REACTIONN: A Nuclear Electric Propulsion Mission Concept to the Outer Solar System

Transcription

1 REACTIONN: A Nuclear Electric Propulsion Mission Concept to the Outer Solar System A. Charania *, B. St. Germain, J. G. Wallace, J. R. Olds SpaceWorks Engineering, Inc. (SEI), Atlanta, GA, A concept assessment is presented of a follow-on mission utilizing the same nuclear electric propulsion technology path envisioned by the Jupiter Icy Moons Orbiter (JIMO) program but to a different destination, namely to the outer solar system including Pluto, Charon, and objects in the Kuiper Belt. The Rapid Electric Acceleration Coupling ION and Nuclear (REACTIONN) concept utilizes currently planned nuclear reactors along with high power ion thrusters. Technology assumptions are derived from power and propulsion technology investment plans from NASA s Project Prometheus. The REACTIONN concept is presented as a follow-on mission to utilize the technology investments made for the JIMO project. A Reduced Order Simulation for Evaluation of Technologies and Transportation Architectures (ROSETTA) model was developed to perform both concept performance and sizing. In addition, life cycle cost and operations estimates are also examined for the concept. Nomenclature V = Delta-V, m/s η = efficiency, % DDT&E = Design, Development, Testing, and Evaluation Isp = specific impulse, seconds T/W = thrust to weight TFU = theoretical first unit TOF = time of flight, years I. Introduction and Motivation here is a new Vision for Space Exploration from NASA that includes multiple programs for transformational T capabilities in the exploration of the Solar System. One tier of that vision rests on NASA s Project Prometheus, a set of programs to develop larger nuclear power and in-space propulsion technologies 1. The first manifestation of this technology development spiral will be the Jupiter Icy Moons Orbiter (JIMO), a follow-on mission to the Galileo spacecraft that will utilize nuclear electric propulsion (NEP) with a suite of new, higher power instruments 2. Such technology investment will have potential for use in subsequent missions to the outer Solar System. The assessment presented here shows the spiral development path of the JIMO mission and the specific costs associated with that path. Acting on a request from NASA s Marshall Space Flight Center (MSFC), SpaceWorks Engineering, Inc. (SEI) fashioned a concept that assumes development of JIMO and its associated technologies. The Rapid Electric Acceleration Coupling ION and Nuclear (REACTIONN) concept utilizes currently planned nuclear reactors along with high power ion thrusters. NASA MSFC representatives provided guidance on a specific followon destination. The results detailed here include performance analysis and life cycle cost assessment of a first order conceptual spacecraft design for a Nuclear Electric Propulsion (NEP) mission to Pluto and the Kuiper Belt (see Fig. 1). The results are a collaborative product of SpaceWorks Engineering, Inc. (SEI) with specific disciplinary * Senior Futurist, SpaceWorks Engineering, Inc. (SEI), and Member AIAA. Director of Advanced Concepts, SpaceWorks Engineering, Inc. (SEI), and Member AIAA. Project Engineer, SpaceWorks Engineering, Inc. (SEI), and Member AIAA. President, SpaceWorks Engineering, Inc. (SEI), and Associate Member AIAA. Copyright 2004 by A. Charania, B. St. Germain, J. G. Wallace, and J. R. Olds. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. 1

2 assistance provided by personnel at MSFC with regards to trajectory determination. A Reduced Order Simulation for Evaluation of Technologies and Transportation Architectures (ROSETTA) model was developed to evaluate concept performance, sizing, and cost. Figure 1. Illustration of REACTIONN Concept at Pluto. II. Design Process A. Selection of Mission SpaceWorks Engineering, Inc. (SEI) proposed various missions for follow-on to JIMO predicated on Project Prometheus power and propulsion technologies. The specific missions included the following: 1) Semi-permanent L-type stations 2) Near earth asteroid manned colonization 3) Multiple comet sample return (multiple return canisters/pods with one mother ship) 4) Saturn orbiter (Saturn ring sample return) 5) Europan orbiter/lender with power beaming 6) Europa ocean underwater station 7) Pluto/Neptune/Uranus/Kuiper Belt probe (all four in one mission, "Voyager"-like but with NEP) 8) Optical interplanetary communication 9) Large, high power antennas across the solar system 10) L-point astronomical observatory 11) Long duration Martian ground or aerial vehicle 12) Cometary impactor 13) Missions to the moons of Mars (power beaming) 14) Jupiter atmosphere sample return 15) VASiMR use of NEP 16) Laser light craft from Martian surface using nuclear power for laser NASA MSFC representatives choose two specific missions from the above: a Pluto/ Kuiper Belt mission and the optical interplanetary communication mission. The former mission was the one subsequently down-selected for examination and reported here (see Fig. 2). 2

