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To Europa and Beyond (Enabled by a Small-RPS)1

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Tibor S. Balint Jet Propulsion Laboratory, California Institute of Technology

4800 Oak Grove Drive, M/S 301-170U Pasadena, CA 91109-8099

818-354-1105 tibor.balint@jpl.nasa.gov

1 0-7803-8870-4/05/$20.00© 2005 IEEE. 2 IEEEAC paper #1585, Version 2, Updated December 11, 2004

Abstract—Solar system exploration (SSE) to the Outer Planets requires internal power sources, since external power generation, relying on solar power, is not very effective or practical. One of the scientifically interesting destinations in this region is Jupiter's moon Europa, where a lander would require internal power sources, such as batteries or Radioisotope Power Systems (RPS) to achieve mission success. This paper discusses a Europa lander concept as a potential adjunct element to NASA's proposed Jupiter Icy Moons Orbiter (JIMO) mission. The lander would operate on the surface for 3 to 14 days, performing astrobiology, geophysics and geological-compositional measurements as defined by the JIMO Science Definition Team (SDT). The power system trade study for this concept compares small-RPSs up to ~50We against primary batteries. Additional SSE destinations are also outlined, where small-RPSs provide a viable option to accomplish science and mission objectives.

TABLE OF CONTENTS

1. INTRODUCTION .....................................................1 2. ESSP CONCEPT.....................................................1 3. RPS APPLICATIONS ..............................................7 4. CONCLUSIONS .......................................................7 ACKNOWLEDGMENTS ...............................................7 REFERENCES .............................................................8 BIOGRAPHY ...............................................................8

1. INTRODUCTION

The National Academy of Sciences identified Jupiter as a critical region of our solar system to search for life's origins and to understand planetary evolution [1]. The Jupiter Icy Moons Orbiter (JIMO) mission was envisioned to enable the exploration of Jupiter's Galilean moons, Callisto, Ganymede and Europa [2]. The JIMO

Science Definition Team (SDT) strongly recommended that a surface science package for Europa be incorporated into the JIMO mission design [3]. It was also suggested that this Europa Surface Science Package (ESSP) could utilize up to 25% of JIMO's payload mass allocation. The current pre-decisional design for the proposed JIMO mission would accommodate a payload of 1500 kg. Consequently, an ESSP could use up to 375 kg mass from JIMO's payload allocation envelope. According to the SDT recommendations, the instruments on the ESSP should target key science objectives by performing astrobiology, geophysics and geological-compositional measurements [3]. Note that cost, planetary protection, and surface contamination issues were not addressed in this study. A detailed feasibility study for the ESSP is provided in [4]. This paper describes a point concept that could fulfill the science objectives for an ESSP, however, the presented configuration is only one of many potential concepts. The study was performed by JPL's Next-Generation Product Development Team (NPDT), in order to cross reference parametric scaling results. The configuration presented here considered small-RPSs to power the ESSP. RPSs present an enabling technology not only for the exploration of Europa, but also for exploration of the Solar System. Some of these mission concepts are briefly introduced at the end of the paper.

2. ESSP CONCEPT

This section introduces the ESSP concept starting with the assumed mission architecture, followed by a discussion on the science instruments, subsystems, power system sizing, and on the radiation and thermal environments. The key drivers for landing on Europa are also discussed.

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Mission Architecture and Assumption

The proposed JIMO orbiter would arrive to Europa (Figure 1), and assume a circular orbit (e=0) at an altitude of 100 km, and inclination of 110° with an orbital period of 2 hr 5 min. During the first 14 days of the proposed 30-day science floor [3] JIMO would map Europa's surface. Once a suitable landing location is selected, the ESSP would uncouple from JIMO and land on the surface. Europa is an airless body. Therefore, regardless of the propulsion system a ∆v of ~1.5-1.65 km/s must be removed propulsively. Three landing methods were assessed and compared, starting with soft landing and using advanced propulsion and throttled engines. Following a 22 m/s de-orbit burn, a ∆v of 1.465 km/s would be removed at perigee, at ~1 km altitude from the ground. Horizontal velocity cleanup, vertical velocity nulling and gravity losses would require an additional 178 m/s to be removed. In order to eliminate the full 1.665

