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Trade study of nuclear space power and propulsion system architectures for advanced interplanetary travel

University of Florida Institutional Repository

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1 TRADE STUDY OF NUCLEAR SPACE POWER AND PROPULSION SYSTEM ARCHITECTURES FOR ADVANCED INTERPLANETARY TRAVEL By JACLYN CICHON A THESIS PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLOR IDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF SCIENCE UNIVERSITY OF FLORIDA 2007

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2 2007 Jaclyn Cichon

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3 I would like to dedicate th is work to my parents.

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4 ACKNOWLEDGMENTS The material presented in this thesis is th e culmination of both hard work and dedicated studies in the fields of aerospac e and nuclear engineering. Presenting this thesis in its final form has only been done through the support and assistance of many people. I would first like to thank my parents for their support throughout my life, and most notably their encouragement of my intellectual pursuit s. I would like to thank Dr. John-Paul Clarke, my UROP advisor at MIT who assured me that I would get through college and someday be a successful graduate student. I would like to thank Dr. Samim Anghaie for convincing me to further my studies in graduate school and a llowing me the opportunity to pursue a Nuclear Engineering masters degree. I owe much gratit ude to my Pratt & Whitney mentor, Russ Joyner, who gave me the time and the tools I needed to begin and complete this thesis. Lastly, I would like to thank all of my friends and peers both at MIT and UF for thei r constant words of encouragement, especially those who truly believ ed I could someday become the President of NASA.

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5 TABLE OF CONTENTS page ACKNOWLEDGMENTS...............................................................................................................4 LIST OF TABLES................................................................................................................. ..........8 LIST OF FIGURES................................................................................................................ .......10 ABSTRACT....................................................................................................................... ............13 CHAPTER 1 INTRODUCTION..................................................................................................................14 Motivation..................................................................................................................... ..........14 Project Statement.............................................................................................................. ......15 Previous Work and Contributions..........................................................................................16 Study Overview................................................................................................................. .....19 Study Methodology.........................................................................................................19 2 MISSION PLANNING..........................................................................................................23 Astrodynamics.................................................................................................................. ......23 Orbital Motion.................................................................................................................23 Patched Conic Method....................................................................................................25 Mission Profiles............................................................................................................... .......30 Considerations for Mission Planning..............................................................................31 Mission trajectory.....................................................................................................31 Mission approach.....................................................................................................32 Mission Profile Selection................................................................................................34 Mission trajectory.....................................................................................................35 Mission approach.....................................................................................................38 Mission profile summary.........................................................................................39 Ephemeris Tools................................................................................................................ .....40 3 MODEL DEVELOPMENT....................................................................................................44 Space Reactor Background.....................................................................................................44 Reactor Configuration.....................................................................................................44 Reactor Fission Spectrum................................................................................................45 Nuclear Thermal Power and Propulsion.................................................................................46 History of Nuclear Thermal Rocket (NTR) Systems......................................................46 Fundamental Research.....................................................................................................47 Development of the NTR Model.....................................................................................50 Power........................................................................................................................52 Propulsion.................................................................................................................55

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6 Nuclear Electric Power & Propulsion.....................................................................................57 History of Electric Prop ulsion (EP) Systems..................................................................57 Fundamental Research and EP Description....................................................................59 Ion thrusters..............................................................................................................60 MPD thrusters..........................................................................................................61 Hall thrusters............................................................................................................61 Development of the Nuclear El ectric Propulsion (NEP) Model.....................................63 Power........................................................................................................................64 Propulsion.................................................................................................................66 Hybrid Power and Propulsion.................................................................................................72 Background..................................................................................................................... .72 Development of the Hybrid Model..................................................................................73 Power........................................................................................................................73 Propulsion.................................................................................................................74 Common Vehicle Subsystems................................................................................................75 TransHab....................................................................................................................... ..75 Spacecraft Functional Components.................................................................................79 Analysis Tools................................................................................................................. .......80 ModelCenter Software.....................................................................................................80 Incorporation of Models in to ModelCenter Framework.................................................81 Ephemeris models...........................................................................................................81 Vehicle architecture models.....................................................................................83 Mission architecture models.....................................................................................84 Validation of models................................................................................................85 4 TRADE STUDY RESULTS..................................................................................................96 Ephemeris Analysis............................................................................................................. ...96 NTR............................................................................................................................ .....97 Departure..................................................................................................................97 Duration....................................................................................................................98 NEP............................................................................................................................ ......99 Departure..................................................................................................................99 Duration..................................................................................................................100 Hybrid......................................................................................................................... ...101 Departure................................................................................................................101 Duration..................................................................................................................102 Vehicle Architecture Analysis..............................................................................................103 NTR............................................................................................................................ ...104 NEP............................................................................................................................ ....106 Hybrid......................................................................................................................... ...108 Tradespace Assessment........................................................................................................109 Methodology..................................................................................................................110 Mission architecture analysis.................................................................................110 Determining MOE scores.......................................................................................112 Tradespace Mission Architecture Results.....................................................................115 Tradespace MOE Results..............................................................................................117