3 Figure 2. Notional Representation of REACTIONN Concept. B. ROSETTA Model Modeling helps to determine the properties of a technically feasible design. In the conceptual design stage, this can include use of monolithic synthesis/sizing codes or an integrated multi-disciplinary environment. These models are representations of the real world based on processes in terms of physics, human operations, financials, etc. A baseline concept, the initial starting point for design space investigation, can be developed from high fidelity analytical tools. In order to negate the computational expense involved with the use of Monte Carlo uncertainty simulation (potentially thousands of converged designs), a time-efficient process is needed for concept simulation and technology evaluation. Meta-models, or representations of these detailed models, can be employed for situations where computation and monetary expense are to be minimized. Therefore, the Reduced Order Simulation for Evaluation of Technologies and Transportation Architectures (ROSETTA) modeling process is employed. A ROSETTA model is a spreadsheet-based meta-model which is a representation of the design process for a specific architecture (e.g., ETO, in-space LEO-GEO, HEDS). ROSETTA models contain representations of a baseline design into which technologies can be infused. The model is based upon higher fidelity models (i.e. POST, APAS, CONSIZ, CHEBYTOP, etc.) and refined through updates from such models. The goal is for the model to execute each architecture simulation in only a few seconds. ROSETTA models can be separated into three categories based upon the output metric provided by the model: 1) Category I: Produces traditional physics-based outputs such as transportation system weight, size, payload, and the NASA metric in-space trip time 2) Category II: In addition to above, adds additional ops, cost, and economic analysis outputs such as turnaround-time, LCC, cost/flight, ROI, IRR, and the NASA metric price/lb. of payload 3) Category III: In addition to above, adds parametric safety outputs such as catastrophic failure reliability, mission success reliability, and the NASA metric probability of loss of passengers/crew Outputs from the model measure progress towards customer goals ($/lb, mass, power level, turn-around-time, safety, etc.). ROSETTA models contain representations of the full design process. Individual developers of each ROSETTA model determine the depth and breadth of appropriate contributing analyses. Generally, such models have more assumptions and fewer links in a typical design structure matrix (DSM) than higher fidelity models due to need for faster calculation speeds. Created at the Georgia Institute of Technology and enhanced at SpaceWorks Engineering, Inc. (SEI), this modeling process was adopted by the Integrated Technology Assessment Center (ITAC), sponsored by NASA Marshall Space Flight Center s Advanced Space Transportation Program (ASTP). The ROSETTA model for assessment of the REACTIONN concept contains 11 disciplinary worksheets and an Inputs/Outputs (I/O) worksheet. These include the following disciplinary components: trajectory, electric propulsion, reactor power, power budget, aft attitude control, forward attitude control, telecommunications, thermal systems, subsystems, sizing, mass, and cost (non-recurring and acquisition). Each component has associated internal calculations and is linked to other disciplinary components with feedback loops present within the most coupled disciplines (i.e. power, propulsion, mass). Most of the sizing algorithms are based upon parametric scaling and or physics-based simulation. 3

