IAC-05-A3.2.A.09

EXPLORING TRITON WITH MULTIPLE LANDERS

by Tibor S. BalintJet Propulsion Laboratory

California Institute of Technology4800 Oak Grove Drive, M/S 301-170U

Pasadena, CA 91109-8099e-mail: tibor.balint@jpl.nasa.gov

ABSTRACT

In our pathway for Outer Planetary Exploration several mission concepts were considered, based onthe proposed JIMO mission architecture. This paper describes a JIMO follow-on mission conceptto Neptunes largest moon. Triton is a target of interest for outer solar system studies. It has ahighly inclined retrograde orbit, suggesting that it may have been a Kuiper Belt object capturedby Neptune. Given this assumption its composition, which may include organic materials, wouldbe of significant scientific interest. The present concept considers a surface mission architecturewith two landers, each powered by a standard multi-mission radioisotope thermoelectric generator(MMRTG). The landers would operate on the surface for several years providing science data, thusexpanding our understanding of the environment, the dynamic surface and atmospheric processes,and some of the seasonal variations. A JIMO class orbiter would provide telecommunication linkbetween the landers and Earth, and would be instrumented to observe both Triton and Neptune.In this paper all key aspects of the mission architecture are addressed, including the science instru-ments, the main subsystems, trade options for the power system and a conceptual design for thelanders.

INTRODUCTION

The Vision for Space Exploration [1] identi-fies three major exploration pathways, targetingMars, the Moon and the Outer Planets. Withinthis roadmap the first planned science missionto the Outer Planets was the Jupiter Icy MoonsOrbiter (JIMO) mission. Potential JIMO follow-on missions were envisioned to target other outerplanetary destinations, such as Saturn, Neptuneor Pluto. This paper describes a landed missionconcept to Neptune’s largest moon, Triton, withthe aim of expanding our incomplete knowledgeof the Neptunian system, which is based on theAugust 25, 1989 flyby of the Voyager 2 space-craft and on more recent Earth and space-basedobservations.Since the completion of this study, the JIMOmission has been deferred. Although this Tri-ton lander concept baselined a JIMO follow-onarchitecture, alternatives were also addressed.Thus, this generic lander concept is not depen-

dent on JIMO follow-on architectures and couldbe considered as a “black box” concept for otherNeptune/Triton exploration missions.

SCIENCE AND MISSION GOALS

Triton is a target of great interest for outer solarsystem studies. Thus, potential science objec-tives for a Triton lander mission would includea more complete characterization of the com-position and circulation of the atmosphere; in-vestigation of the physical processes responsiblefor plume formation; surface composition mea-surements; and geophysical monitoring. In par-ticular, seismological measurements could po-tentially refine our knowledge of the physicsof plume eruptions, and could detect Triton-quakes, if such are occurring at the present time.A pair of landers, one located in the summerhemisphere and the other in the winter hemi-sphere, could collect complementary information

1

Figure 1: Triton images: (a) South Pole; (b) Tenuous Clouds; (c) Icy-slurry filled crater.

on atmospheric processes and on the interiorstructure. It would be very desirable to landatop one of the dark streaks (Figure 1(a)), inorder to perform in-situ compositional analyses.This information would help constrain models ofthe physics of plume eruptions (Figure 1(b)) onthis exceedingly cold, icy body.The complement of instruments for such alanded mission could include a sophisticatedweather station (pressure, wind, and tempera-ture measurements), along with an imaging sys-tem and a micro-seismometer system. Remoteor contact instruments for determining the com-positions of surface materials would also havehigh priority. Last but not least, a separate in-strument for determining the atmospheric com-position (a mass spectrometer of some type)should be included.Although not described here in detail, the ac-companying orbiter should likewise carry instru-mentation to measure the magnetic field (if any)and to determine the internal structure throughgravity measurements. (The latter would be aradio science investigation employing the com-munications signal from the spacecraft to theEarth). Imaging, remote sensing of surface com-position, and fields and particles instruments ap-propriate for Neptune’s magnetospheric environ-ment should also be considered. Remote mea-surements of surface composition may be diffi-cult due to the extremely low surface tempera-ture; active illumination of the surface might berequired. Further details on Triton can be foundin [2].The mission goals for this Triton lander mis-

sion would include a successful landing of onebut preferably two spacecraft on the surfacefollowed by a long duration surface science op-eration (measured in years) to characterize theenvironment and to extend our knowledge baseon the Neptunian system. Measurements wouldbe taken to identify mineral composition in thevicinity of the landing site. Seismic activity andgeyser eruptions would be monitored as well.Visual observations would be taken and localmeteorological conditions could be monitoredover a long duty cycle in order to characterizeseasonal changes through a small portion of aNeptunian year.

MISSION ARCHITECTUREOVERVIEW AND ASSUMPTIONS

The primary architecture assumes a JIMOfollow-on configuration utilizing a low thrustnuclear electric propulsion (NEP) system, withthe mission referred to as the Neptune Icy MoonOrbiter or NIMO. The NIMO spacecraft, just asits predecessor, would require a heavy launch ca-pability, which does not exists at this time. Thevarious launch options could include a numberof Delta IV-H LVs and in orbit assembly or asingle heavy LV with a Saturn LV capability.Although this study will focus on this primaryconfiguration, a second high thrust trajectorybased architecture will be also mentioned forcomparison purposes. The various mission ar-chitecture options are summarized in Table 1.