km/s of ∆v, about half of the ESSP mass would be used for propulsion system dry mass and for propellant. Two other landing methods were also considered. With a “stop-and-drop” landing maneuver most of the velocity (~1.4-1.5 km/s) would be removed propulsively, letting the ESSP free fall from ~1 to 2 km. Landing could be achieved with either airbags or with crushable materials, depending on the free fall altitude. However, the propellant saved from these latter methods would be less than the mass penalty from implementing a secondary landing system. Therefore, soft landing was selected, since airbag- and crushable material-based landing would further constrain the already mass limited mission. A comparison of the three landing methods is illustrated in Figure 2. For the de-orbiting maneuver an advanced propulsion system (with hydrazine & oxygen-difluoride bipropellant) would be employed, with an Isp of 355 s. Three pressure-regulated and throttled main engines, and six small thrusters for attitude control, would enable the soft landing maneuver.

For the present analysis the landing site would be located at the coordinates of 55S and 251E (Europa True Equator of date), although the precise site is unimportant for viewing statistics. The line-of-sight visibility from the lander to JIMO was calculated for a total of 14 days. The visibility between the lander and the orbiter has a cyclical nature. Therefore, these results would be repeatable, providing a generic environment to assess telecom needs independent from actual mission dates. All passes given in Table 1 are subject to a 20° horizon constraint. Total

Figure 1. Jupiter’s Galilean moon, Europa

Figure 2. Comparison of landing methods

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telecom times for the assumed 14, 7 and 3 day mission durations would be 133.08 min, 61.78 min, and 36.28 min, respectively. The mission duration was considered in relation to Europa’s ~3.5 day orbital period around Jupiter, where the 3, 7 and 14 days correspond to about 1, 2 and 4 orbits.

The landing surface on Europa is not known. Therefore, an assumption was made to land the ESSP on its lander deck, which would be designed without legs. This simplification from a Viking type lander would save mass and remove the uncertainty of landing the ESSP in a small crevice. Another advantage of this configuration would be the good surface contact. During landing, the (impinging) propulsion plume would likely to melt the icy

surface into slush. The ESSP would land there and would likely freeze into the ice almost immediately following the engine cut-off. This would provide an excellent contact bond between the ice and the lander, allowing for good seismic measurements. Seismic measurements represent a primary science objective. However, this type of soft landing would have a significant impact on sample collection, where the introduced surface contamination could present measurement complications.

In order to reduce mission risk and increase redundancy, dual string design was considered for the ESSP. Note that design principles required 30% contingency on mass and power, which impacted the mass allocation.

Science Instruments

Two sets of science instruments were assessed through detailed concept studies to demonstrate the flexibility of the available options that could achieve the science objectives defined by the JIMO SDT. Both sets were selected based on discussions with the SDT. The first set included a Gas Chromatograph/Mass Spectrograph (GC/MS), a micro-seismometer, a magnetometer, an imager and radiation sensors. Wet chemistry and Raman spectroscopy instruments were left out due to mass limitations. (Preliminary calculations indicated that each additional kilogram of mass from instruments or batteries would result in an almost five fold increase in total wet mass on the ESSP.) The second instrument set included an organic detector connected to a mini-mass spectrometer, a seismometer, a magnetometer, a radiation sensor and a panoramic camera. A summary of the second set of instruments is shown in Table 2. These instruments were only defined for mass, power and system sizing purposes, without an attempt to influence instrument choices on future landing concepts. It is also expected that any of the heritage instruments may have to be modified for the Europa environment. The ESSP concept described in this paper is based on the second set of instruments.

