the Tiger Stripe valleys,which include 101 identified distinct vapor and ice jetswhich form the plumes towering above Enceladus [50](Fig. 3). The term Tiger Stripe describes a V shapedvalley enclosed by two, nearly parallel ridges. These ridgesare about 100150 m high, while the valley is about 200250 m deep. The total width of the formation is about 25 km. The South-Polar Terrain features 4 valleys (Damas-cus, Baghdad, Cairo and Alexandria Sulcus) separated fromeach other by 35 km wide plains. As one moves closer to aTiger Stripe valley, the terrain rises with a relatively gentlegradient to the summits of the ridges. The terrain texturealso changes from highly fractured to a more undulating

Table 2Landing site selection traceability matrix.

Landing site selection criteria Measurements Instruments

Closeness to a plume source Perform high resolution mapping Hi-ResReconnaissanceCamera

Sublacial morphology Assess aquiferous fracture morphology for landingsite selection purposes

Ice Penetrating Radar

Thermal emissions (higher plume activity means a greater likelihoodof water nearer to surface)

Perform thermal mapping Thermal mapper

Thermal emissions (higher melting efficiency in warmer ice) Perform thermal mapping Thermal mapper

Surface roughness Estimate distribution of smaller features, surfacetexture

Radar

Measure ice block abundance Thermal mapper

Terrain morphology Detect slopes, large features and other landinghazards

Hi-ResReconnaissanceCamera

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638968one, covered with numerous icy blocks. Once over thelateral ridges, the slope is initially steeper but changes to amore moderate gradient the deeper one descends, withunconsolidated material sliding from steeper down toflatter sections, where material tends to accumulate. Thevalley floor is expected to be very narrow, in the order of50100 m [51], and interspersed with obstacles, such asnarrow ridges and elongated domes. An indicative pictureof the South-Polar Terrain at Enceladus can be seen inFig. 2.

Another important aspect is the texture of the surfacematerial in the South-Polar Terrain. It is mainly the falloutfrom the plumes that modifies the texture: ice particlestend to accumulate more near the plumes and less furtheraway. Studies have shown that nearer the plumes, particledeposition rates can reach up to 0.5 mm/year, indicating adeposit layer thickness of tens of meters if we assume thatthe plumes have been active in the past million years [52].Still, exposed icy crust can also be encountered, especiallyon slopes on which less consolidated material has sliddownward.

The mechanical behavior of the fallout is also crucial tothe understanding of the surface texture. In a firstapproach the fallout can be approximately treated assuper-fine snow, comprising grains of about 7.5 m out-side the SPT, 40 m in the vicinity of the Tiger Stripes and75 m inside the valleys, where larger particles tend to fallnearer to the plumes [52]. Grains are expected to have losttheir crystalline shape due to collisions with the ventwalls, and have a roughly round shape. These microscopicproperties are translated to macroscopic properties of thesurface material, namely: increased force transmissioncapacity due to the fine grain size, non-consolidated layersof material and increased compressibility due to the lowgravity.

Cassini data indicate that the South-Polar Terrain iswarmer than would be expected if it were heated only bysunlight. Moreover, the ice is warmer within the TigerStripe fractures, with observed temperatures reaching upto 157 K, significantly warmer than the expected 68 K forthis region of Enceladus [5].The radiation environment in the magnetosphere ofSaturn is not as severe as e.g. that of Jupiter. Moreover,Enceladus sweeps up energetic particles along its orbit,resulting in significant reduction in radiation levels in loworbits around it and on the moon's surface [53]. The icesof Enceladus also provide a reduction of orders of magni-tude in radiation barely a few centimeters under thesurface [54].

1.4. Landing-site selection

A critical aspect for a lander mission to deploy asubglacial melting probe near one of the plumes onEnceladus is the choice of landing site. Based on Sections1.2 and 1.3, we can derive the requirements such a landingsite must satisfy. First, it should be scientifically interest-ing, i.e. in this case close enough to an active plume source,where the water in the aquiferous fracture is close enoughto the surface, and possessing other potentially scientifi-cally interesting characteristics. In addition, the landingsite must satisfy certain landing safety criteria, such as theabsence of steep slopes, boulders and other hazards for alander. The instrumentation needed to assess these twoparameters is derived from the above two criteria. Therespective traceability matrix is given in Table 2.

2. The Enceladus explorer project (EnEx)

Since 2012 the joint research collaboration EnceladusExplorer (EnEx) funded by the German Space Administra-tion (DLR), investigates necessary technologies for a futureexploration of the Saturnian moon Enceladus. The goal isthe development of a terrestrial navigation system for asubglacial research probe that fulfills the followingrequirements: detection of liquid water in a body of iceand measurement of the distance to a set target point;continuous determination of the attitude and position inice relative to a surface station and to the target point;detection of gaps and obstacles in the ice; autonomousdetermination of the optimal route to the target point. The

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 69EnEx consortium is led by FH Aachen University of AppliedSciences and consists of eight partners, namely:

typFH Aachen University of Applied Sciences, Germany(Bernd Dachwald and team from the Faculty of Aero-space Engineering, Gerhard Artmann and team fromthe Institute of Bioengineering). Bundeswehr University Munich, Germany (Bernd Eiss-feller and team, Roger Frstner and team, both from theInstitute of Space Technology & Space Applications). University of Bremen, Germany (Kerstin Schill and teamfrom the Cognitive Neuroinformatics Group of theFaculty of Computer Science). Braunschweig University of Technology, Germany (PeterHecker and team from the Institute of Flight Guidance). RWTH Aachen University, Germany (Christopher Wie-busch and team from the Astroparticle Physics Group ofthe Faculty of Physics). University of Wuppertal, Germany (Klaus Helbing andteam from the Astroparticle Physics Group of thePhysics Department).

The developed navigation solution is integrated into aresearch melting probe of the IceMole type, a meltingprobe concept, which has been developed at FH AachenUniversity of Applied Sciences since 2008. We will refer tothis integrated system as the EnEx-IceMole throughoutthis paper1. It is validated and tested within field testsconducted as part of an ongoing collaboration betweenEnEx and MIDGE, an Antarctic exploration initiativefunded by the US National Science Foundation (NSF) [55].

In the context of the EnEx project, the Institute of SpaceTechnology & Space Applications (ISTA) of the BundeswehrUniversity Munich is responsible for the overall missionand system concept for a mission to land near a plume atEnceladus and deploy the IceMole there. The missionscenario for the extraterrestrial application of such a probeshould be studied, in order to determine the conditions forthe complete design and extraterrestrial operation of thesubglacial navigation system in detail.