2

Option CommentsBaseline (NIMO) 300 kWe fission reactor; Low thrust through all stages;

Payload: ∼3000 kg; Trip time: ∼15 years to Neptune(+ 3 years orbit transfer to Triton)

Chemical / SEP C3=12.1 km2/s2; SEP: 50 kWe;(Low C3 / High mass) Payload: ∼1940 kg; Trip time: ∼12.5 yearsChemical / SEP C3=18.4 km2/s2; SEP: 30 kWe;(High C3 / Low mass) Payload: ∼790 kg; Trip time: ∼10.25 yearsJIMO class launcher (a) multiple launches plus in orbit assemblyfor the baseline option (b) Saturn class launch vehicle

Table 1: Mission architecture options for a Triton lander mission.

Baseline Architecture

It is assumed that a 300 kWe nuclear reactorwould power NIMO, built upon the 100 kWeJIMO heritage. NEP enables the highest massdelivery, but results in the longest transfer timewhen compared against high thrust trajectories.Venus and Jupiter gravity assists could furtherreduce the trip time to Neptune. A NEP en-abled mission to Neptune would take about 15years, which includes spiraling out of Earth andspiraling in to Neptune. It would take an ad-ditional 3 years to reach orbit around Triton.JIMO’s currently proposed payload allocation is1500 kg. Since the NIMO spacecraft would onlypass Jupiter during a gravity assist flyby, thehigh radiation environment of the Jovian sys-tem would not have a significant effect on it.Thus the mass allocated on JIMO for shield-ing could, in part, be reallocated as payload onNIMO. (Note that mass differences will in partbe a function of the size of the power system, andmore importantly will reflect differences in thepropellant needs to account for traveling about6 times farther from the Sun to Neptune thanJIMO’s journey to Jupiter.) Increasing the pay-load envelope to as high as 3000 kg could beassumed for the present mission, effectively sup-porting two landers on Triton and still allowingfor additional science instruments on the orbiterfor remote sensing / mapping of Triton and forobserving Neptune from Triton’s orbit. Since thepresent mission concept focuses on the landers,it is assumed that NIMO is already in a 1500km circular orbit around Triton with a suitable

payload allocation for two landers. Therefore,details of the NEP enabled low thrust trajectoryfrom launch to arrival are not addressed due tothe limited scope of this study.

According to the primary mission configuration,NIMO would achieve a 1500 km circular orbitaround Triton and spend the first weeks map-ping the surface. The returned data would en-able the science team on Earth to select suit-able landing locations for the two landers fora “stop-and-drop” type landing. Each landerwould de-orbit to the surface using the lander’sown propulsion system. An initial 137 m/s smallde-orbit burn would lower the lander’s periapsisto about 20 km, where a large 1200 m/s burnwould remove the horizontal velocity. Soft land-ing would re-quire a small 195 m/s throttledburn from the bi-propellant system, assisted fur-ther by a sky-crane, which would be based on2009 Mars Science Laboratory (MSL) concept.Lowering the landing platform from a skycranecould help with minimizing surface contamina-tion from the impinging exhaust of the thrusters.Once the lander was released, the skycrane plat-form would disengage and crash land at a safedistance from the payload base. Soft landing wasselected for two reasons. First, when landing onan airless body the option for an aeroshell andparachutes is not feasible, thus all or at leastmost of the velocity must be removed throughpropulsive means. (From the aspects of descentand landing, Triton, with its very thin atmo-sphere, can be considered an airless body.) Fur-ther to this, the fuel mass saving from cutting offthe engine at a few kilometer altitude, free falling

3

Figure 2: Triton lander with a skycrane platform

and landing with airbags is significantly less thanthe additional mass required for an second land-ing system, such as airbags for hard landing orcrushable mate-rials for rough landing. There-fore, as demonstrated by Balint [3] adding asecond landing system, such as airbags similarto the landing configuration of the Mars Explo-ration Rovers [4], or crushable materials such asproposed for the Mars Net Landers [5] woulddecrease the landed payload mass. Second, thepower source proposed for this mission, a Multi-Mission Radioisotope Thermoelectric Generator(MMRTG), is designed for a maximum acceler-ation load tolerance of about 40g. Therefore, onTriton soft landing presents the only viable masseffective landing configuration with the givenpower source. A conceptual design of the Tritonlander with the skycrane, thrusters and propul-sion system is shown in Figure 2.As discussed above the first lander would touchdown on the illuminated side of Triton, such asthe South Pole shown in Figure 1(a). The land-ing location could be either inside or outside ofthe dark streaks. Another potential landing lo-cation could be inside the remains of an ancientimpact crater that is thought to be filled withice, probably formed by eruptions of water orwater-ammonia slurry (Figure 1(c)).The second lander would touch down on the op-

posite side of the moon, which would allow forstudying atmospheric processes such as poten-tial migration of atmospheric constituents fromthe illuminated side to the dark side of Triton.