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7 5 CONCLUSION.....................................................................................................................154 Study Contributions............................................................................................................ ..154 Areas of Future Work...........................................................................................................155 Final Comments................................................................................................................. ...156 APPENDIX A LIST OF ABBREVIATIONS...............................................................................................158 B TRADESPACE MISSION ARCHITECTURE RESULTS.................................................161 C MODELCENTER TEMPLATE AND FILEWRAPPER.....................................................175 LIST OF REFERENCES.............................................................................................................180 BIOGRAPHICAL SKETCH.......................................................................................................184

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8 LIST OF TABLES Table page 2-1 Design Reference Mission Categories...............................................................................43 3-1 NTR Architecture Systems................................................................................................88 3-2 Fuel Operating Parameters.................................................................................................88 3-3 NTR Reactor Mass Sizing.................................................................................................88 3-4 NEP Power Reactor Descriptions......................................................................................88 3-5 EP Thruster Characteristic Parameters..............................................................................88 3-6 TransHab Mass Breakdown...............................................................................................88 3-7 Common Subcomponent Mass Breakdown.......................................................................89 3-8 NTR Model Validation Mass Estimates............................................................................89 4-1 MOE4 TRL Score.............................................................................................................120 4-2 Tradespace MOE Scores..................................................................................................120 4-3 Tradespace Mission-Based Rankings..............................................................................121 4-4 Tradespace Planet-B ased and Overall Rankings.............................................................121 B-1 NTR Optimized Ephemeris Results.................................................................................161 B-2 NEP Optimized Ephemeris Results.................................................................................162 B-3 Hybrid Optimized Ephemeris Results.............................................................................163 B-4 NTR Graphite Tradespace Results...................................................................................164 B-5 NTR Composite Tradespace Results...............................................................................165 B-6 NTR Carbide Tradespace Results....................................................................................166 B-7 NTR CERMET-Unimodal Tradespace Results...............................................................167 B-8 NTR CERMET-Bimodal Tradespace Results.................................................................168 B-9 Hybrid Ion Tradespace Results........................................................................................169 B-10 Hybrid MPD Tradespace Results.....................................................................................170

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9 B-11 Hybrid Hall Tradespace Results......................................................................................171 B-12 NEP Ion Tradespace Results............................................................................................172 B-13 NEP MPD Tradespace Results........................................................................................173 B-14 NEP Hall Tradespace Results..........................................................................................174

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10 LIST OF FIGURES Figure page 1-1 Tradespace Matrix......................................................................................................... ....22 2-1 Escape Hyperbola.......................................................................................................... ....43 3-1 NTR Unimodal Vehicle Architecture Schematic..............................................................90 3-2 NTR Bimodal Vehicle Architecture Schematic.................................................................90 3-3 NEP Vehicle Architecture Schematic................................................................................91 3-4 NEP Reactor Mass vs. Power............................................................................................91 3-5 Power and Efficiency Flow Diagram.................................................................................92 3-6 PMAD Mass Sizing Equation............................................................................................92 3-7 Hybrid Vehicle Architecture Schematic............................................................................93 3-8 NEP Ephemeris Tool GUI.................................................................................................93 3-9 Darwin Optimization Tool GUI.........................................................................................94 3-10 ModelCenter Screenshot................................................................................................... .94 3-11 ModelCenter Mission Architecture GUI...........................................................................95 4-1 ModelCenter Parametric Tools........................................................................................122 4-2 NTR Delta-V vs. Departure Year....................................................................................122 4-3 NTR Mars Delta-V vs. Departure Year and Month.........................................................123 4-4 NTR Delta-V vs. Transfer Time......................................................................................124 4-5 NTR Delta-V vs. Stay Time.............................................................................................124 4-6 NTR Mars Delta-V vs. Transfer and Stay Time..............................................................125 4-7 NTR Jupiter Delta-V vs. Transfer and Stay Time...........................................................126 4-8 NTR Saturn Delta-V vs. Transfer and Stay Time............................................................127 4-9 NEP Delta-V vs. Departure Year.....................................................................................128 4-10 NEP Mars Delta-V vs. Departure Year and Month.........................................................129