2.1. The EnEx-IceMole: design and tests

Ice melting probes have been used since the 1960s forterrestrial research in ice [56]. Those pure melting probes,however, have two major drawbacks: they penetrate onlyvertically downwards and it is difficult to (intentionally)change their direction. In addition they cannot penetratedust/sand/grit layers. The EnEx-IceMole remedies thesedrawbacks of traditional ice melting probes. Its designis based on the novel concept of combined melting anddrilling (or more precisely screwing) with a hollow icescrew as it is used in mountaineering and a meltinghead at the tip of the probe (see Fig. 4). The screw assuresclose-contact between the melting head and the ice, whichoptimizes conductive heat transfer into ice and aids steeringin the desired direction.1 Whereas we mean a yet to be built melting probe of the IceMolee when referring to the IceMole.The current IceMole bodies have the shape of a rectan-gular tube with a 15 cm15 cm cross section. Their lengthdepends on the payload dimensions and the desiredmaneuverability. The rectangular shape of the EnEx-IceMole provides support for the torque of the rotatingice screw. The EnEx-IceMole can change melting directionby differential heating of the melting head, which gener-ates a torque (perpendicular to the longitudinal axis of theEnEx-IceMole) that forces the probe into a curve. Side wallheaters support its curve-driving capability. The requiredelectric power for the melting head, the ice-screw motor,as well as for the payloads and the other subsystems, isgenerated by a surface generator and transmitted via athree-conductor power cable. In future IceMole versions,this cable can be uncoiled from the probe to overcomerefreezing of the melting channel, and for depths of severalhundred meters the cable can be packed into separatecable containers, each containing tens of meters of lengthof cable, that are attached to the back of the IceMole.Communication and data transfer to the surface can beperformed via the power cable or via an Ethernet cable.

For navigation, the EnEx-IceMole is equiped with aninertial measurement unit (IMU) and a magnetometer tomeasure its attitude. A second (reference) magnetometeris installed at the surface ground station. Together with theadvance of the ice screw, the attitude history from the IMUand the magnetometer will then be used to calculate theprobe's position. Four ultrasonic phased arrays in themelting head are designed to detect the targeted aquifer-ous fracture as well as potential obstacles in the ice. Inaddition, acoustic pingers on the surface emit signals thatare received by four separate on-board detectors. Theyallow for an independent determination of the probe'sposition using the principle of trilateration. A multi-sensordata fusion system combines raw data streams from thenavigation sensors to generate a consistent scenario and todisplay it in a user-friendly format to support the decision-making process of the human probe operator on thesurface. In the future, this may serve as a basis for theautonomous control of the probe. This system will alsoinclude a module to optimize the probe's trajectory in theice with respect to parameters such as available resources,time, and risk. The EnEx-IceMole is described in moredetail in [57] and its navigation system in [58].

Several field tests of the various IceMole versions since2010 have been conducted on the Swiss Morteratsch andIcelandic Hofsjkull glaciers. The tests have successfullydemonstrated horizontal, upward, and downward meltingcapabilities and maneuverability, the stability and theinteroperability of the newly developed navigation sub-systems as well as the clean access and sampling sub-system. The latest IceMole version (EnEx-IceMole, seeFig. 4) was tested in NovemberDecember 2013 on theCanada glacier in Antarctica (see Fig. 5). The EnEx projectwill culminate in the Antarctic field season 2014/15 withthe deployment of the EnEx-IceMole at Blood Falls, aunique glacial feature at the terminus of the Taylor Glacierin the McMurdo Dry Valleys, Antarctica, where it willcleanly collect samples for chemical and microbiologicalanalyses within a collaborative exploration missionbetween EnEx and MIDGE (NSF funded) [55].

Fig. 4. Technical drawings of the current EnEx-IceMole design (top) and the EnEx-IceMole head (bottom) as tested on the Canada glacier, Antarctica(NovemberDecember 2013) [59].

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389703. The current mission concept

3.1. Mission statement

As part of the EnEx project, a mission and operationalconcept to deploy and operate the IceMole subglacialprobe near one of the plume sources on the surface ofEnceladus is being developed at the Institute for SpaceTechnology and Space Applications (ISTA) of the Bundes-wehr University Munich. The objective of the missionpresented here is to deploy a Lander with an IceMoleprobe at the surface of Enceladus. Because landing in closevicinity to a water-bearing fracture is too risky, we proposeto land at a safe distance and use the maneuverableIceMole probe to navigate to an aquiferous fracture at adepth of 100 m or more below the surface. Once there, theprobe can sample and analyze the materials in the frac-ture. Navigating from the Lander to the targeted fracturerequires an elaborate navigation solution that can detectthe aquiferous fracture but also potential obstacles alongthe way (e.g., dry fractures and mineral inclusions).

The mission concept study is a work in progress. Designfeatures are therefore not final and could change until thefinal mission concept definition.3.2. Mission architecture trades

In the following we focus on the driving architectureelements for the mission. Decisions on these elements arecrucial, and will shape the subsequent design of thesystem.

Fig. 6. Required heating power P as a function of melting velocity v,without (solid lines) and with conductive losses (broken lines).

Fig. 5. Field testing of the IceMole by the FH Aachen operations team onCanada Glacier, Antarctica during NovemberDecember 2013. Credit: FHAachen.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 713.2.1. Power sourceDue to the high energy demand by the IceMole, the

driving system element for the mission is the primarypower source.

As mentioned in Section 2.1, electricity from the powersource is converted to thermal power in the melting headson the IceMole, which in turn is conducted to the icemelting or sublimating it, depending on the ambientpressure. The pressure buildup between the melting headand the ice can be regulated by the screw, which pulls thehead against the ice. Thus enough pressure is produced toinduce ice melting, which is an order of magnitude moreefficient than sublimation.

The typical shape of a melting probe is that of acylinder with length L and diameter D2 R. In a simpleenergy balance approximation that neglects all losses, theminimum power requirement P0 to a given meltingvelocity v is given by:

P0 AcpTHv 3 1

where AR2 is the probe's cross-section, is the icedensity, and cp is the specific heat capacity of ice,TTpc Tis the difference between the phase change temperature Tpcof ice and the local ice temperature T (the phase change tem-perature is the melting temperature, Tm, or the sublimationtemperature, Ts, respectively), and H is the phase changeenthalpy.

Because the triple point of H2O is at a temperature of273.16 K (0.01 1C) and a pressure of 6.1173 mbar, the operat-ing mode of melting probes under vacuum conditions issublimation, unless the closing of the hole (by refreezing orcovering) raises the pressure in the hole above the triplepoint pressure, which happens quite rapidly. We thusassume in the following that the appropriate measures aretaken, so that the operating mode of the probe is melting.

To obtain a more accurate result, losses need to beconsidered, most importantly the losses due to lateralconduction into the ice. The lateral conduction losses havebeen estimated by Aamot [60] for a cylindrical probe. Theresulting numerically problematic equation for the lateralconduction losses can be approximated by:

Plc 932Ws

Km3Ux0:726vR2T 3 2

where x is the value for L/(vR2) in s/m2.Hence, the total power required by the probe to

penetrate the ice with a velocity v is given by:

P P0A; v; TPlcA; L; v; T 3 3

Fig. 6 shows the required heating power of a typicalmelting probe (L1 m, D10 cm) as a function of meltingvelocity for ice temperatures 50 KrTr270 K. The solidlines represent the minimum power requirement accord-ing to Eq. 3-1, while the broken lines depict the totalpower including losses according to Eq. 3-3.