The environment on the surface of Triton isharsh. Therefore, the lander concepts are de-signed with the philosophy of simplicity andreliability in mind to meet mission lifetime re-quirements using an appropriate combination ofhigh-reliability components and dual-string de-sign. The very long 18 years cruise phase com-bined with the 38K surface temperature makesmobility with a rover or even with a roboticarm undesirable. Thermal cycling or freezingcould cause an early end to a long awaited mis-sion. Upon arrival to the surface, the landerwould initiate the investigations of science tar-gets. First it would deploy its only mobile com-ponent, the mast, on which the panoramic cam-era (Pan Cam) and remote sensing instruments(Raman Spectrometer / Laser Induced Break-down Spectrometer or LIBS) are located. Goodcontact with the surface would allow for seis-mic measurements. A meteorology sensor suitewould monitor the temperature and pressurechanges in the atmosphere, while a Gas Chro-matograph / Mass Spectrometer (GCMS) wouldmake compositional measurements. Details onthe instruments and their operations are given

4

in the relevant sections below. To minimize risk,the lander design would use as much design andflight mission heritage as possible from previousouter planets orbiters and landers.

Additional Architectures

In comparison with the baseline concept usingNEP, a chemical propulsion system combinedwith a 30 kWe solar electric propulsion (SEP)system, could reduce the trip time from ∼15years to ∼10.25 years. Due to the distance be-tween Earth and Neptune, the trajectory wouldrequire a high C3. Since Triton’s gravity is verysmall, Neptune’s gravity field would be used tocapture the spacecraft. From there the orbitwould be changed to an orbit around Triton,still with the propulsion system of the motherspacecraft. Such a mission is expected to uti-lize a significantly sized propulsion system forthe orbiter/mother spacecraft. For this baselinetrajectory the spacecraft would be launched ona C3 of 18.4 km2/s2, followed by a Venus anda Jupiter Gravity Assist (VJGA). The inertialentry velocity for a Neptune aerocapture, in therange of 28-30 km/s, would offer the best combi-nation of highest delivered mass to a Neptunianorbit with the lowest entry heating. Althoughthis second option would cut trip time by abouta third and would be achievable with availableDelta IV-H or Atlas V launch vehicles, the de-liverable total mass to Neptune’s orbit would beonly around 790 kg [6] [7]. This mass is notsufficient for a Neptune-to-Triton orbit trans-fer while still supporting an orbiter and a lan-der at Triton. With the reduction of C3 ve-locity from 18.4 km2/s2 to 12.1 km2/s2 (withVinf=12.8 km/s) and scaling up to a 50 kWeSEP stage (with 2800 kg wet mass), about 3330kg could be inserted into Neptune’s orbit on a2017 launch opportunity. This is based on aDelta IV-H launch vehicle with a 7250 kg pay-load mass inserted to a C3 of 12.1 km2/s2; a12.5-year VJGA trajectory; and an advanced ae-rocapture vehicle (<1250 kg). (Note that for thisC3 the maximum payload on a Delta IV 4050-Hlaunch vehicle is 7510 kg [8].) Following an orbittransfer to Triton the total mass of the space-craft would be around 1940 kg. Assuming anapproximately 900 kg lander the spacecraft in

orbit could only support one lander and an ap-proximately 1000 kg orbiter. This configurationwould not fully satisfy the science goals of themission. Although this option was not selected,the velocity and propellant mass calculations forthe lander shown in Table 2 would be the sameas for the primary mission configuration.Note that this second mission architecture wouldbe more power limited. Due to an average 30AU distance from Triton to Earth, telecommuni-cation would present a significant challenge, re-quiring power in the multi hundred-watt range.For example 3 MMRTGs could provide about250 We power to the orbiter after 18 years. Insummary, this second architecture would require4 MMRTGs, between the orbiter and a singlelander.A summary of the various mission architecturesis provided in Table 1. For the chemical /SEP options the SEP stage would be ejected atJupiter and the S/C would utilize aerocapture atNeptune. The baseline option would require oneof the launcher options, while the chemical/SEPoptions would use a single Delta IV-H launchvehicle.

SCIENCE INSTRUMENTS

The lander concept shown in Figure 2 was de-signed to fulfill the key science objectives ofthe mission. Potential instruments on the lan-der can be broken down to three categories,such as remote sensing, contact and analyticalsuites. Remote sensing instruments are locatedon the mast and include a panoramic camera(Pan Cam), and sensors for a Raman spectrom-eter and a Laser Induced Breakdown Spectrom-eter (LIBS). Contact instruments are the seis-mometer and, to a certain extent, the meteorol-ogy station. The Gas Chromatograph & MassSpectrometer (GC/MS) is an analytical instru-ment. All of these instruments must have suf-ficient sensitivity to measure the relevant envi-ronmental conditions. After describing the in-struments shown in the Triton lander conceptdrawing (Figure 3), additional potential instru-ments will be considered. These might be placedeither on the lander or NIMO on the orbiter.The imaging system, a MER derivativepanoramic camera, is located at the top of the

5

Item Delta V Required Total(m/s) Mass (kg) Mass (kg)