Avionics and Telecom Systems

The command and data handling (C&DH) system was sized for data collected by the science instruments and impacted by the telecom windows. The C&DH system was assumed with a dual-string Harris R3000 controller and radiation hardened field programmable gate arrays. The data would be collected on 256 MB of memory before being transferred to JIMO. While most of the data processing was assumed using the C&DH system, an alternative approach could shift data processing to the instruments, reducing both mass and power requirements on the C&DH system. However, this would require additional technology development. The dual-string ElectraLite telecom package would communicate with JIMO at L-band through an UHF monopole antenna with a ground plane. (L-band was chosen to mitigate the

TABLE 1. Telecom opportunities over 14 days

Start (day)

Duration (min)

Start (day)

Duration (min)

0.90 5.03 8.15 5.51 0.99 5.54 8.23 4.76 1.94 3.88 8.32 2.78 2.03 5.18 9.18 5.22 2.11 5.37 9.27 5.28 4.48 5.09 9.36 3.71 4.57 5.52 11.55 2.89 4.66 4.84 11.64 5.19 4.74 2.95 11.73 5.48 5.52 3.96 11.81 4.68 5.60 5.20 12.68 4.09 5.69 5.33 12.76 5.22 5.78 3.89 12.85 5.22 7.98 2.63 12.94 3.50 8.06 5.14

TABLE 2. Science instrument assumptions

Instrument Mass (kg)

Power (W)

Data rate & Data volume (compressed)

Operation

Organic Detector

3.00 10.0 ~1 kb 1 hour

Mini-MS

0.50 1.0 ~1 Mbits 1 hour

Seismo-meter

2.11 2.7 ~16.6 kbps Cont.

Magneto-meter

0.70 0.5 ~2.5 kbps Cont.

Radiation Sensor

0.11 0.3 ~1 Mbits/day Cont.

Pan Cam 0.50 4.3 ~8 Mbits / frame (raw); 5:1 compr.

~15 min

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magnetic field interference associated with Jupiter's environment.) This configuration could support a data rate up to 4096 kbps (assuming JIMO is in a 100 km circular orbit around Europa). The daily data volume over the communication opportunities could be as high as 300 Mbits.

Power System and Sizing

The ESSP described here was considered with small Radioisotope Power Systems. RPSs present a reliable continuous internal power source for missions where solar power generation is not feasible: at Jupiter for example. The initial mass limitations for the ESSP make it evident that large RPSs currently under development (i.e., MMRTG ~42.7 kg and SRG ~34 kg) would be mass prohibitive. Although smaller radioisotope power systems are not yet available, ongoing efforts reported in [5] forecast the potential availability of such systems by as early as 2011. The power system assumed in this assessment utilized a number of single GPHS module based small-RPSs, with an estimated mass of ~5-6 kg each. A conceptual design is shown in Figure 3. Further information on this new and exciting technology can be found in [5]. Radioisotope power systems have a low acceleration load tolerance of ~40 g. This makes soft landing a good candidate for RPS enabled landers. (Airbag landing with acceleration loads up to 600g or crash-landing with acceleration loads in the thousands of

g’s would not be suitable for this power system, based on the present MMRTG requirement of 40 g. Requirements for small-RPSs are not yet established). Outer planetary missions, especially when using electric propulsion, could result in mission durations as high as 10 to 20 years. For such missions the radioisotope power systems should be sized to account for their graceful degradation, due to the radioisotope decay process and to the degradation of the power converters. The Plutonium-238 fuel, used in RPSs, has a half-life of 87.74 years. This causes a natural (exponential) degradation of the generated heat, which can be approximated as a 0.79% drop in thermal power output per year. In addition, the thermoelectric converters degrade at a rate of about 0.8%. This results in an approximate total power loss of ~1.6% per year. For example, after 10 years (not including 2 years of storage) a single GPHS module generating ~250 Wt at BOL (fabrication) would produce about 227.4 Wt, while the electric power would drop from ~13.75 We to ~11.36 We, based on today’s typical static thermoelectric conversion efficiency of 5.5%.