In general, the melting efficiency (EP/P0) becomesvery small for melting velocitieso1 m/h and very cold ice.For example, in moderately cold 250-K Antarctic ice, theefficiency drops below 50% for melting velocitieso0.55 m/h, while in extremely cold 150-K Enceladan ice, theefficiency at this melting velocity is below 15%.

We can now estimate the power required by theIceMole to melt through ice on the south pole of Encela-dus, by making several assumptions. First, the ice tem-perature near the plumes is approximated at around 150 Kas discussed in Section 1.3. As mentioned above, theIceMole is assumed to operate in the melting regime.The IceMole is also approximated by a cylinder with alength of 1 m and a radius of 10 cm, and its melting speedis set to 1 m/h. From Fig. 6 we thus get an approximatevalue for the required input electrical power of 5 kW.

For the stated melting speed, and assuming that theIceMole will only be operated for 50% of the time, and alsothat it will melt through 100 m of ice or more, the durationof operation until it reaches its subglacial target is no lessthan 9 days.

Power of this order of magnitude and for that durationat large heliocentric distances where sunlight is notintense enough for solar power generation, can only be

Fig. 7. Regime diagram showing the typical applicability of variousenergy sources in the Saturn system (10 AU heliocentric) as a functionof power and duration. The yellow shaded area is an estimate of utility ofsolar power at 10 AU, while the yellow dotted line is an estimate at 1 AUfor reference. Adapted from [61]. (For interpretation of the references tocolor in this figure legend, the reader is referred to the web version of thisarticle.)

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638972provided using nuclear power sources, either radioactiveisotopes, or a space nuclear reactor (Fig. 7).

The principle of operation of space-based nuclearpower systems is converting heat from a radioactive heatsource to electricity. Radioisotope power sources (RPS)using plutonium-238 as the heat source and dynamicconversion methods such as Stirling converters are themost efficient option, with a maximum achievable specificpower for this technology of up to 9 W/kg [62]. These highperformance Stirling RPSs would be an evolution of theAdvanced Stirling Radioisotope Generator (ASRG). How-ever there are significant future Pu-238 availability issues:although production is scheduled to be restarted in the USaround 2020 at the rate of 1 kg/year [63], supplies are stillexpected to be limited and hard to obtain. In addition,further development of the Advanced Stirling RadioisotopeGenerator (ASRG) was recently canceled, which represents asignificant setback for this technology [64]. A similar evolvedASRG design using Americium-241 as the heat source couldconceivably be used. This system would be somewhat morecomplex than the Pu-238 ASRGs due to the existence ofpumped loops to conduct heat from the radioisotope heatsource to the Stirling converter rather than using simplecontact conduction. The maximum specific power of such apower source is about 4W/kg [65]. Am-241 is currentlyselected as the European fuel/heat source baseline anddecision on the development of a full-scale production plantin Sellafield, England will be made by 2015. The estimatedyearly production rate will be between 10 and 17 kg [66],making it potentially much more readily available andcheaper than Pu-238. However there are no plans for thedevelopment of such a high performance Am-241 basedsystem, with the focus being instead on simpler systemswith smaller specific power, similar in design to the currentMulti-Mission Radioisotope Thermoelectric Generator(MMRTG). A plausible limit for power production using RPSsis considered to be around 2 kWe [67].

Using a nuclear reactor would circumvent the problemof fuel availability. Space nuclear reactors have significantheritage, however not for landed applications. Of the threeoptions, it is the most complex system. Heat is transferredfrom the reactor core using heat pipes to compactthermoelectric heat exchangers for power levels of up to5 kWe or Stirling converters for power levels of up to10 kWe [68]. A nuclear reactor would remain subcriticalduring plausible launch failure scenarios. The specificpower of this kind of small nuclear reactors is between 3and 7 W/kg [69].

High performance fuel cells were also considered as theprimary power source for the IceMole. However, due totheir low technological maturity and the uncertaintyconcerning the IceMole operations duration on Enceladusthey were discounted. In case a nuclear reactor provesunfeasible due to Planetary Protection concerns (Section4), this option will have to be reinvestigated.

Due to the above, a small nuclear reactor is chosen asthe baseline for the mission. The resulting planetaryprotection issues are addressed in Section 4.

3.2.2. Relay Orbiter elementOne of the basic trades critical to the system architec-

ture concerns whether to use two separate elements, anOrbiter around Enceladus or Saturn to act as a relay and aLander (OL), or one single element, dubbed an Orbi-lander (OL) [70] communicating directly with Earth. Theconfiguration chosen in turn influences the communica-tions architecture between Lander and Earth.

The OL option is clearly advantageous in terms of savedcost, launch mass (no need for an extra functional Orbiter)and reduced operational and systems complexity becausecommunication takes place directly from the Lander to Earth.The deciding factor in this case, however, is that the mostprobable location of the candidate landing spots is near thebottom of one of the canyons that form the Tiger Stripes.They are located at very low southern latitudes near the poleand are characterized by rugged surrounding terrain (seeFig. 2).

The above means that the Earth will be visible from theOrbilander only in low elevation angles, and only duringSummer at the south pole of Enceladus (20252040) whenthe Sun and the Earth are above the horizon, with theassociated high risk that the Earth will be hidden by localterrain features, or that the antenna mobility will be compro-mised after landing on a large but survivable slope. Alterna-tively, we could limit our candidate landing sites to the onesfurther away from the pole to try and circumvent this problem(the northernmost plume source is at around 701 latitude,Fig. 3). This, however, would impose unnecessary extraconstraints to the mission. Due the above reasons, theseparate Orbiter and Lander option is chosen for the mission.

3.2.3. Propulsion method and trajectoriesThe propulsion method used and the flown trajectories

up to Enceladus landing are interconnected, the latterdepending on the type of thrust used (high/chemical orlow/electric thrust) and the method used to capture inorbit around Saturn. The options considered are: All chemical propulsion (Chem):The spacecraft will launch at an Earth escape trajectoryand will follow a gravity assisted interplanetary trajec-tory to Saturn, where an orbital insertion maneuver