Initial payload mass at launch to C3 ⊗ ⊗ 7250Mass to place into Neptune orbit ⊗ 3460 3790SEP module (wet mass) n/a 2800 ⊗(ejected before aerocapture) ⊗Cruise propellant 25 70 ⊗Neptune aerocapture aeroshell / TPS n/a 400 ⊗(discarded after aeropassage) ⊗Aerocapture control 30 40 ⊗Neptune Perineps Raise to 4000 km 110 120 ⊗(including 3% gravity loss) ⊗Neptune Aponeps Correction 40 30 ⊗(358,000 km Neptune orbit) ⊗From Neptune orbit to Triton orbitInsert to 1500 km circular Triton orbit 2800 1850 1940Orbiter mass at lander-orbiter n/a 1000 940separation & mass available for landerFrom Triton Orbit to surface (propellant only to land 300 kg payload, includingthe propulsion system dry mass for tanks, thrusters etc., and 30% contingency)Lander de-orbit burn 137 37 ⊗(to a 1500 km by 20 km orbit) ⊗Triton Pericenter Burn 1200 241 ⊗(including 2% gravity loss) ⊗Soft landing (incl. 10% gravity loss) 195 28 ⊗Attitude Control Allotment (total) n/a 15 ⊗Propulsion system dry mass n/a 97 ⊗

Table 2: Mass allocation and Delta V requirements for a Chemical / SEP system. [9]

Figure 3: Triton Lander Instrumentation

6

mast. The high-resolution stereoscopic cameraprovides needed context and aids in character-izing the geomorphology of the surface throughthe generation of terrain maps, slope maps andranging. It can generate 360◦ panoramas andmulti-spectral images of the surface, which helpsto characterize the nature of the materials sam-pled with other instruments. Thus in effect thePan Cam would work in conjunction with theRaman spectrometer and with the LIBS. ThePan Cam camera is at TRL9.The combined Raman spectrometer and LIBSsystem could measure elemental abundance andmineralogy of surface materials. By activelystimulating the target these instruments avoidthe negative consequences of the low surfacetemperature that reduce the capabilities of TESand other IR-sensitive instruments. The Ramanspectrometer fulfills the astrobiology driven sci-ence goals by performing mineral characteriza-tion and assisting in the detection of water, or-ganic and inorganic forms of carbon. It identi-fies many major, minor and trace minerals andtheir relative proportions (i.e., Mg/Fe ratios),and carbon ratios. Sharp Raman spectral fea-tures and statistical point counting help identifyminerals in complex mixtures and morphologies.LIBS would use a higher energy excitation ofthe surface than Raman, consequently ablatingthe studied surface. Compositional informationwould be drawn from spectral analysis of theresulting plasma. The instrument is based onthe Mars Microbeam Raman Spectrometer, cur-rently at TRL4. LIBS is proposed for upcomingMars missions and is at TRL5.The lander would be equipped with a two-component seismometer measuring both highand low frequencies. The 2-axis very broadbandseismometer would capture tidal and long pe-riod motions up to 10 Hz. The 3-axis short pe-riod micro-seismometer would measure high fre-quency movements from 1 Hz to 50 Hz. The twosets of seismometers would achieve the highestsensitivity in an ultra broad band from 5×10-5Hz to 50 Hz. In addition, a partial redundancywould be achieved due to their significant over-lap in frequency band. Triton’s geysers produceplumes rising several kilometers in height. Likegeysers on Earth, these must produce seismicwaves within crustal materials, which could be

measured with this seismometer. Detection andanalysis of seismic energy can provide informa-tion on the eruptive processes (their energy, fre-quency, time evolution) and on the properties ofcrustal materials traversed by the waves. Thisproposed seismometer is currently at TRL4.Triton is one of two satellites in the solar systemwith an appreciable atmosphere, beside Saturn’smoon Titan. Every aspect of this frigid atmo-sphere is of scientific interest: its composition,its circulation, its exchange processes with thesurface, its evolution with time. A pair of so-phisticated weather stations situated in oppositehemispheres could yield a very significant sci-ence return. The very cold environment and thinatmosphere requires significant modifications tothe sensitivity of existing weather monitoringequipment. Such instruments are at TRL5.The gas chromatograph & mass spectrometer(GC/MS) measures isotopic gas ratios of traceatmospheric components. If a sampling mech-anism were implemented, the GC/MS could beused to identify the presence of organics as wellas mass spectra and isotopic ratios of evolvedgas constituents from rock and soil samples. Theinstrument is proposed for upcoming Mars mis-sions and is currently at TRL5/TRL6.Although not included in this concept, addi-tional instruments on the lander and orbitercould also be considered. For example on thelander a small sampling mechanism in the formof a robotic arm with a scoop could be used toposition the contact instruments (Raman, LIBS)closer to the target objects. If a sample ac-quisition (scoop) is included, then the collectedsample could be analyzed by a Thermal andEvolved Gas Analyzer (TEGA). TEGA offers amore complete characterization of the volatilecomponent of surface materials than is possi-ble with Raman and LIBS. However, TEGA isheavier and more complex than a simple ovento heat the samples and analyze them by theGC/MS. Sample handling introduces additionalcomplexities especially in a cold environment asTriton, hence this is not included in the presentconcept. Beside LIBS, a heat lamp or conduct-ing fins could warm the surface near the lander.Heating or thawing the surface could potentiallyinitiate small geyser-like eruptions after creatinga sub-surface greenhouse effect.