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11 4-11 NEP Jupiter Delta-V vs. Departure Year and Month......................................................130 4-12 NEP Saturn Delta-V vs. Departure Year and Month.......................................................131 4-13 NEP Delta-V vs. Transfer Time.......................................................................................132 4-14 NEP Delta-V vs. Stay Time.............................................................................................132 4-15 NEP Mars Delta-V vs. Transfer and Stay Time..............................................................133 4-16 NEP Jupiter Delta-V vs Transfer and Stay Time............................................................134 4-17 NEP Saturn Delta-V vs. Transfer and Stay Time............................................................135 4-18 Hybrid Delta-V vs. Departure Year.................................................................................136 4-19 Hybrid Mars Delta-V vs Departure Year and Month.....................................................137 4-20 Hybrid Jupiter Delta-V vs Departure Year and Month...................................................138 4-21 Hybrid Saturn Delta-V vs Departure Year and Month...................................................139 4-22 Hybrid Delta-V vs. Transfer Time...................................................................................140 4-23 Hybrid Delta-V vs. Stay Time.........................................................................................140 4-24 Hybrid Mars Delta-V vs. Transfer and Stay Time...........................................................141 4-25 Hybrid Jupiter Delta-V vs. Transfer and Stay Time........................................................142 4-26 Hybrid Saturn Delta-V vs. Transfer and Stay Time........................................................143 4-27 NTR IMLEO vs. Delta-V................................................................................................144 4-28 NTR Mass Breakdown vs. Fuel Type..............................................................................144 4-29 NTR Burn Time vs. Fuel Type........................................................................................145 4-30 NTR Vehicle Mass Breakdown.......................................................................................145 4-31 NEP Power vs. Delta-V...................................................................................................146 4-32 NEP IMLEO vs. Delta-V.................................................................................................146 4-33 NEP Mass Breakdown vs. Thruster Type........................................................................147 4-34 NEP IMLEO vs. Burn Time and Delta-V........................................................................148 4-35 NEP Vehicle Mass Breakdown........................................................................................149

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12 4-36 Hybrid IMLEO vs. Delta-V.............................................................................................149 4-37 Hybrid Mass Breakdown vs. Thruster Type....................................................................150 4-38 Hybrid IMLEO vs. NEP and NTR Delta-V.....................................................................151 4-39 Hybrid Vehicle Mass Breakdown....................................................................................152 4-40 Mars Mission IMLEOs....................................................................................................152 4-41 Tradespace MOE Scor es for Mars Missions...................................................................153

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13 Abstract of Dissertation Pres ented to the Graduate School of the University of Florida in Partial Fulfillment of the Requirements for the Degree of Master of Science TRADE STUDY OF NUCLEAR SPACE POWER AND PROPULSION SYSTEM ARCHITECTURES FOR ADVANCED INTERPLANETARY TRAVEL By Jaclyn Cichon May 2007 Chair: Samim Anghaie Major: Nuclear Engineering Sciences Past mission analysis studies for interpla netary missions have typically focused on assessing one power and propulsion system applied to a singular mission. This has led to a lack of comparable results for different power a nd propulsion systems, and limited systematic procedures for mission planning. We attempted to address these faults through a comprehensive trade study on NTR, NEP, and Hybrid power a nd propulsion systems, grouped by NTR fuel type or electric thruster type into eleven configurations. Ephemeris codes were used to investigate mission planning and departure options, and vehicl e models were developed and then analyzed through parametric performance plots. Lastly, m easures of effectiveness were defined and used to assess a tradespace matrix containing all vehi cle configurations applied to twelve optimized reference missions. This assessment provided a simplified ranking of specific vehicle designs for each given interplanetary mission, along with a comparison of vehicle performance based on important figures of merit such as propellant mass and IMLEO.