4 C. Detailed Cost Assessment This assessment utilized two specific cost models to estimate Design, Development, Testing, and Evaluation (TDDT&E) and Theoretical First Unit (TFU) costs. These costs include accounts for both specific hardware costs and associated system integration costs. The first method relies on an ad-hoc cost model developed specifically for this concept within the ROSTETA model. The other method is based upon the NASA/Air-Force Cost Model (NAFCOM) 2004 edition which includes the Spacecraft Operations Cost Model (SOCM). Unless otherwise noted, cost estimates presented here do not reflect the cost of technology maturation to a Technology Readiness Level (TRL) of six, science instrument development costs, operations costs, or launch vehicle/in-space assembly costs. III. Disciplinary Assumptions For all of the disciplines contained within the ROSETTTA model the objective was to develop a sufficiently robust design simulation such that the entire spacecraft could be parametrically scaled based upon various combinations of specific top-level input parameters (such as required payload mass). Assumptions were in part based upon both JIMO specific and other outer planet mission designs 3,4. Trajectory analysis performed for this concept was based on a direct, rendezvous with Pluto. The CHEBYTOP trajectory code was utilized by NASA MSFC personnel to generate a curve fit of spacecraft thrust-to-weight (T/W) ratio versus V required for Pluto rendezvous (see Fig. 3). CHEBYTOP is more accurate when the T/W values are in the 0.1g range and below. The analysis accounted only for the heliocentric portion assuming a gamma of zero degrees at Earth departure and Pluto arrival travel distance to Pluto of 30.5 Astronomical Units (AU). For a NEP mission nominally some other propulsion system needs to move the spacecraft to a nuclear safe orbit, such as a 1,000 km circular Earth orbit or higher. For this analysis the starting point of the trajectory is just such an orbit. A hydrazine (N2H4) propellant attitude control system (ACS) is included with locations at the forward and aft sections of the spacecraft. Each section is sized to provide 50 m/s of attitude control with an Isp of 220 seconds. 60 Input: Vehicle Power, Initial Mass, Isp Output: Delta-V, Time of Flight (TOF) Pluto Rendezvous Delta-V, km/s Data Curve Fit 0 1.E-05 1.E-04 1.E-03 1.E-02 1.E-01 1.E+00 NEP Spacecraft Initial T/Wo, referenced to Earth g Figure 3. NEP Trajectory Curve Fit. The overall configuration of the spacecraft, similar to JIMO reference designs, consists of a central truss section with multiple sub-systems latched at various positions (see Fig. 4). The nuclear reactor system (reactor, containment vessel, and cylindrical shielding) is located at the notional front of the spacecraft along with a power conversion subsystem. A geometrical representation of the spacecraft is included in the ROSETTA model to attempt physically size the system to determine the mass of the required central truss. The major components of the spacecraft (given in order from the front to back) include: reactor, shield, power conversion, forward ACS (thrusters, tank, and propellant feed system), radiators (primary), communications/science payload, magnetometer booms, science payload, radiators (secondary), propellant tanks, aft ACS (thrusters, tank, and propellant feed system), power processing unit/propellant feed system, thrust structure, thruster platform connectors, and thruster grids. A mass breakdown statement (MBS) is developed that encompasses all the subsystem masses and spacecraft sizing. Overall mass is then fed back to various other disciplines including trajectory and electric propulsion. 4

5 Figure 4. REACTIONN Baseline Spacecraft Schematics. 5

6 Propulsion and power systems were based upon ion thrusters driven by an advanced nuclear reactor system (see Fig. 5). Electrostatic ion thrusters similar to NASA s NSTAR and NEXIS thrusters were assumed with a specific power of 1.2 kg/kw with a total thruster throughput of 2,000 kg. The efficiency of converting electric power to thrust power (thruster efficiency) is 79.7% based upon xenon propellant and an Isp of 4,050 seconds. The baseline power system utilized for the concept was a Particle Bed Reactor (PBR) consisting of seven fuel elements (with an assumed core density of 1,600 kg/m 3 ) 5. An advanced power conversion system was assumed with an efficiency of 30%. A power budget was developed that accounted for several losses throughout the power chain (see Fig. 6). Losses for several key system systems were included including thruster efficiency, power conversion efficiency, and power conditioning efficiency. Secondary losses such as that from the propellant feed system, power processing unit, shielding, cabling, radiation, and thermal losses were also included in the power budget. The telecommunications system consists of two 5 m X/Ka band antennas, each with a transmitting power of 5 kw. The antenna system was sized for maximum line of sight distance of 45 AU utilizing 34 m Deep Space Network (DSN) receiving antennas. A five section primary Liquid Drop Radiator (LDR) system is included on the spacecraft with a mass/area ratio of 0.25 kg/m 2 and radiator thickness of m. Different radiator systems to dissipate heat for the main reactor system and electric propulsion system are included. Additional subsystems included on the spacecraft include magnetometer booms (2), data processing, navigation sensing systems, and command/data handling. Science instrumentation was not defined specifically for this concept and only represented as a required payload that the spacecraft must transport. POWER PROPULSION RADIATORS POWER PROCESSING UNITS ELECTRIC THRUSTERS REACTOR SHIELDING POWER CONVERSION POWER MANAGEMENT AND DISTRIBUTION VEHICLE SYSTEMS XENON PROPELLANT RADIATORS SPACECRAFT BUS SCIENCE PAYLOAD SPACECRAFT Figure 5. Components of NEP System on REACTIONN Spacecraft. A cost model specific to this concept was developed. Both non-recurring and acquisition costs were estimated based upon various categories of historical and analogous cost data. Both weight based and unit-based cost estimating relationships (CERs) were used for this model. Additional program costs such as System Test Hardware (STH), Integration, Assembly, & Checkout (IACO), System Test Operations (STO), Ground Support Equipment (GSE), System Engineering & Integration (SE&I), and Program Management (PM) were estimated as various percentages of the based hardware cost. A fifty-five percent margin was applied to all output costs. For CERs that used input mass estimates from the MBS, these masses did not include associated performance margins (input masses do not include margin). 6