7

On the orbiter, an Ion and Neutral Mass Spec-trometer (INMS) could directly sample the ten-uous atmosphere surrounding Triton. It couldconfirm the presence of the major gases andcould detect others (not yet known to bepresent) in the atmosphere. The orbiter shouldcarry additional atmospheric remote sensing in-struments to fully characterize the compositionof the atmosphere, such as (LIDAR) radar;ground penetrating radar; cameras; and a Ther-mal Emission Spectrometer (TES). (Note thatthermal emissions at 38K are very low, com-pared to the ∼150K on Mars. Therefore, TES atTriton may not have the sensitivity to performmeaningful measurements.)

POWER SOURCE SELECTION

This section discusses power source trade op-tions, radioisotope power system characteristicsused for the baseline configuration and describesalternative RPS options.

Power Source Trade Study

Insolation decreases with one over distancesquared from the Sun. In fact, at Neptune (30AU from the Sun) solar radiation is only about0.1% of that at Earth. It has been shown in[10] that beyond ∼3-4 AU solar power genera-tion with current technology is less mass efficientthan power generation with RPSs. Low Inten-sity Low Temperature, or LILT, solar panel tech-nology – which is planned for the Juno spacecraftto Jupiter – could potentially work at 5 AU, butbeyond that it could be mass prohibitive andlikely not feasible. Consequently, missions toJupiter and beyond (such as to the Neptuniansystem) require a different kind of power source,independent from the Sun. Batteries may sup-port limited duration mission operations, how-ever, longer missions require nuclear fission orradioisotope decay-based power systems. Be-side an MMRTG, a lander mission to the Nep-tunian system could utilize a Stirling Radioiso-tope Generator (SRG) or a small fission reac-tor. SRG-110 is a 110 W (BOL) Stirling systemunder development with a TRL9 target date of2009. It is based on two GPHS modules andtwo pistons positioned head-to-head, resulting

in a significantly lower Pu238 fuel requirementthan the 8 module based MMRTG. With fewerGPHS modules the thermal output of the SRGwould also be reduced by 75%, to about 500Wt (BOL). However, with an MMRTG morewaste heat would be available to heat the space-craft, which is more desirable on the presentmission. It should be also noted that both theMMRTG and the SRG have the same 14 yearslife-time requirement [11] [12], while the pro-posed Triton mission would have a mission lifeover 18 years. Outer planets missions should ad-dress and resolve mission lifetime issues in thiscontext. Small fission reactors, for example aHOMER type reactor, could generate ∼3 kWeof power. A small 3 kWe surface reactor withStirling power converter could weight about 775kg [13]. This power and mass configuration isbeyond the power requirements and mass lim-its of a Triton lander mission. Therefore, fissionreactors are not considered as power source al-ternatives. (Note that, a fission reactor can re-main inactive until the beginning of the surfaceoperation. During the inactive “cold” phase thesystem produces negligible radiation and is notaffected by the long cruise phase.)Although not the focus of this study and there-fore not discussed in detail, the NIMO orbiterwould perform remote sensing measurements tocharacterize Triton and Neptune. On NIMO, theonboard 300 We nuclear reactor would power thescience instruments. Other subsystems such astelecom and command and data handling wouldalso be supported. Thus, the baseline missionarchitecture would rely on a nuclear reactor onthe NIMO orbiter and two MMRTGs on the twolanders.

RPS Characteristics

Each lander would use a single MMRTG. It isrequired that the power system should operatecontinuously during the entire mission, whichincludes ∼15 years of cruise phase to Neptuneand about 3 additional years to reach Triton’sorbit. During these 18 years the power gener-ated by the MMRTG would degrade by about1.6% per year. Half of it is due to natural de-cay of the plutonium fuel and the other halfis to degradation of the thermoelectrics. Thus

8

at the beginning of the science mission, definedby landing on Triton, the generated electricaland thermal power would be only ∼82.6 Weand ∼1735 Wt, respectively. During surface op-eration, the power would continue to decreaseat a rate of about 1.2 We/year. Soft landingon Triton would impart acceleration loads onthe MMRTG within design limits. Waste heatwould be utilized through radiation from theMMRTG to a hot plate of the warm electronicbox (WEB), and through conduction along theMMRTG fins and thermal straps from the powersource to the WEB. Consequently, the presentdesign would benefit from two of the advantagesof an RPS, namely continuous electrical powergeneration and utilization of its excess heat.

Alternative RPS Power System

Small radioisotope power systems with TE con-version, in a modular configuration, could beconsidered for a Triton lander mission. However,the minimum number of small-RPSs required tosupport a Triton lander mission would be closeto the number of GPHS modules in an MM-RTG.. The mission could also consider SRGs,providing the same electric power output as theMMRTG. However, an MMRTG would generatefour times more heat than an SRG,, providingan advantage at the cold Triton environmentthrough waste heat utilization. Therefore, anMMRTG is considered the best choice for a Tri-ton lander mission in order to reduce complexityand potentially lower cost through the use of asingle system.

OPERATIONS AND ENVIRONMENTS

This sections discusses operational issues, suchas power requirements, data collection and com-munications, and thermal and radiation envi-ronments.