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14 CHAPTER 1 INTRODUCTION Motivation The desire to travel though space and reach distant planets has no doubt existed for ages, yet the actual planning of interpla netary missions truly began in the United States approximately 60 years ago. This is when Werner Von Br aun, former aerospace engineer and National Aeronautics and Space Administra tion (NASA) administrator, c onducted the first engineering analysis of a manned mission to Mars. By the late 1960s Von Braun had become a proponent for nuclear thermal rocket-powered spacecraft, wh ile his Soviet contemporary Korolev believed nuclear electric propulsion was the best option to go to Mars. Cancellation of the Apollo program in the 1970s led to a sharp decline in nuclear propulsion and Mars exploration research.1 In 1988 President George H. Bush appointed NASA to create the Mars Office of Exploration as part of the Space Exploration Init iative, resulting in a resurgency in reports on manned missions to the moon and Mars.2 It was believed that the moon would be a stepping stone to Mars, as one report states: the Moon provides a unique database for life scie nce and operational verification in a reduced gravity environment, combined with the psychological rea lism of operations at a harsh extraterrestrial location.3 Although the goals sought by the Space Exploration Initiative never came to fruition, a renewed interest in moon and Mars explorati on was expressed by George W. Bush under the Moon Mars and Beyond initiative in 2002. NASA has indeed answered this call for a return to the moon and then to Mars, doing so at a pace commensurate with its appropriated funding and resources. The government agency, with the help of its industry partners, is currently in the initial phases of its Constellation project, which hopes to return man to the moon by 2014 and eventually replace the outdated and complex Sp ace Shuttle with the Or ion Crew Exploration

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15 Vehicle. Despite all the prep aration a return to the moon wo uld provide for a future Mars mission, one of the main differences would be that a vastly different prop ulsion system would be needed to make the much longer journey to Mars. In fact it has been noted by mission planners that the choice of a transportation system is the key trade in performance for Mars exploration.3 This is why it is paramount to invest now in Mars mission planning studies that assess the various power and propul sion architectures that have the capabilities that will be required to carry man to the surface of Mars and back. Project Statement Our study aimed to provide this necessary mission analysis and to do so from as broad a scope as possible. Past studi es have provided results for sp ecific missions using only one propulsion system, making it difficult to try and compare all possi ble propulsion systems for the same mission. Our study attempted to provide a comparison of the major power and propulsion system architectures considered for future manned interplanetary missions and to compare them on an even playing field. This involved elem ents of mission analys is, system modeling and optimization, and parametric analysis that culm inated in a large trad e study. Although the endgoal was to provide an answer to which power and propulsion system model best completes the attempted interplanetary missions using the aide of a tradespace matrix, analysis of non-missionspecific vehicle performance and inte rplanetary trajectories was also assessed. It was the hope of this author that the approach taken was broad enough in context to allow application to a myriad of future space mission profiles, yet specific enough to provide valid and useful results of vehicle performance. This mission analysis study prom ises a unique perspective on the subject matter given the combined systems engineering perspec tive and fundamentals of nuclear reactor design, astrodynamics, and rocket propulsion.

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16 Through investigation of previously comp leted Mars mission studies, a commonality found was that the primary objective of each study was to safely send and return humans to Mars.2 Although this objective was inherent as well to this st udy, a very important secondary objective was to provide a comparison of different vehicle architectures on an apples-to-apples level, such that bias towards or against a particul ar system was completely withheld. Each of the architecture models was pieced together follo wing the same methodology, with the same design constraints and without regard to particular mi ssion profiles. Thus, the optimization of vehicle performance for each mission profile produced resu lts that could provide an honest comparison of the different propulsion systems currently under consideration by both NASA and other members of the aerospace comm unity for future manned interplanetary missions. An extra feature of this study was that even more cha llenging missions to Satu rn and Jupiter were attempted so that a more in-depth understandi ng of the capabilities of the technologies being assessed could be gleaned. Previous Work and Contributions Many of the first formal studies comp aring advanced propulsion methods for interplanetary missions came about during the Space Exploration Initiative period, which focused on missions to the moon and then to Mars. At that point in time, it was desired to compare the conventional chemical propulsion system with that of a nuclear thermal or electric propulsion system. One report that was done by Boeing for pr esentation to NASA was the Space Transfer Concepts and Analysis for Expl oration Missions report. This comprehensive report attempted to compare the possible propulsion methods through various analyses, diagrams, and charts. Initial Mass in Low Eart h Orbit (IMLEO) charts were presented in many different fashions, but one of the most useful wa s a chart showing the optimal IMLEO versus trip time for all the opposition opportunities in the Earth-Mars synodical cycle.