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 73using the chemical propulsion system will be per-formed. From the capture orbit, the spacecraft will starta gravity assisted moon tour similar to the one pro-posed in the 2007 NASA GSFC Enceladus Flagship study[71], to shave off enough orbital energy to reach theorbit of Enceladus, and then capture in orbit around it. All electric propulsion (EP):The spacecraft will use the on-board small nuclearreactor to power electric thrusters (Nuclear ElectricPropulsion or NEP). It will follow a gravity assistedinterplanetary trajectory, with intermediate low thrustarcs, with a final thrust arc applied to capture aroundSaturn. After Saturn capture, the spacecraft will per-form a gravity assisted moon tour, while also perform-ing low thrust maneuvers, to reach and capture in orbitaround Enceladus. Electric propulsion using a chemical stage for Saturncapture (EPChem):This option is similar to the all electric option above,with the difference that no low thrust decelerating arcis applied in the final leg before arriving to Saturn,relying on a chemical stage for Saturn capture instead.The moon tour is then performed using low thrust,same as for the all electric option. Chemical propulsion using an Aeroshell to performAerocapture (ChemAero):The spacecraft is launched and follows an interplane-tary trajectory as described for the all chemical option.Once at Saturn, an Aerocapture maneuver is performedin the atmosphere of Titan. The spacecraft then reachesa Titan Enceladus orbit, fromwhich it begins a gravityassisted moon tour as described above to reach andcapture around Enceladus. Electric propulsion using an Aeroshell to performAerocapture (EPAero):The spacecraft follows the interplanetary trajectory ofthe EPChem option. The same Aerocapture maneuveris performed as described for the ChemAero optionabove. EnEx then starts the gravity assisted moon tourwhile also applying low thrust to reach Enceladus, asdescribed under the EP option.Options involving direct landing after Saturn orbitinsertion are discounted as too demanding in propellant.Also, in addition to Nuclear Electric Propulsion consideredTable 3Key mission architecture choices. The selected architecture elementabove, conceivably a Solar Electric Propulsion (SEP) stagecould be added to the spacecraft to power electric thrus-ters at higher energy levels and therefore significantlyreduce mission duration. Such a combined Solar-NuclearElectric Propulsion (SNEP) stage has been discussed e.g. in[72]. However adding a SEP stage while there is already anon-board nuclear reactor, albeit small, was consideredredundant and discounted as an option.

Options including chemical propulsion benefit from thehigh flight heritage of this propulsion method and its lowrelative operational and system complexity. Chemical pro-pulsion involves drawbacks however, such as increased pro-pellant mass due to its low specific impulse (Isp) and longermission duration due to the need for multiple planetaryflybys in the inner Solar System. Options involving electricpropulsion on the contrary have less flight heritage, parti-cularly in the outer Solar System and are more complexoperationally and as systems. However they offer consider-able mass savings and can shorten mission duration sig-nificantly. Aerocapture can potentially save significantpropellant mass as well because it eliminates the need fora capture maneuver. However, since in this case the targetbody lacks an atmosphere, the benefits of Aerocapture arereduced.

Considering that the landed mass on Enceladus is goingto be high, we prefer solutions that minimize launch mass.Since a relatively high-powered energy source is necessaryto power the IceMole anyway, it would be logical to use italso to power electric thrusters. Since the interplanetarytransfer duration is already within reasonable limits (o 10years) using only electric propulsion, there is no need forthe added mass and complexity of a chemical boost stagefor Saturn capture. Aerocapture would be best avoidedaltogether, due to its low technological readiness and highoperational risk.

A purely electric thrust system powered by the on-board small nuclear reactor thus emerges as the preferredsolution.3.2.4. Lander typeThe target of the mission is to land and deploy the

IceMole near a plume source that will most likely belocated on the floor of a Tiger Stripe valley. This wouldrequire high landing accuracy and autonomous hazarddetection and avoidance (HDA). Due to the long distancess are given in bold red letters.

Fig. 8. Preliminary design of the Combined Spacecraft, comprised of the Orbiter and Lander, before Lander separation.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638974and attendant signal delays involved, remote control of thelanding is not an option, and a high degree of landingautonomy must be employed.

Ball et al. [73] categorize landers based on the velocityat which the contact the target body and on their landingaccuracy: Pod landers can survive an initial landingimpact, which may send the vehicle rolling and/or boun-cing across the surface, and then commence operationshaving come to rest in whatever orientation is finally rea-ched. Penetrators are bullet-shaped vehicles designed topenetrate a surface at high velocity and emplace experi-ments at some depth. Legged landers use a system of legsto cushion the landing and provide a stable platform forsurface operations.

Due to the above constraints of the mission concept(high accuracy, soft landing, unknown terrain) the onlysuitable lander type is the legged lander, and it is thereforechosen as the baseline for the Lander.3.2.5. Architecture overviewTable 3 handily summarizes the main mission archi-

tecture characteristics chosen for the mission.Fig. 9. Preliminary launch configuration of the Combined Spacecraft, withina provisional launch vehicle fairing with dimensions of 5.417 m2 (ArianeV fairing).3.3. System description

The combined Orbiter and Lander before separationwill be referred to as the Combined Spacecraft. Theconfiguration of the Combined Spacecraft is shown inFig. 8. This Combined Spacecraft will transfer to Enceladusand communicate with Earth, with the Orbiter serving asthe propulsion module. It will also perform remote sensingof potential landing sites during the landing site recon-naissance phase. After Lander separation, the main func-tion of the Orbiter will be to relay data between landerand Earth.

The nuclear reactor on the Lander provides power tothe electric propulsion system, while three MMRTGsprovide power to the rest of the Orbiter subsystemsthroughout the mission lifetime. An extended fixed struc-ture is used to put distance between the reactor and theunshielded components of the Orbiter. Fig. 9 shows thelaunch configuration for the Combined Spacecraft.3.3.1. Orbiter3.3.1.1. Instruments. Table 4 shows a list of the instrumentsto be carried by the Orbiter, as derived from the landing siteselection traceability matrix (Table 2). All three instruments

Table 4Reference instrument list for the Orbiter.

Instrument Mass(kg)

Power (W) (Max) Data rate Dimensions Heritage

Recon camera (RC) 33.4 38 706 Mbps Approx. 1.51 1 m3 envelope MRO HiRise Camera [74]Ice Penetrating Radar(IPR)

17.128 (acquisition) 2400 kbps (in flyby

phases)16 m deployable long antenna MRO SHARAD, RIME [75]

13 (standby)

Thermal Mapper (TM) 11 14 0.6 Mbits/sApprox. 5050 50 cm3envelope

Mars Odyssey THEMIS[71]

Table 5Component list for the communications subsystem [76].

Component Mass (kg) Power (W) (Max) Dimensions (max. envelope)

AntennaHGA (incl. LGA-1)a 100.6 3-m diam. reflectorLGA-1 LGA-2 0.5 6.4-cm diameter, 33.5-cm length (max)Transmission lines 12.5

Radio frequency subsystemDeep Space Transponder (2 ) 8.0 10.2 18.83.011.5 cm3Command Detector Unit (2 ) 0.7 12.712.71.8 cm3Telemetry Control Unit 7.3 5.1 21.119.415.3 cm3Ultra Stable Oscillator 1.8 3.0 19.410.212.8 cm335 W X-band TWTA (2 )a 10.8 53.7 16.718.441.8 cm3Waveguide Transfer Switch (4 ) 1.5 8.24.310.5 cm3X-band diplexer (2 ) 3.4 passive 49.49.48.9 cm33 dB Hybrid coupler (1 ) 0.1 passive

Radio frequency instrument subsystem 41.842.517.8 cm335 W Ka-band TWTAa 4.9 33.7S-band Transmitter 2.7 41.3Ka-band Exciter 2.4 3.1Ka-band Translator 3.5 8.0Microwave Components 0.1

a Different from Cassini communications subsystem components, see text.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 75should be pointed towards the candidate landing site to besensed. The Ice Penetrating Radar can also sense the surfaceby operating at a different frequency.