These instrument mass and power assumptions, helped to size the power, C&DH, and telecom systems, and consequently influenced the final mass. The instruments listed here are at TRL 3 or above as of today, and could be brought to TRL 6 within a few years of development.

Figure 3. Small-RPS concept with one GPHS module

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The power system was sized, based on the assumed power requirements. The shortness of the mission allowed for the construction of full operational scenarios for 3, 7 and 14 days. The radioisotope power systems were sized using a JPL-developed power simulation tool (MMPAT). The results indicated that 50 We of electric power provided by small-RPSs, with an addition of two 10 Ah Li-Ion secondary batteries, would enable continuous operation over any mission duration. However, longer mission times would further increase the shielding mass requirements, and rapidly exceed the mass allocation. (In comparison, a battery could power a 3-day long mission. For that configuration, the primary battery was sized from a direct summation of the energy requirements over the mission. It was found that for a 3-day mission a 100 Ah lithium thionyl chloride (LiSOCl2) primary battery would be able to support the operation. With increased mission duration the batteries – and shielding – would size up proportionally, exceeding the mass limit at ~4 days on the surface.) Accounting for the exponential natural decay of the radioisotope and for the degradation of the thermoelectric converter, the electric power generated with 3 GPHS modules after 10 years (not including 2 years of storage) would be ~49.6 We. This calculation assumed an 8% thermal to electric conversion efficiency, which could be achieved for this technology by a technology cut-off date of 2012. (Note that an MMRTG will have a conversion efficiency of around 5.5-6.5%, while laboratory demonstrations on segmented thermoelectric converters reached above 12%.)

The ESSP concept with 3 small-RPSs is shown in Figures 4 and 5. Note that this conceptual design was not optimized and the results were used for comparison purposes against parametric scaling assessments. Therefore, further studies would be necessary to refine these designs concepts. During landing, throttled engine drivers require high power, typically supplied by thermal batteries. A thermal battery was selected for descent, which is normally sized for lifetime (tens of minutes) and

for peak power (up to ~1 kW). During powered descent it would provide 70.1 W to the attitude control system, 11.5 W to the command and data handling, 20.7 W to the power system, 181.7 W to the propulsion system, 40 W to the telecom system and 16 W to the thermal system, totaling 340 W without the 30% contingency. Power requirement for the powered descent with contingency would be 442 W. During surface operation three operating modes were assumed with the second set of science instruments. There were high- and low-powered science modes and a telecom mode. During the high-powered mode the organic detector and mini-MS would operate for 1 hour, using ~10 W and 1 W of power in addition to the continuously operating instruments. Engineering sensors for pressure and temperature would take measurements in a periodic basis to conserve energy. During low-powered mode the seismometer (~2.7 W), magnetometer (~0.5 W), and the radiation sensor (~0.3 W) would operate continuously. When all other subsystems are accounted for, namely the C&DH (~9-12.5 W), power conditioning (1-2 W), and telecom in idle mode (2 W), the total power would be 44.3 W for the high-powered science mode and 22.3 W for the low-powered science mode. The ElectraLite telecom package would use about 40 W of power during the telecom opportunities, and when the subsystems are accounted for, the total system power requirement during the telecom window would be 76.3 W. During the high-powered mode, the small-RPSs and the batteries would supply power simultaneously, while during low-powered mode the RPS would recharge the batteries. Two instruments with short operations are also considered. The imager would use ~4.3 W of power during its 15 minutes operation at the beginning of the mission. A sample acquisition system (not listed in Table 2) would require about 6 W of power during the 10-15 minutes of sample acquisition. Its assumed 2.6 kg mass is highly dependent on the actual design. A sampling device for soft landing could require special care in order to ensure

Figure 4. ESSP concept

Figure 5. ESSP concept – deck layout

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the collection of non-contaminated samples.