Power Issues

Each lander would utilize an MMRTG for elec-trical power. In addition, a 25 Ahr Li-Ion bat-tery would provide backup during high poweroperating modes, such as during telecom oppor-

tunities. Power calculations for a 3-day repeat-able mission scenario with an 82 We continuouspower source demonstrated that the present hy-brid system would provide sufficient power tothe lander and would keep the secondary bat-tery power positive. This would permit a repeat-able cycle lander operations through the wholemission lifetime. In the power analysis threeoperational modes were considered. In highpower mode all science instruments and sup-port subsystems would work simultaneously. Inlow power mode, some of the instruments wouldnot be operated, such as the Raman spectrome-ter, LIBS and the GC/MS. In telecom mode theUHF transmitter would operate in conjunctionwith the power, electronics and thermal subsys-tems. Science instruments designated as “Al-ways on// would be also operational through-out all three modes, as shown in Table 3. Theelectronics subsystem would provide permanentsupport to the instruments, by processing thecollected information, storing it and sending itto NIMO through the telecom system (10.4 We).The power subsystem would use 15.4 We topower the peripheral control unit (PCU), thepower distribution unit (PDU), the battery con-trol, the universal switch and the shunt limiter.Thermal heaters would also be used continuously(5.9 We) to keep the Pan Cam and the warmelectronics box above survival temperature. Thetelecom system would use 52 We of power butonly during the telecom opportunities.The power analysis confirmed that an MMRTGenabled lander to Triton would be feasible.

Data and Communications Issues

The command and data handling system, as-sumed to be a dual string Harris RH3000 elec-tronics unit with radiation tolerance over 100kRad, is sized by the data collected from the sci-ence instruments and communicated to NIMOduring the telecom opportunities. The high-est data volume would be generated by theseismometer, which would collect up to 16.6kbits/sec and would operate continuously. Thecollected data would include both high and lowfrequency measurements. This data volumecould be significantly reduced by data compres-sion and by stand-by monitoring of the activi-

9

Instrument/Subsystem Power Duty cyclePanoramic Camera 5.6 We 1 hour on demandRaman Spectrometer 23.4 We 2-3 hoursLIBS 23.4 We 2-3 hoursSeismometer 3.5 We Always onMeteorology sensors 1.4 We Always onGC/MS 11 We 8 hours totalHarris Electronics 10.4 We Always onPower subsystem 15.4 We Always onTelecom 52 We 1 hr / day (max)Thermal heaters 5.9 We Always on

Table 3: Power requirements (with 30% contingency).

Instrument/Subsystem Data ratePanoramic Camera 50.4 Mbits/frame raw. Average compression

∼3:1; 12 frames required for a full panoramaRaman Spectrometer 17 kbit / spectra (raw), 50 Raman spectra per hourLIBS Similar to RamanSeismometer 16.6 kbits/secMeteorology sensors Few kbits / hourGC/MS ∼10 kbits per mass spectra (MS);

∼200 kbits per evolved gas sample (GC)

Table 4: Data rate.

Systems CBE Mass CBE Mass +(kg) 30% cont. (kg)

Propulsion System 321.5 418Structures & Mechanisms 135.0 175.5Power System 77.1 100.2Thermal System 7.9 10.3Avionics 2.2 2.9Telecom 8.3 10.8Science Instruments 12.4 16.2Total 564.5 733.8

Table 5: Triton Lander mass Breakdown.

10

ties. For time periods without seismic activitythe data handling system would simply discardthe data. In case of an activity, a memory loopwould retain the immediate time period priorto the event and record throughout the activ-ity. The Pan Cam would take 12 frames for afull panorama, where each frame would use 50.4Mbits of raw data. This data could be com-pressed at a ratio of 3 to 1. Following the initial360◦ panorama taking, the Pan Cam would beon standby mode until a seismic activity, whichcould allow to capture geyser events in the vicin-ity of the lander. The Raman spectrometer andLIBS would each generate less than 1 Mbits ofdata per measurement. The GC/MS would per-form only a limited number of measurements;therefore, the data obtained by these instru-ments would not have an impact on the C&DHsystem. (see Table 4)The lander operation would include two modes,based on operation time frames. The initialmode, following the landing, would include a fullset of measurements. The second mode wouldswitch most of the instruments into standbymode, keeping only the seismometer and meteo-rology sensors operational. This second modewould generate only a small amount of data,which would reduce data transfer from the sur-face to NIMO and back to Earth. For the secondmode only a reduced staff would be required tooperate the orbiter and landers, and to analyzethe data.The distance of Neptune from the Sun is 30AU, therefore, direct to Earth (DTE) commu-nication from the landers is not likely. Becauseof the limited power availability from the MM-RTG and potential visibility issues, the datato Earth would nominally be relayed throughNIMO. Each lander would utilize redundantElectraLite UHF radios with a 5 W transmitterto communicate with NIMO. This telecom sys-tem would support a data rate up to 500 kbps.(The UHF system could transfer data at ratesbetween 1 kbps and 2048 kbps, while the receivedata rate would correspond to a range between 1kbps and 8 kbps.) Based on the telecom oppor-tunities for an assumed 1500 km orbit the dailydata volume could be over 200 Mbits, transmit-ted through the lander’s monopole antenna. TheElectraLite UHF radio would be placed inside

the warm electronics box, while the base plateof the UHF monopole antenna would be posi-tioned above the MMRTG, utilizing its wasteheat through conduction and radiation to pre-vent the antenna from freezing (see Figure 3).