3.3.1.2. Communications. For communications between theOrbiter and the Earth a system based on that of Cassini isforeseen, as described in [76], with the modification that a3-m diameter High Gain Antenna (HGA) is used instead ofthe original 4-m one, in order to make sure that thespacecraft can fit in the fairing. In order to retain theoriginal system gain, the X-Band and Ka-Band TWTAs(Traveling-Wave-Tube Amplifier) are replaced withhigher power versions (see Table 5). The HGAcommunicates at the X-Band. Two 6.4 cm33.5 cm UHFLow Gain Antennas (LGAs), one mounted on the HGA andthe other on the opposite side of the Orbiter, both withwide fields-of-view, provide uniform coverage.

According to [76] this system is capable of transmitting4 Gb in 9 h during high activity periods while in orbit inthe Saturn system, requiring the use of a 70-m receivingstation on Earth. This translates to a data rate of 130 kbps(link margin of 3 dB), which we will consider a maximumcapability of our system.

A list of the subsystem components is shown in Table 5.3.3.1.3. Command and data handling. During the missionthere are two sources of data that must be stored; beforeLander release the Reconnaissance Camera will takeimages that will be stored and then downlinked overtime. After Lander release, the Orbiter will act as a relayfor the Lander data. As discussed in Section 3.4.3, theOrbiter generates about 4 Gbit per pass over the South-Polar Terrain. This data can be stored multiple times in alarge, redundant, solid-state data recorder (6.4 Gbit percard) [74]. The data will then be transmitted back to Earthas described in Section 3.3.1.2. Based on [74], a single-fault-tolerant RAD750 single-board computer is selected. Itprovides greater than 6.4 Gbit of storage using multipleFlash memories as stated above.

3.3.1.4. Structures and mechanisms. A hexagonal primarystructure is foreseen for the Orbiter body. During theinterplanetary flight and before Lander separation, theLander and Orbiter will be connected with an extendedhexagonal structure, so that the Orbiter is at a safedistance from the radiation emmiting reactor.

For separation of the Lander mechanisms must beincluded, that provide a safe separation, imparting enoughV so that no recontact occurs. To avoid the potential for

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638976unstable dynamics of the Orbiter, the supporting longstructure might also need to be separated from the Orbitervia the use of additional separation mechanisms identicalto the ones used for Lander separation, but with stricterno-recontact requirements.

3.3.1.5. Power. Before Lander separation, the on-board rea-ctor provides power only to the electric thrusters. The drivingoperational phases for the Orbiter are the Reconnaissanceand Landed operations (data relay from the Lander) phases.For these phases, a maximum of 80W and 100W is requi-red respectively (Table 8 and Section 3.4.5 respectively). Wealso allocate an additional maximum power requirementof 100W for C&DH, GN&C, and thermal control [74]. Theref-ore, three MMRTGs [77] providing total End-of-Life (EOL)(after 15 years) power of 300W are foreseen to fulfill all thepower needs for the Orbiter with significant margin. Such aseparate power system reduces complexity of interfaces andadds robustness to the system.

3.3.1.6. Thermal. Excess heat from the on-board MMRTGon-board the Orbiter or even from the reactor on theLander before separation can be used to heat Orbitersubsystems to minimize electrical power consumption.The temperature of the instruments, electronics andpropellant tanks is regulated with the help of louvers.The entire Orbiter is covered with multi-layer insulation(MLI). A radiator of 1 m with louvers is foreseen for theOrbiter.

3.3.1.7. Propulsion. The combined Orbiter and Lander willreach the orbit of Enceladus and capture in orbit around itusing Nuclear Electric Propulsion. Power from the reactoris varied via a Power Processor Unit (PPU) to power theMain Engine Assembly (MEA) comprising two NEXTthrusters with a maximum input power of 5 kW. Thepropellant is pressure-regulated to maintain constantpropellant mass flow. Attitude control is performed byfour Attitude Control Thruster Clusters (ACTC) with fouroperational and four redundant thrusters each, fed withpropellant out of two hydrazine (N2H4) tanks mounted atthe sides of the Orbiter. Two more tanks contain the Xenonpropellant used for the NEP system.

3.3.1.8. Guidance, navigation and control (GN&C). The Orbiter'sGN&C system will have to serve both the combined Orbiterand Lander prior to Lander separation, and the Orbiter duringthe communications relay.

As mentioned in Section 3.3.1.7, there are four AttitudeControl Thruster Clusters. Three High Torque ReactionWheels are also included, mainly for fine pointing theremote sensing instruments during the reconnaissancephase. Redundant sun-sensors, Inertial MeasurementUnits (IMUs) and star-trackers are included for navigation.

3.3.2. Lander3.3.2.1. Configuration. A preliminary illustration of theLander as described below is shown in Fig. 10. Theconfiguration was based on the Mars Cryobot Landermission concept [78].A major driver for the Lander's design and configura-tion is the intense radiation emitted from the reactor, alsoduring the interplanetary transfer phase. Sensitive elec-tronics, such as the IceMole, instruments and Landeravionics will be protected behind radiation shielding inRadiation Protected Regions (Fig. 10). The Lander designwill need strenuous attention to radiation levels formaterials and equipment outside of these shielded plat-forms. The water ice environment on Enceladus is astrong absorber of neutrons, thus mitigating radiationdosage due to scattering. At this stage we have bookkeptan additional 500 kg for shielding. Detailed radiationdose analysis and shielding sizing is currently ongoingusing the LANL MCNP tool [79].

3.3.2.2. Instruments. Table 6 shows a list of the IceMolenavigation auxiliary instruments to be carried by theLander, as derived from the IceMole navigation require-ments (Section 2.1). A site imaging camera (Site ImagingSystem, SIS) is included in the Lander instrument suite, forlanding site context imaging.

The SIS will be placed on a raised platform, to have anunobstructed view of the surroundings. Because theLander will land deep within a Tiger Stripe valley, thereis the possibility that it will not be illuminated by the sunat all. A floodlight must then be added next to the camerato illuminate the immediate surroundings. The GroundPenetrating Radar (GPR) dual parallel antennas will bedeployed extending from the side of the Lander. Radiofrequency Pingers aiding in IceMole navigation will needto be deployed (Table 6).