Radiation Environment and Mitigation Strategies

Europa has a high radiation environment of ~180 kRad per day. In addition, JIMO’s low thrust trajectory and long mission time would result in a significant amount of time spent in high radiation environments. This includes spiraling out of the Van Allen radiation belt and into Jupiter’s own high radiation environment. Additional radiation exposure, although smaller in magnitude, consists of galactic radiation and radiation from JIMO’s fission reactor. It was estimated that without radiation shielding the JIMO spacecraft would be subjected to about 60 MRad of total ionizing radiation during its tour [7]. Mitigation of this high radiation necessitates significant shielding on JIMO and on the ESSP, and the use of high radiation tolerant components. For a 2012 technology cutoff date, 3 years prior to the proposed launch date, a radiation tolerance level of 1 MRad was assumed, which is about twice of the current state-of-the-art (~300-500 kRad), but reasonable.

Two options were considered for shielding calculations. For the first option, spot shielding was assumed on science instruments and on electronic components, based on bounding boxes representing the respective surface areas. For the second option, housed shielding was considered, placing a bounding dome over the radiation sensitive components. The shielding surface area for the present not optimized concept for spot shielding and for housed shielding was estimated as ~2.09 m2 and 2.66 m2, respectively. It was also calculated that 1” of Aluminum could effectively shield against the radiation during the proposed JIMO tour [7]. If all of the required shielding were placed on the ESSP alone it would increase its total mass beyond the allocated mass limit. For spot shielding or housed shielding the shielding mass would be 144 kg and 182 kg. If adequate shielding would be provided on JIMO (~1”) until the release of the ESSP, the radiation environment acting on the ESSP should be mitigated for the surface operation only. For 7- or 14-day operations 5 mm or 10 mm or shielding would be required on the

ESSP to reduce the radiation dose below 1 MRad (including a radiation design factor – RDF – of 2). The corresponding shielding mass for 7-days would be ~28 kg and 36 kg for spot or housed shielding, respectively. For 14-days the shielding mass would be twice as much. It should be noted that these are conservatively high mass values, based on a non-optimized ESSP design. Detailed point designs would allow for refined shielding mass estimates, which could significantly reduce the relevant mass requirements. The refinements could include: (1) better estimates of instrument volumes; (2) optimized integration of components and accounting for natural shielding from the propellant on the ESSP; and (3) potential integration with JIMO to benefit from it's shielding.

Thermal Environment

The thermal environment was calculated for an ESSP powered by small-RPSs. The major thermal design goals for the ESSP were to maintain the electronics above their survival temperature and to manage the heat generated by three small-RPS systems. It was assumed that a Warm Electronics Box (WEB) would house all of the instruments. The WEB would be covered by Multi Layer Insulation (MLI). During normal operation the instruments and subsystems would generate sufficient heat to keep the WEB above –20°C. Each RPS system would generate about 220 Wt of thermal power (dependent on the JIMO mission duration and the resulting power degradation) that should be radiated out by the RPS housing and fins or conducted to the WEB. To assist in the radiated heat transfer, each small-RPS unit would receive high-emittance coating, such as black paint, while conductive heat transfer could be improved by a 1/8-inch Aluminum doubler plate. The thermal system weight associated with the added doublers and the MLI blankets (with minor impacts from heaters and thermal sensors) would amounts to about 10.1 kg. When the Pan Cam is deployed, additional heaters will be needed to maintain the temperature survival limits for the Pan Cam and for the mast motors, by using 1.8 W and 0.3 W heaters, respectively. It was also found that heat generated through normal operation of the electronics would maintain the Warm Electronics Box above a survival temperature, making the use of Radioisotope Heater Units (RHU) unnecessary for component heating.