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17 This chart showed these opportunity bands, w ith the lower boundary of each propulsion system option formed by the best year (2018), and the upper boundary by the most difficult year (2025), given all opportunities between the year s 2010 and 2025. The nuclear electric propulsion (NEP) system was the front-runner for the shorter mi ssion trips, yet it did not seem realistic to look at total mission trips to Ma rs below 400 days, especially given the radically high power requirement (120 MW) and IMLEO (1000 t) that would be needed for a fast NEP mission to Mars. Looking at the more reasonable time span of 400-600 day missions, the advanced NTR system was the front-runner, with the NEP system close behind, and lastly the rather impractical chemical system. It should be noted, however, th at if longer mission trip times were acceptable, the NEP system would present the lowest IM LEO option given all the systems assessed.4 In more recent studies in line with th e 2002 Moon, Mars, and Beyond initiative, the mission analysis performed has commonly fo cused on one specific power and propulsion system, going into a bit more detail on mission specifics, hardware, and mass and performance estimates. Three such studies were reviewed th at specifically focused on one main architecture that is currently considered feasible for advanc ed interplanetary travel These architectures include a Nuclear Thermal Rocket (NTR) system, NEP system, and a Hybrid system that uses one reactor to power both electric and thermal propulsion systems. A study entitled Nuclear Thermal Rocket Vehicle Design Options for Future NASA Missions to the Moon and Mars that was conducted by the NASA Lewis Research Center analyzed a 2010 human mission to Mars. This mission consisted of a 344-day outbound transfer, 1153-day stay at Mars, and a 180-day return. Th ree NTR burns were performed with an Earth swingby used for the last maneuver in order to reduce energy requirements. Two different reactors were considered, one based on a NERV A reactor, and the second utilizing ternary

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18 carbide fuels. The first had a specific impulse (I SP) of 900 (s), and the la tter an ISP of 960 (s). The reactors provided thermal power for the propulsion system and electric power for any spacecraft needs. This bimodal reactor de sign functioned by flowing hydrogen coolant through the core to heat the propellant, and then a heli um-xenon working fluid to remove heat for power conversion purposes. The results for the NTR sp acecraft included an IMLEO of 234 (t) for the NERVA-type reactor, and 207 (t) fo r the carbide reactor, given four reactors each operating at a thrust level of 15 (klbf).5 A team of engineers at NAS As Glenn Research Center conducted the NEP mission analysis study entitled, High Power MPD Nuclear Electric Propulsion (NEP) for Artificial Gravity HOPE Missions to Callisto. The desi gned mission was a 4.5-ye ar round-trip crewed journey to Jupiters moon Callisto, beginning in th e year 2041. The vehicle was to be proceeded by both a cargo and tanker vehicle, support a crew of six, provide artificia l gravity for most of the mission duration, and be refueled before leavi ng Callisto to return to Earth. The spacecraft was to be powered by a high temperature gas-cool ed fission reactor with a tungsten metal matrix CERMET fuel element. Electrical power was generated using a high power, closed-cycle Brayton heat engine. Hydrogen MPD thrusters provided propulsion with advanced performance parameters including 8000 (s) ISP and 64.5% efficien cy. A full mission analysis of this system revealed that the 120-day stay, 2.1-year transfer to and from Callisto would require an IMLEO of 262 (t). This value accounts for an outbound prop ellant mass of 74 (t), but does not include the return propellant load of 53 (t), which was carried to Jupiter aboard the tanker vehicle.6,7 The results of a mission analysis study performed to evaluate the feasibility for a mission to Mars using a Hybrid propulsion system were found in Mission to Mars Using Integrated Propulsion Concepts: Considerations, Opportunities, and Strategies. Integrated Propulsion