3.3.2.3. Communications. The Lander will communicatewith Earth via the Orbiter that will function as acommunications relay. A pointable LGA is foreseen, thattracks the Orbiter when it is above the local horizon andtransmits stored data to it. This is necessary instead of a fixedzenith pointing antenna because the Orbiter will remain atan orbit with inclination close to 601, and therefore willremain close to the local horizon of the Lander on the South-Polar Terrain. Achieving high data rates between the Landerand the Orbiter is not hard, owing to the relatively smalldistances involved. A UHF helical antenna with a length of40 cm and a diameter of 3 cm can transmit close to 1.3 Mbps(3 dB margin) to an identical antenna on the Lander over375 km, the distance between the Orbiter and the Landerwhen the former rises above the local horizon. The systemhas a second LGA for redudancy.

3.3.2.4. Command and data handling (C&DH). Followingthe approach of [74], a RAD750 single-board computerwas selected. Onboard data storage will accommodatemultiple copies of the Lander and IceMole science andhousekeeping data.

3.3.2.5. Structures & mechanisms. The Lander body is atypical hexagonal pallet lander platform. Six lander legswith incorporated crushable honeycomb for shockattenuation are foreseen. The legs are crushable up to apredefined height from the surface, to allow for thenecessary clearance to be maintained below the Lander.

Table 6Reference instrument list for the Lander (IceMole navigation auxiliary instruments and Camera).

Instrument Mass(kg)

Power (W)(Max)

Data rate Dimensions Heritage

Site Imaging System(SIS)

2.8 4 1.2 Mbps Approx. 503030 cm3 envelope, on gimballedplatform, 1.0-m mast

MER Pancam [74]

Ground PenetratingRadar (GPR)

0.15 1 2400 kbps2 deployable antennas 2 m each

JPL RoverGPR [80]

Magnetometer(MAG-L)

2.2 4 1.8 kbps electronics: 1310.48.6 cm3, sensor:84.84.6 cm3, on 2 m deployable boom

MESSENGER MAG, Galileo MAG

Pingers (4x) 1 (each,TBR)

1 (each,TBR)

10 cm diameter spheres (TBR)

EnEx Pinger development team,personal communication

Fig. 10. Preliminary design of the Lander.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 77On the bottom of the landing legs snow shoes will beattached, that must be able to cope both with landing onvery soft snow and on hard and slippery ice. The snowshoes must therefore have both sufficient surface area andsome sort of spikes to grip on any underlying icy surface.The Lander structure must be elevated enough above thesurface to avoid protruding features like boulders thatcould damage the underside of the Lander, but also low

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638978enough so that the center of mass of the Lander ismaintained low, thus avoiding a tip-over.

The IceMole deployment mechanism must deploy theIceMole from its stowed position on the Lander and positionit above a patch of ice, suitable for the initiation of IceMoleoperations. An evolution of the deployment mechanismdesign currently used in the field testing campaigns of thecurrent EnEx-IceMole design is the current baseline.

Due to the high levels of radiation from the reactor,Lander avionics, reactor electronics and instruments aregoing to be placed on a fixed truss type mast at a heightof 4 m. On the top of that mast a radiation shield with aweight of 100 kg will be placed, on which shield the plat-form for the electronics mentioned above will rest. The toppart of the mast will be connected to the sides of the Landerwith guy wires for stability. Before deployment the IceMoleis also shielded from radiation from the reactor.

Ice particles and water vapor ejected from the plumescan also pose a hazard to sensitive surfaces, for examplethe lens of a camera on the Lander could be damaged byan ice particle or covered by a thin layer of ice, slowlyaccumulating over time. This will require the shielding ofsensitive upwards facing surfaces, similar to that used forprotection against micrometeoroids.

3.3.2.6. Power. The chosen power source for the mission isa small 5 kWe nuclear reactor. Several such small spacenuclear reactor designs exist (see e.g. [57]).

The currently baselined reactor is the Heatpipe Oper-ated Mars Exploration Reactor (HOMER) [81]. This reactorhas been proposed for landed applications on Mars, bothon a pallet lander [78] and a rover [82].

The main component of the reactor is the core. Inside it,thermal power is transmitted from the UO2 fuel rods toheatpipes. The criticality of the reactor is controlled byslidable control rods. The core is surrounded by a radialneutron reflector. The heatpipes transfer heat to a Stirlingconverter, where it is converted to electrical power. Theexcess heat is radiated. Assuming a conversion efficiency of20%, and for a required electric power output of 5 kWe, atotal of 25 kW of thermal power has to be produced by thereactor. Excess heat that is not used for other purposes, e.g.heating various components, is rejected via the radiator.

The total weight of the reactor is approximated at800 kg [81].

Two redundant batteries are included on the Lander fordescent and landing, and deployment of the IceMole [78].

3.3.2.7. Thermal. The thermal system of the Lander willmaintain avionics, batteries, propellant and instrumentswithin their allowed temperature ranges, through allphases of the mission. Excess heat from the nuclearreactor can be used to heat components as needed.

The driver for the thermal system is the HOMER on-board nuclear reactor. Due to its high thermal load, thesides of the reactor will be exposed to space. Excess heatwill be transferred to the fixed 4 m diameter radiator byheat pipes and removed via the top side. [81] suggest thatHOMER is very robustly designed from a thermal point ofview, albeit for a 3 kWe power output. In case furtheranalysis shows that more intensive thermal control isneeded, additional changes to the design can be made,like using pumped loops instead of heat pipes, addingradiators on the sides of the mast, etc.

The propellant tanks could conceivably use heatradiated from the reactor core to passively remain withinthe required temperature limits. They will be individuallycovered with MLI. Thermal control of instruments andelectronics in the IceMole after it has been inserted in theice can be managed locally, using e.g. the abundantelectrical power for electrical heating units.

A preliminary thermal distribution on the Lander isshown in Fig. 11. The top of the mast and the IceMole beforedeployment are kept within operational temperature ranges,e.g. via heatpipes and proper use of MLI. Propellant tanktemperatures were not calculated, but will be in the future,and measures will be taken to insulate them from heat givenoff by the reactor. The effects of water vapor from the plumeon radiators and other components of the thermal controlsubsystem will be looked into in the future as well.

3.3.2.8. Propulsion. The Lander propulsion system is a mon-opropellant system working on blowdown mode. The pro-pellant used is hydrazine and it is divided among fourtanks: two tanks to feed the Main Engine Assembly (MEA)and two to feed the Attitude Control Thruster Clusters(ACTC). The MEA is mounted at the bottom of the Lander,so that the thrust vector runs through the Lander's center ofgravity. Alternatively, each of the component thrusters ofthe MEA can be divided among the thruster clusters, thusfreeing the bottom of the Lander and allowing for a largerclearance between bottom of the Lander and hazardousterrain features. The total thrust provided by the MEA isestimated based on a thrust-to-weight ratio of 5. This givesa total thrust requirement of 1200 N, which will be appliedby four 400 N thrusters. Each of the three ACTCs is placedon a side of the Lander. Helium gas is used as thepressurant, stored in the same tank as hydrazine andseparated by elastomeric diaphragms.

3.3.2.9. Guidance, navigation and control (GN&C). Toachieve the goal of autonomous safe pinpoint landing, alanding Guidance, Navigation and Control system must beemployed, involving the necessary hardware (sensors,actuators) and software (algorithms). The top levelarchitecture of the Lander GN&C is illustrated in Fig. 12.