Mass Allocation and Configuration Summary

Based on the above considerations, a summary of the mass breakdown for a 7-day operating small-RPS powered ESSP concept is provided in Table 3. For a 14-day mission the mass would be higher due to the increased shielding requirement. (Shielding mass is distributed and included in the mass allocation table entries). In addition, concept drawings were generated to visualize the designs and to help with a better understanding of the requirements for shielding and

Table 3. Mass allocation for the ESSP (RPS; 7 days)

Mass allocation Science Instruments 6.9 kg 1.8% GN & C 17.1 kg 4.5% Avionics 3.2 kg 0.8% Thermal 10.1 kg 2.7% Power 38.0 kg 10.0% Telecom 6.4 kg 1.7% Structures 56.7 kg 14.9% Contingency 54.6 kg 14.4% Propulsion (Dry) 43.7 kg 11.5% Propellant (Wet) 142.7 kg 37.6% Total Wet Mass 379.4 kg 100.0%

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structures. Figures 4 and 5 show the ESSP concept powered by 3 small-RPSs. The systems and instruments were placed on the lander deck in a logical order. For example, the organic detector, the mini mass spectrometer and the sampling mechanism are positioned adjacent to each other. Small differences could occur between the two shielding strategies, which are not shown in the CAD drawings. For a spot-shielding configuration, with the dome removed, the present drawings provide a reasonable instrument placement. For housed shielding, the radiation sensor should be placed outside of the shielding dome. The three small-RPSs were added to the lander outside of the shielding and thermal protection, allowing the waste heat to radiate to space. The Pan Cam was located on the truss system. After landing, the mast (shortened from the MER design) would deploy to a vertical position allowing for good panorama visibility. Thermal (MLI) blankets – not shown in the figures – would wrap the ESSP, providing thermal insulation not only to the instruments, but also to the fuel tanks. The UHF antenna (not shown) could be placed on the top of the ESSP for good visibility with JIMO. The lander deck would have a diameter of about 1.2 m, and the height of the ESSP would be ~0.9 to 1 m.

3. RPS APPLICATIONS

Radioisotope power systems could provide advantages to other outer planets missions as well where solar power is not available. Solar flux decreases as a function of one over distance squared from the Sun. As shown in [6] solar power generation is less mass effective beyond ~3.5-4.5AU. This necessitates the use of internal power sources, such as nuclear fission reactors or radioisotope power systems. These two enabling technologies address different power ranges. Fission reactors could provide power from a multi-kilowatt to the multi-megawatt range, and are planned for flagship class missions to the outer planets (e.g., JIMO). Looking at lower power ranges, the present generation of RPSs could generate power from about 110We (MMRTG, SRG) to about a kilowatt (three GPHS-RTGs on the Cassini mission). Note that the GPHS-RTG technology is discontinued. The last unit will be used on the 2006 Pluto-Kuiper Belt mission. It will be replaced by the MMRTG and SRG designs as early as in 2009. Smaller radioisotope power systems are also under investigation, providing power in the tens to few hundreds of milliwatts and in the tens of watts ranges. Small-RPSs could power single instrument based adjunct mission elements to large outer planetary missions (due to the economy of scale), while stacked standard RPSs could scale up to the kilowatt power level. Multi-hundred watt RPSs could be used to explore a large number of outer planetary destinations both in space and on planetary surfaces, while kilowatt class RPSs could be used for radioisotope electric propulsion (REP).

4. CONCLUSIONS

The study described here identified three key-drivers bounding the trade space for an ESSP design. The first driver was the initial constraints, imposed by the proposed JIMO mission with a 375 kg initial mass limit for the ESSP, and by the TRL cutoff date of 2012. The second driver was the radiation environment, potentially exposing JIMO and the ESSP to as much as 60 MRad of total ionizing radiation dose. This must be mitigated at a potentially significant shielding mass penalty. The third driver was the landing method, which according to the rocket equation would require about half of the mass allocation for the propulsion system alone (including propellant and propulsion dry mass).

It was also demonstrated that an ESSP with a mass allocation of 375 kg would be feasible using either targeted or un-targeted soft landing methods. For a 7- or a 14-day mission three small-RPSs could provide sufficient power for the ESSP. Improved thermoelectric power conversion efficiency could further reduce the number of small-RPSs. A short 3-day mission could be achieved with a primary battery. The mass cross over between battery and RPS powered landers is expected to be around 3 to 4 days of surface operation. With these configurations the ESSP could achieve the science objectives identified by the JIMO SDT.