Thermal Issues

Thermal design of the landers requires maintain-ing them at an appropriate operating tempera-ture during all phases of the missions. It wouldbe desirable also to reduce and potentially elim-inate the number of moving components on thelanders in order to minimize the potential forthermal-mechanical failures. For this, the ther-mal environment could be sustained by utilizingwaste heat from the MMRTG.Thermal control for the Triton landers would beaccomplished by a combination of passive andactive components. Both landers would havethe same thermal design. The Pan Cam camerawould require a 2.5 W resistance heater. Tomaintain the electronics box warm, two possibleoptions could be considered, either by coveringeach of the exposed surfaces with a 1/16 inchlayer of Aerogel (a high performance, lightweightinsulator) or by covering them with MLI in addi-tion to two 2 W resistance heaters to compensatefor a less efficient thermal blanket design. AnMMRTG would generate about 1735 W of ther-mal heat, a portion of which would be dissipatedradiatively through the fins. The remaining heat(∼520 Wt assuming a 70% efficient fin) wouldneed to be conducted to the warm electronicsbox (WEB) using a high performance thermalstrap (such as a K1000 heat strap). A doublerplate could also be used to help remove a por-tion of the heat and to minimize the hot spotslocated along the mounting plate. It was foundthat waste heat utilization with the present ther-mal management strategy would be sufficient tokeep the warm electronics box continuously withits operating temperature range, while the MM-RTG would provide enough power to allow forresistance heating of components outside of theWEB (e.g., Pan Cam).

Radiation Issues

During the total mission lifetime the NIMO

11

spacecraft and the landers would be exposed tovarious radiation sources. These are the VanAllen radiation belts near Earth, cosmic radi-ation through the cruise phase, Jupiter’s radi-ation during the flyby, and radiation from theNIMO reactor and from the MMRTGs. A cal-culation was performed comparing the radiationenvironment for this mission against the prelimi-nary calculations for the JIMO mission. The cal-culations assumed 100 mil of aluminum shield-ing to protect NIMO and its landers. During thespiraling out phase from an 1000 km Earth orbitthe Van Allen radiation belts would expose bothNIMO and the landers to about 100 kRad of to-tal ionizing dose (TID) radiation. The Jupiterflyby could add about 40 to 60 kRad, based onthe flyby distance (the above values assumed 10and 6 Jupiter radii at 0◦ inclination). The NIMOreactor would add ∼40 kRad TID. Radiationfrom the MMRTG would be dependent on thedistance from the radiation source and was ex-trapolated from calculations performed for theMars Science Laboratory mission [14]. The re-sulting TID for an 18-year cruise phase is es-timated to be below ∼250 kRad. The naturalradiation environment at Neptune results in anadditional 0.003 kRad/year, which is negligible.It is evident that such a mission does not re-quire the amount of shielding seen on JIMO.Since the shield mass per unit area for 100 milsof Al is 0.686 gm/cm2, the mass savings due toshielding could be significant, accommodatingan equally larger payload. The currently pro-jected JIMO payload allocation envelope is 1500kg. The JIMO follow-on spacecraft assumed forthis Triton lander mission could accommodate apayload in the 2000 to 3000 kg range.

MASS ALLOCATION

It is assumed that NIMO could deliver up to3000 kg of payload mass to Triton’s orbit. Thepayload allocation on the NIMO orbiter is higherthan that of JIMO due to two factors. First, the300 kWe NIMO reactor would allow for largerelectric propulsion thrusters, thus increasing thedeliverable mass to the Neptunian system. Sec-ond, Neptune’s radiation environment is benigncompared to that of Jupiter. Thus, the shielding

requirement is much lower, resulting in a masssavings, which can be re-assigned to the payloadfraction. The larger mass in effect allows forup to two landers in addition to science instru-ments allocated on the orbiter for remote sensingand data relay. The mass breakdown is shownin Table 5. The mass of the propulsion sys-tem accounts for about 57% of the lander massat the time it detaches from the orbiter. Thisincludes the propellant wet mass (bi-propellantand pressurant) and the propulsion system drymass (thrusters, tanks, valves). The structuresand mechanisms, about 24% of the total mass,account for the lander base plate, WEB housing,mast, skycrane and miscellaneous items such ascabling. The power system is almost 14% ofthe total mass, and includes the MMRTG powersource, the batteries and other components suchas PCU, PDU, shunt limiter and battery controlboards. The thermal, avionics and telecom sys-tems account for less than 3.3% of the mass. Fi-nally, the science instruments utilize about 2.2%of the total lander mass. With 30% mass mar-gin, required by design principles for conceptualdesigns, the Triton lander mass at the time itdetaches from NIMO would be about 733.8 kg,while two landers would be less than 1500 kg,leaving half of NIMO’s assumed payload for or-bital science instruments.