7 95.0% hppu 94.1% Total 25.8% PPU Reactor EFFICIENCIES FOR NUCLEAR POWER SOURCE AND ELECTRIC PROPULSION 100.0% Efficiency η-other 99.0% hshielding 98.0% Total η-cabling Shielding η-shielding 98.0% η-power-conversion η-power-conditioning 30.0% hpower-conversion 29.7% Total η-ppu Power Conversion η-electric-thrusters 29.1% η-propellant-feed-system 95.0% hpower-conditioning 94.1% Total PMAD / Power Cond. 27.4% Value Description 99.5% Including radiation and thermal, for both nuclear and solar power systems 99.5% For both nuclear and solar power systems 99.0% 30.0% The efficiency of power conversion for the reactor 95.0% The efficiency of power conditioning for the reactor 95.0% The efficiency of converting electric power to thrust 79.7% power (thruster efficiency), based upon xenon propellant 95.0% 99.0% Total Hotel Loads 99.0% Total Science Loads 99.0% Total Communication Loads 27.1% 27.1% 27.1% 79.7% helectric-thrusters 78.9% Total Electric Thrusters 20.3% 95.0% hppu 94.1% Total Propellant Feed System 24.2% Figure 6. Nuclear Power Source: Baseline Spacecraft Efficiency Chain. IV. Concept Assessment The concept shown here will have a primary science mission to orbit Pluto and Charon with additional capability to tour the Kuiper Belt. Some of the fundamental assumptions include use of Project Prometheus power and propulsion technologies in a post JIMO timeframe (past calendar year 2015). Specifically, this Rapid Electric Acceleration Coupling ION and Nuclear (REACTIONN) concept incorporates Nuclear Electric Propulsion (NEP) consisting of a fission reactor and electrostatic ion thrusters. A baseline design was developed using the ROSETTA modeling process with subsequent trade studies to develop a better understanding of the design space for such a mission. A. Baseline Overview The baseline REACTIONN concept is a nuclear electric propulsion (NEP) spacecraft with a baseline destination of Pluto rendezvous and orbit capture with additional mission requirement for Kuiper Belt follow-on mission (see Fig. 7). Baseline trajectory analysis yielded a V of 47.7 km/s with flight time to Pluto of approximately 5.2 years and an additional V of 2 km/s for a Kuiper Belt excursion. The nominal payload mass of 1 MT as shown in Table 1 is sized for a reactor power level of 1 MW (30% power conversion efficiency). This results in a dry mass of approximately 10.8 MT (with payload) and a near Earth departure mass (NEDM) of approximately 50 MT (including a 15% mass growth margin). Table 2 shows a detailed two-level summary Mass Breakdown Statement (MBS). The entire spacecraft stack is slightly longer than 100 m. The baseline cost assessment yielded a total nonrecurring and acquisition cost of $2.82 B (FY2003 with 55% cost margin) consisting of $342 M in acquisition costs (see Table 3). As shown in Fig. 8, a large percentage of the entire non-recurring cost is due to the nuclear reactor and power and conversion system. 7

8 Total Length = 115 m Maximum Width = 101 m Total Power Required = 1,000 kw Isp = 4050 sec IMLEO = MT Dry Mass (with 1MT payload) = 10.8 MT 100 meter 50 meters 0 meters Figure 7. REACTIONN Baseline Spacecraft Summary. Table 1. REACTIONN Baseline Spacecraft Power Budget. Item Power, kw Communication Loads 5.0 Science Loads 25.0 Hotel Loads 5.0 Propellant feed systems 2.1 Power required for electric thrusters PPU 13.6 PMAD / Power Conditioning 14.4 Power conversion losses Shielding losses 10.0 Total cabling losses 19.3 Total other losses 19.4 Total power required from reactor 1,