It is further emphasized here that the present concepts is not representative of a potential final ESSP design. The concept is derived to assist with the understanding of the trade space and to provide reference points against parametric scaling results. Since the concepts were derived from a fast turnaround study, the accuracy of the results could only be used to support an argument that an ESSP mission is conceivable within the 375 kg mass allocation. Any further conclusion should require detailed follow on studies, exploring point designs.

While small-RPSs may not be able to support larger New Frontiers class science missions to the Outer Planets, they could be considered in a supporting capacity. They could be used as adjunct elements on large flagship class missions, such as on the proposed JIMO mission.

ACKNOWLEDGMENTS

I wish to thanks my colleagues at JPL who contributed to this report, including members of the Next-Generation Product Development Team. The ESSP feasibility study, which provided parametric scaling results, was funded by Curt Niebur, Program Scientist for the Jupiter Icy Moons Orbiter at NASA HQ. I also wish to extend my thanks for the support of Fuk K. Li (Deputy Director for Office-6X), Frank Jordan (Program Manager for Office-61X), Dave

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Senske (Program Manger for Office-63X), and Erik Nilsen (Section Manager for Section 312) at JPL. This work was performed at the Jet Propulsion Laboratory, California Institute of Technology, under contract to the National Aeronautics and Space Administration.

REFERENCES

[1] NRC. New Frontiers in the Solar System, an integrated exploration strategy, Technical Report, Space Studies Board, National Research Council, Washington, D.C., 2003.

[2] NASA. Jupiter Icy Moons Orbiter, Web site: http://www.jpl.nasa.gov/jimo/, accessed July 2004.

[3] Greeley, R., Johnson, T., Report of the NASA Science Definition Team for the Jupiter Icy Moons Orbiter (JIMO), Technical Report, National Aeronautical and Space Administration, February 2004.

[4] Balint, T. Europa Surface Science Package Feasibility Assessment, Report Number JPL D-30050, Jet Propulsion Laboratory, California Institute of Technology, September 2004.

[5] Abelson, A.D., Balint, T.S., Marshall, K.E., Noravian, H., Randolph, J.E., Satter, C.M., Schmidt, G.R., and Shirley, J.H., Enabling Exploration with Small Radioisotope Power Systems. Technical Report JPL Pub 04-10, National Aeronautical and Space Administration, Washington, D.C., September 2004.

[6] Balint, T., Small Power System Trade Options for Advanced Mars Mission Studies, 55th International Astronautical Congress, Paper Number: IAC-04-Q.3.b.08, Vancouver, Canada, October 4-8, 2004

[7] Jun, I., Single Brick GPHS Total Dose Estimates, Personal Communications, Mars Smart Lander Study, Jet Propulsion Laboratory, California Institute of Technology, December 23, 2003.

[8] Jun, I., Peer Review for Radiation Shielding Approach. Government Study Team Technical Baseline Review-2 Presentation, Personal Communications, Jet Propulsion Laboratory, California Institute of Technology, February 2004.

BIOGRAPHY

Tibor Balint is a senior engineer within the Mission Systems Concepts Section (312) at the Jet Propulsion Laboratory, Pasadena, CA. His research interests include nuclear systems, mission architecture, systems engineering, and biomechanical engineering. Dr. Balint obtained an MSc degree in mechanical engineering from the Technical University of

Budapest, Hungary, in 1987; an MPhil in chemical engineering from the University of Exeter, UK, in 1989; a PhD in engineering from the University of Warwick, UK, in 2001; and a Master of Space Studies (MSS) degree from the International Space University, Strasbourg, France, in 2002. He also worked as a nuclear design engineer for 9 years at Ontario Hydro, Canada, conducting nuclear safety analysis.