The GN&C software comprises several parts. Concern-ing guidance there are two different types: primaryguidance and piloting guidance. The primary guidancecomputes a trajectory for the landing craft and the pilotingguidance is a simplified guidance that chooses a landingsite and a trajectory or acceleration profile is computed foreach of them. These trajectories are compared in a decisionmaking process and if necessary a new landing site ischosen by the piloting guidance. Both types of guidancehave different requirements. Furthermore both types needinformation about the current state of the landing vehicle.This information is provided by the navigation algorithm.A Hazard Detection and Avoidance system is necessary todetermine a safe landing site, and in case of hazardsinitiate avoidance maneuvers and retargeting.

Fig. 12. Illustration of a top-level GN&C system architecture [83].

Fig. 11. Thermal distribution on the Lander. Heat pipes are used to warm up areas of interest such as the top of the mast, to instrument operationaltemperatures.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 79There is a large variety of algorithms for each of thesefunctions, thus extensive trade-offs had to be performed.

Regarding primary guidance, Convex guidance [84] wasselected. The algorithm is based on solving a Second OrderCone Problem (SOCP) based on convex optimization andshows advantages in its ability to handle constraints,robustness, accuracy and fuel optimality. For pilotingguidance, D'Souza's guidance was chosen [85] that is based

Table 7Reference instrument list for the Lander (Landing guidance, navigation and control instruments).

Instrument Mass (kg) Power (W) (Max) Generated data

Sun sensors (2 ) 0.1 (per unit)

100 Mb (all GN&C instruments during landing, as per [88])IMUs (2 ) 4.4 (per unit) 22HDA Lidar (2 ) 3.7 (per unit) 22TRN Camera 3 10Total 19.4 54

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 638980on minimizing a performance index that depends on theweighted flight time and commanded acceleration. Advan-tages associated with it are fuel optimality, computationalefficiency and simplicity.

For navigation, the Vision Based Autonomous RelativeNavigation (VARN) algorithm is chosen [86]. The basicprinciple of VARN is the observation of features/landmarksusing only a passive sensor and the combination of thesemeasurements with the estimation of the current attitudeof the spacecraft to deduce its position, using a Kalmanfilter. It comes with advantages in regards to its ability tohandle constraints, robustness, and accuracy.

For HDA, the VB-CR-HDA algorithm was chosen [87],which analyzes the terrain topography and detects slopes,shadows and analyzes the textures using an image proces-sing algorithm using a single optical camera.

The navigation sensors used by this system during thevarious phases of landing include sun sensors, IMUs, LIDARfor HDA and a hi-resolution camera for Terrain RelativeNavigation (TRN). A list of instruments used for landingGN&C is given below. Control of the calculated trajectory andattitude is then performed using the Lander's propulsionsystem (Table 7).

3.3.3. IceMoleWithin the scope of the EnEx project, the EnEx-IceMole

design is adapted to the subsurface investigation of ter-restrial glaciers and ice shields as seen on Section 2.1(1515200 cm3, 60 kg). A first miniaturized version ofthe IceMole, called MarsMole (6630 cm3, 5 kg) is cur-rently built for tests under simulated Mars conditions. As isevident from Section 3.2.1, the cross sectional area of theIceMole version that would be deployed on Enceladus shouldbe minimized to reduce power requirements for the meltinghead. On the other hand, reducing the size of the IceMolewould require significant miniaturization of the instrumentsand electronics it must carry. A trade-off must thus beperformed between the two. As a current baseline we haveconsidered that an IceMole version with a size of1010100 cm3 is a reasonable compromise.

The configuration of the IceMole to be landed onEnceladus will be based on that of the EnEx-IceMole(Fig. 4), considering the size and component modificationsdescribed in this chapter.

3.3.3.1. Instruments. In Table 1, six methods that can beconceivably used to detect biosignatures are listed:antibody microarray/sequencing, nanopore-based devices,microscopy, flow cytometry, fatty acid markers, and massspectrometry. The instruments associated with the MCTs ofmolecular replication and cellularization should have thehighest potential for detection according to what is knownfrom terrestrial biology, as discussed in Section 1.2.3.

As also discussed in Section 1.2.3, the candidate instru-ments require significant effort for their miniaturization.According to the results of that miniaturization effort, onlya certain limited number of these instruments will be ableto fit on an ice melting probe, due to the volume con-straints stemming from the need to minimize the probe'scross-section as mentioned above. Such miniaturizationefforts are not within the scope of the EnEx project. Due tothe development, volume, and mass uncertaintiesdescribed above, an exact downselection of instrumentscannot be given beyond the candidate instruments derivedin Section 1.2.3 and listed above. For the following calcula-tions we will use data for the payload of the HADES icemelting probe [88]. Engineering requirements areexpected to be in the same order of magnitude as for theinstruments eventually selected for a future IceMole ver-sion deployed on Enceladus.3.3.3.2. Melting head, side heaters, and screw. The generalprinciple of melting and maneuvering through the ice usingmelting head, side heaters and screw is as described inSection 2.1 for the current version of the IceMole, the dif-ference being the increased power input to the heating units.3.3.3.3. Guidance, navigation and control (GN&C). Again,the same principles apply for navigation under the icesof Enceladus, as the ones for the current EnEx-IceMoleversion (IMU, IceMole and surface magnetometer, deadreckoning by measuring the screw turning).3.3.3.4. Tether. The tether will serve a dual purpose:transfer power from the nuclear reactor from the Landerto the IceMole as well as data and commands between theIceMole and the Lander. As also seen in Section 2.1, thetether must comply with strict mechanical, thermal,electrical and volume requirements. It must be able tooperate at the high voltages necessary for powertransmision through a cable of 100 m or more. It has tobe thin enough, so that the entire length of the cable canbe stowed in one or multiple possibly jetissonablecontainers in the back of the IceMole, due to the meltingchannel refreezing behind it. It must also minimizethermal losses in the cold ice, to achieve maximumpower transmission and survive the mechanical stressesit might encounter.

Fig. 13. Two-EGA transfer of the Combined Spacecraft to Saturn with lowV1 approach at Saturn. The inset illustrates the first leg of the trajectorycontaining the two EGA maneuvers, with thick line arcs denoting low-thrust arcs [90] (grid of the inset shows 1 AU from center to edge). Thefinal transfer leg after the second EGA to Saturn is shown in green. Thrustarcs for this leg take place after the final EGA and before Saturn capture(not shown). The orbits of Earth and Saturn are shown. (For interpreta-tion of the references to color in this figure legend, the reader is referredto the web version of this article.)

Fig. 14. The trajectory of the spacecraft (green) from the original captureorbit around Saturn, to a low energy Saturn orbit close to the orbit ofEnceladus (red). (For interpretation of the references to color in thisfigure legend, the reader is referred to the web version of this article.)