9 Table 2. REACTIONN Baseline Spacecraft Mass Breakdown Statement (MBS). Item Level 2 Mass, kg Level 1 Mass, kg 1.0 Nuclear reactor power system 4,680 Nuclear core 300 Containment vessel 2,370 Radiation shield 1,175 Power conversion 200 Power conditioning Propulsion 2,740 Electric propulsion system 2,700 Attitude control system Thermal Control 105 Primary radiators 60 Secondary radiators 20 Misc. blankets, heaters, thermostats Primary Central Structure Data Processing 70 Attitude/Orbit determination 20 Attitude/Orbit control 20 Device pointing 20 Integrated function Navigation Sensing/Control 40 Celestial 20 IMU Telecom 200 TCM Module 25 Command and data handling 10 Communications payload Growth Margin (15%) 1,280 Dry Mass (w/o payload) 9, Payload 1,000 Dry Mass (w payload) 10, Propellants 39,230 NEP propellant 36,700 Forward attitude control 1,265 Aft attitude control 1,265 Near Earth Departure Mass 50,030 Note: Any errors due to rounding Table 3. REACTIONN Baseline Spacecraft Cost Assessment: ROSETTA Cost Model. Hardware Cost Item Non-Recurring (DDT&E) Cost, $M-FY2003 Acquisition Cost, $M-FY2003 Nuclear reactor power system $ M $ M Propulsion $ M $58.56 M Thermal Control $30.00 M $0.52 M Main structure $70.00 M $0.10 M Data Processing $8.40 M $4.20 M Navigation Sensing/Control $2.00 M $0.60 M Telecom and Data $60.49 M $21.90 M Cost Summary Sub-total $1, M $ M Total Programmatic Costs (30%, and 10%) $ M $20.09 M Total Cost (without Margin) $1, M $ M Margin (+55%) $ M $ M Total Cost (with margin) $2, M $ M Total Cost-Development and Acquisition (with margin) $2, M Note: Any errors due to rounding 9

10 Data Processing 0.7% Telecom and Data Navigation Sensing/Control 4.9% 0.2% Main structure 5.7% Thermal Control 2.4% Propulsion 27.1% Nuclear reactor power system 59.0% Figure 8. Non-Recurring (DDT&E) Cost (Without Program Costs and Margin): ROSETTA Cost Model. B. Trade Studies Figures 9 through 11 reveal the results of several trade studies performed to examine sensitivity to input payload mass and Isp in terms of overall spacecraft lmetrics such as departure mass and reactor power. At various larger Isp levels, the reactor power required for a 1 MT payload plateaus around 0.5 and 1 MW. This could be indicative of the non-linear scaling effects of the reactor subsystem. For Isp values under 4,000 seconds, the difference in NEDM becomes more apparent for different time of flight (TOF) values. Payload generally has a linear relationship to near Earth departure mass for this concept. Near Earth Departure Mass, MT 160 Payload = 1.0 MT ,500 3,750 4,000 4,250 4,500 4,750 5,000 Isp, seconds TOF = 3 yrs TOF = 5 yrs TOF = 7 yrs TOF = 9 yrs TOF = 11 yrs TOF = 13 yrs TOF = 15 yrs Figure 9. Trade Study A: Isp Versus Near Earth Departure Mass (1 MT Payload). 10

11 Payload = 1.0 MT Reactor Power, MW ,500 3,750 4,000 4,250 4,500 4,750 5,000 Isp, seconds TOF = 3 yrs TOF = 5 yrs TOF = 7 yrs TOF = 9 yrs TOF = 11 yrs TOF = 13 yrs TOF = 15 yrs Figure 10. Trade Study B: Isp Versus Reactor Power (1 MT Payload) Isp = 4,050 seconds Near Earth Departure Mass, MT Payload, MT Reactor Power = 0.50 MW Reactor Power = 0.55 MW Reactor Power = 0.65 MW Reactor Power = 0.75 MW Reactor Power = 1.00 MW Reactor Power = 1.25 MW Figure 11. Trade Study C: Payload Versus Near Earth Departure Mass (Isp = 4,050 seconds). V. Detailed Cost Assessment As an adjunct to this study a more detailed parametric cost model was utilized and compared with the ROSETTA cost model. Similar sets of assumptions were utilized for systems integration and margin percentages. The NAFCOM 2004 tool utilized here allowed for more adjustment of complexity factors to reflect heritage technology from an assumed earlier JIMO mission. This additional modeling effort also allowed for calculation of operations cost for the mission. 11