CONCLUSIONS

A study has been performed to demonstratethe feasibility of a landed mission to Neptune’smoon Triton with two landers, deployed froma JIMO follow -on fission reactor powered or-biter (NIMO). Each lander was designed to usea standard MMRTG as its power source. AtBOL an MMRTG would generate 2000 Wt and110 We, however, due to natural decay of theplutonium power source, and degradation of theTE converters, the power would drop to about1735 Wt and 82.6 We after the 18 years transferfrom Earth to Triton. Since Triton is consideredto be the coldest place in our solar system witha mean surface temperature of -235◦C, the lan-der was de-signed with no moving parts (besidethe panoramic camera), and with arrangementsto utilize the excess heat from the MMRTG. AnMMRTG would provide sufficient electric power

12

to support science instruments fulfilling the sci-ence objectives outlined above. It can also beconcluded that radioisotope based power sys-tems provide the best solution for a landed mis-sion to Triton, since solar energy at this distanceis inefficient and mass prohibitive for power gen-eration and fission reactors would be oversizedfor this type of missions.

ACKNOWLEDGMENTS

The author of this paper wishes to thank JimShirley for the science and instrument payloaddefinition, and the Next-Generation ProductDevelopment Team (NPDT), lead by KnutOxnevad. The NPDT team members wereRobert Carnright, Raymond Ellyin, Timothy Y.Ho, Guillermo Olarte, Cesar Sepulveda, ThomasValdez, Kevin Anderson, Nickolas Emis, MichaelD. Henry, Robert Haw, Daniel A. Nigg, SalvadorDistefano, David Hansen, Ellis Miner, LutherBeegle, Eric Wood and Steven D. Keates. Aspecial thanks is extended to JPL’s Prometheusoffice, under the leadership of Garry Burdick,who funded the study. The Office was managedby Erik Nilsen and the RPS studies were coor-dinated by Robert Abelson.

This work was performed at the Jet PropulsionLaboratory, California Institute of Technology,under contract to NASA. Any opinions, findings,and conclusions or recommendations expressedin this paper are those of the author and donot necessarily reflect the views of the NationalAeronautics and Space Administration.

REFERENCES

[1] President G.W. Bush. A Renewed Spirit ofDiscovery, The President’s Vision for U.S.Space Exploration, January 2004.

[2] L.A.M. Benner. The Encyclopedia of Plan-etary Sciences. Chapman-Hall, New York,1997.

[3] T. Balint. Europa Surface Science PackageFeasibility Assessment. Technical ReportJPL D-30050, National Aeronautics andSpace Administration, Jet Propulsion Lab-oratory, Pasadena, CA, September 2004.

[4] NASA. NASA’s Mars Exploration Pro-gram. http://marsprogram.jpl.nasa.gov/,2004.

[5] T. Balint and N. Emis. Thermal Analysisof a Small-RPS Concept for the Mars Net-Lander Network Mission. In M.S. El-Genk,editor, AIP Conference Proceedings #746,Space Technology and Applications Interna-tional Forum (STAIF-2005), Melville, NewYork, 2005.

[6] M. Noca and R. W. Bailey. Mission Tradesfor Aerocapture at Neptune. In AIAA-2004-3843, AIAA/ASME/SAE/ASEEJoint Propulsion Conference and Exhibit,Fort Lauderdale, July 2004.

[7] R. Bailey, J. Hall, T. Spilker, andN. Okong’o. Neptune Aerocapture Missionand Spacecraft Design Overview. In AIAA-2004-3842, AIAA/ASME/SAE/ASEEJoint Propulsion Conference and Exhibit,Fort Lauderdale, July 2004.

[8] NASA-KSC. Launch vehicle database.http://elvperf.ksc.nasa.gov/ elvMap/, July2004.

[9] R. Haw. Mass allocation and Delta V re-quirements for a Triton Lander. Personalcommunications, June 2004.

[10] T.S. Balint. Small Power System TradeOptions for Advanced Mars Mission Stud-ies. In Proc. 55th International Astronauti-cal Congress of the International Astronau-tical Federation, the International Academyof Astronautics, and the International In-stitute of Space Law, (IAC-2004), numberIAC-04-Q.3.b.08, Vancouver, Canada, Oc-tober 2004.

[11] P. Anderson. SRG110 Program Overview.Presented at the SRG110 Program Briefing,Jet Propulsion Laboratory, Pasadena, CA,June 28, 2004.

[12] R. Rovang. Multi-Mission Radioiso-tope Thermoelectric Generator (MMRTG);MMRTG Preliminary Design Review DataPackage. The Boeing Company, 24 Febru-ary, 2004.

13

[13] R.J. Lipinski, S.A. Wright, M.P. Sherman,R.X. Lenard, R.A. Talandis, D.I. Poston,R. Kapernick, R. Guffee, R. Reid, J. Elson,and J. Lee. Small Fission Power Systems forMars. In M.S. El-Genk, editor, AIP Con-ference Proceedings #608, Space Technol-ogy and Applications International Forum

(STAIF-2002), pages 1043–1053, Melville,New York, 2005.

[14] I. Jun. Peer Review for Radiation ShieldingApproach. Government Study Team Tech-nical Baseline Review-2 Presentation, JPL,February 2004.

14