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389 813.3.3.5. Thermal. The IceMole must be able to maintain allits electronics within their operating temperatures. This isexpected to pose no significant problem during nominaloperations; however problems might occur during non-nominal power outages. For this contingency it isconceivable that on-board batteries can power localheaters in critical areas.Fig. 15. Reference stable orbit around Enceladus [91].3.4. Operations

3.4.1. Launch and interplanetary transferThe low-thrust nuclear electric trajectories for the

entire transfer to Enceladus (Sections 3.4.1 and 3.4.2)where modeled using the InTrance software package(Intelligent Trajectory Optimization using NeurocontrollerEvolution). InTrance combines evolutionary algorithmswith neuronal networks in order to find an optimalstrategy for a given optimization problem [89].

The mission will be launched directly on an interpla-netary trajectory with a C3 of 0 km2/s2. An Earth gravityassist (EGA) maneuver is performed one year after launch,and a second EGA 2 years later. These two EGAs, inconjunction with intermediate NEP thrust arcs will impartan outgoing velocity of 8 km/s to the spacecraft for thefinal leg of the transfer to Saturn. This final leg includes athrust arc during its final part to reduce the incomingvelocity at Saturn and enable capture. This final leg takes8.7 years. The complete interplanetary transfer takes thus14.55 years and is illustrated in Fig. 13.

Since the trajectory includes Earth gravity assists only, alaunch opportunity occurs every 1.04 years according tothe Earth Saturn synodic period.The gravity assists at Earth with an operating nuclearreactor may be problematic. In the future the flyby altitudesmust be elaborated.3.4.2. Moon tourAfter Saturn orbit insertion, the moon tour phase will

be initiated, using low thrust propulsion (Fig. 14). The aimof the moon tour will be for the Combined Spacecraft tocircularize its orbit close to that of Enceladus, and even-tually capture around the moon. This will last 2.85 years.

K. Konstantinidis et al. / Acta Astronautica 106 (2015) 6389823.4.3. Reconnaissance and Landing site selectionPolar orbits around Enceladus that would be ideal for

remote sensing of the south-polar region are not long termstable due to strong 3rd body effects from Saturn. Highlyinclined, long term stable orbits believed to be near themaximum feasible inclination are presented in [91]. Theorbit chosen for the mission has altitudes ranging from190 to 290 km, eccentricities ranging from 0.06 to 0.19,and inclinations ranging from 551 to 641. It is depicted inFig. 15. The high altitudes compared to the body radiusmakes both poles viewable (including the Tiger Stripesnear the southern pole) at near nadir pointing angles.To achieve better observation angles, if needed, the Orbitercan go to a polar orbit and remain there for no longer thanthree days, but such a maneuver is not currently consid-ered. The orbital period is about 6 h, allowing for foursouth-polar passes per day. During those passes theinstruments can sense the surface and subsurface of oneor several candidate landing sites.

In detail, assuming that an orbit around Enceladus has aperiod of approximately 6 h [91], we allocate 30 min perorbit for the Orbiter to perform pointed instrumentobservations of the area around one candidate landingspot. The data collected and the maximum power usage byeach instrument is shown in Table 9. For the remaining 5 h30 min of the orbit, the Orbiter is pointing to Earth,transmitting the obtained data. According to the 130 Kbpsdata rate given in Section 3.3.1.2, the 3.93 Gb of reconnais-sance data obtained from an overflight of a candidatelanding spot can be transmitted over the next 2 orbitalperiods, or 11 h 50 min with a margin of 30%. Eclipsing ofthe spacecraft by Enceladus and Saturn is not expected togreatly affect the link duration to Earth, due to therelatively high altitude of the used orbit (close to 1 RENC).

The process of remote sensing a particular candidatelanding spot, and then forwarding the data to Earth overtwo orbital periods is repeated for several of the initiallydesignated plume sources as shown in Fig. 3, completingthe reconnaissance after several days. In parallel to the thisprocess, a team of experts processes the downlinked datain order to choose the optimal landing site based onTable 8Reference instrument data for the IceMole (Scientific in situ instruments).

Instrument Mass(kg)

Power (W)(Max)

Data volume(Mb)

IceMole scientificinstrument suite 7 38 147

Table 9Data volume generated, and maximum power usage during a pass over the Sou

Instrument Data volume(Gb)

Power (W)(Max)

Comme

Reconnaissance Camera (RC) 2.4 38 Data raRadar/Ice Penetrating Radar(IPR)

1 28 Source

Thermal Mapper (TM)0.53 14 Source

duratioTotal 3.93 80scientific interest and landing safety characteristics (seeSection 1.4). Once this is completed for the initiallydesignated candidate landing spots, the procedure can berepeated for newly discovered areas of interest or repeatobservations as needed. This operation phase is thusexpected to last up to a few months.

3.4.4. LandingThe driving requirements for landing accuracy are the

width of the canyon floor, at around 100 m as discussed inSection 1.3, and the desire to minimize the length of the tetherconnecting the IceMole to the Lander for power and commu-nications purposes (Section 3.3.3.4). Landing too close to theplume source should also be avoided, to minimize planetaryprotection related concerns (see Section 4). Taking theseconsiderations into account, the target landing area shouldbe a circle with a radius smaller than 50m. Such a landingaccuracy is categorized as pinpoint landing (see e.g. [74]).

Once the landing site selection has been performed, thelanding sequence can be initiated: the Lander separatesfrom the Orbiter and performs a deorbit maneuver totarget the South-Polar Terrain. The GN&C landing subsys-tem described in Section 3.3.2.9 is designed to achieve apinpoint landing. The landing operational sequence wasbased on that of the ESA Lunar Lander mission [93]. Thelanding sequences of other icy moon landing missionswith similar accuracy requirements were taken intoaccount [71,74].

After separation from the Orbiter, the Lander performs aburn to simultaneously lower its periapsis to 5 km abovethe south pole and raise the inclination of its orbit from 601to 901. The Lander then coasts to periapsis, where it starts asecond burn to dump most of its excess velocity. Once theLander has achieved a sufficiently reduced velocity theengines begin to throttle down and the targeted descentphase begins. During this phase the HDA is active. If thedesignated landing spot proves unsuitable, retargeting cantake place and a new landing spot within the Lander's reachselected. Once the Lander is above the landing site with nohorizontal velocity and a vertical attitude at an altitude of1020 m, it shuts it engines and slowly free-falls to a softlanding on the surface.

3.4.5. IceMole operationsAfter Lander touchdown a first 3D subglacial map is

generated via sensing the relevant area of operation usingthe Ground Penetrating Radar to identify the position ofpossible hazards and targets. The IceMole is deployed andstarts melting. For the first stage of melting, pressure mayth-Polar Terrain during the Reconnaissance phase.

nt

te from Table 4 for 15 min. operational durationfor data volume: [92]

for data volume: data rate from Tables 4 for 15 min. operationaln

Table 10Data generation per Lander instrument and the IceMole during a six hourperiod.