12 A. NAFCOM 2004 with SOCM The NASA/Air Force Cost Model (NAFCOM) is a parametric cost-estimating tool based upon historical data of space projects. Outputs from the model include development and production costs down to a subsystem level. Included in the latest revision of the NAFCOM model is the NASA Space Operations Cost Model (SOCM). This tool is based on both parametric data and constructive approaches to operations at various NASA field centers. Costs are calculated for various portions of the mission. Utilization of NAFCOM 2004 tool resulted in a total development and acquisition cost relatively close to the ROSETTA cost model output (see Table 4). The acquisition costs were slightly less for the NAFCOM derived model, as was the overall cost. This is most likely due to the extra capability to adjust subsystem complexity within the NAFCOM 2004 model. These complexities were adjusted to reflect more design maturity in the REACTIONN concept given a previously assumed development effort for a JIMO mission (and thus a subsequent lower Technology Readiness Level or TRL for each subsystem technology). Table 4. REACTIONN Baseline Spacecraft Cost Assessment: NAFCOM 2004 Cost Model. Hardware Cost Item Non-Recurring (DDT&E) Cost, Acquisition Cost, $M-FY2003 $M-FY2003 All sub-systems $1, M $ M Cost Summary Sub-total $1, M $ M Systems Integration $ M $33.83 M Fee (+5%) $79.49 M $12.79 M Program Support (+10%) $ M $26.85 M Contingency (+15%) $ M $44.30 M Total Cost (with margin) $2, M $ M Total Cost-Development and Acquisition (with margin) $2, M Note: Any errors due to rounding SOCM was used to develop a scenario of operations for the REACTIONN mission. A ten to fifteen year cruise portion along with a total on-orbit portion of one year was assumed. The SOCM also has the capability to estimate science instrument operations costs. Based upon a suite of sample science instruments including radars, SOCM estimated a total operations cost of $106.4 M (FY2003) which consists of $77.7 M for flight operations, $13.6 M for navigation and tracking, and $15.1 M for science operations. VI. Conclusion A concept assessment is presented of a follow-on mission utilizing the same nuclear electric propulsion technology path envisioned by the Jupiter Icy Moons Orbiter (JIMO) program but to a different destination, namely to the outer solar system including Pluto, Charon, and objects in the Kuiper Belt. The Rapid Electric Acceleration Coupling ION and Nuclear (REACTIONN) concept utilizes currently planned nuclear reactors along with high power ion thrusters. A Reduced Order Simulation for Evaluation of Technologies and Transportation Architectures (ROSETTA) model was developed which included trajectory, performance, weights, power, sizing, and cost disciplines. The baseline REACTIONN concept has a near Earth departure mass of MT (10.8 MT dry mass) for a total V of 49.7 km/s and an anticipated development and acquisition cost between $2.45 B and $2.82 B (FY2003). The resultant spacecraft is relatively large and will require in-space assembly of constituent parts in Low Earth Orbit (LEO).The spacecraft s subsystems are generally small enough to be launched individually or in combination with other subsystems. Trade studies indicate that for lower payload classes (under 1 MT), larger reactor power does not necessarily relate to smaller near Earth departure mass. At such low payloads the power reactor seems to be oversized for the payload required. This effect is most noticeable for power levels approaching 1 MW and beyond for the payload range (0.25-2MT) in question. Acknowledgments The authors would like to thank NASA Marshall Space Flight Center (MSFC) and specifically Norm Brown for their contribution in providing financial and technical support for this assessment. Appreciation is also extended to 12

13 Tara Polsgrove at NASA MSFC for specific assistance in assessment of various trajectories for this mission. Specific acknowledgements are extended to Andy Gamble, also at NASA MSFC, who provided guidance on the mission down-selection process. References 1 Anon, A., NASA - Space Science - Project Prometheus, NASA Project Prometheus Homepage [html document], URL: [cited 14 July 2004]. 2 Anon, A., Jupiter Icy Moons Orbiter Fact Sheet, NASA JIMO Homepage [PDF document], URL: [cited 14 July 2004]. 3 Casani, J., Jupiter Icy Moons Orbiter: Mission Characteristics Overview to the Forum on Concepts and Approaches for Jupiter Icy Moons Orbiter, Lunar and Planetary Institute [PDF document], URL: [cited 14 July 2004]. 4 Noca, M., Polk, J. E., and Lenard, R. L., Evolutionary Strategy for the Use of Nuclear Electric Propulsion in Planetary Exploration, Proceedings of the Space Technology and Applications International Forum (STAIF), edited by M. S. El-Genk, Vol. 1, American Institute of Physics, New York, Humble, R. W., Henry, G. N., and Larson, W. J., (ed.), Space Propulsion Analysis and Design, Space Technology Series, McGraw-Hill, New York,