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19 Systems (IPS) is the term used to refer to a Hybrid system, one that uses both NTR and NEP systems for propulsive means. The mission scenar io in this study was a manned mission to Mars in which the spacecraft was assembled in Low Ea rth Orbit (LEO) at about a 300 (km) altitude, and then the reactor engine was turned on to be gin the mission. In this particular study, the nuclear and electric engines were both used to spiral out of the Earths atmosphere until the required Earth escape velocity was reached. The spacecraft then separated into two by means of a tether for on-orbit artificial gravity, and restarted both pro pulsion systems for simultaneous operation during the remaining voyage to Mars.8 The technology assumed in this study consis ted of a Rubbia Nuclear Rocket and MPD electric thrusters. In the Rubbia nuclear reactor, the heat exchange is essentially reversed from a typical NERVA-type reactor, with fission fragme nts from subcritical fissions of an isotope of Americium heating the coolant. This allows lower fuel operating temperatures, while enabling higher ISP values to be reached. Four superconductive Magneto-Plasma-Dynamic (MPD) thrusters were assumed for the electric propulsion system. It was assumed in the study that the ISP of the Rubbia systems was 3500 (s) and that of the MPD system was 56000 (s). The mass breakdown for this IPS spacecraft was found to be 378 (t) with an outbound propellant mass of 132 (t) and return propel lant load of 92 (t).8,9 Study Overview Study Methodology Our study attempted to go farther than prev ious studies, which primarily focused on specific power and propulsion systems applied to a singular mission, by as sessing three types of vehicles broken down into eleven different conf igurations over each of twelve possible mission scenarios. Basic methodology used to perform su ch an expansive mission trade study includes four major steps:

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20 1. Defining the general mission concept and propaga ting this to twelve different reference mission profiles 2. Creating ephemeris models a. Generating optimized trajecto ries for each mission profile b. Obtaining results of parame tric study of ephemerides 3. Developing models for the space po wer and propulsion architectures a. Running the vehicle models with mission trajectory requirements b. Performing parametric study of vehicle design attributes 4. Comparing all vehicle configurations in tradespace matrix a. Defining measures of effectivene ss (MOEs) and scoring algorithm b. Assessing all mission architecture blocks in tradespace matrix The reference missions, or mission profiles, us ed in this study we re built upon missions studied in the past. Though it was desired to encompass a broad spectrum of possible planetary destinations, stay times and transfer durations this spectrum was derived from missions in previously successful studies. Literature source s were also very useful in confidently making assumptions and ground rules as to trajectory ch oices, vehicle staging, an d planetary escape and capture mechanisms. The choice of mission prof iles, however, is inevitably up to the mission planner, with no real wrong or right manner in whic h to make the choices. Major considerations in designing the profiles were thus focuse d on designing missions that would be unbiased towards a particular vehicle, and to obtain a se t of missions that would adequately represent a subset of realistic mission scenarios for the tr adespace. It was essentially the design of the tradespace matrix and designation of appropriate MO Es that will, in the e nd, determine the utility of the final results.

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21 The next step in the trade study methodology wa s to set up ephemeris tools to model the trajectories of the chosen mission profiles. The te rm ephemeris refers to a table of the predicted positions of astronomical bodies such as the planets or moon, and by extension, the predicted positions of artificial satellites.10 The space vehicles modeled in this study can be seen as these artificial satellites, thus determining their traj ectories from Earth to the destination planets depended on complex calculations that took into c onsideration the orbits of planets at different dates in the future. Two ephemeris code s originally devel oped by NASA (IPREP and CHEBYTOP) were used in order to model these trajectories with ease. By passing the codes simple inputs based on the reference mission profile s, outputs were obtained that designated the propulsive requirements for the system architectur es. Optimization of the transfer time, stay time, and departure date inputs was done so as to minimize these energy requirements. Modeling the architectures of the space vehicles was the ne xt phase of the study. The main vehicle models created were for NTR, NEP, and Hybrid architectures. These models were then further broken-down into multiple configurations that varied by fuel type for the NTR system and thruster type for the NEP and Hybrid systems. The models were generated with Excel spreadsheets that calculated requirements fo r power, thrust, propellant load, and vehicle component sizing. The main consideration in these models was to accurately generate mass estimates for each system component in orde r to calculate IMLEO based on the trajectory requirements input into the vehicle models. Some of the components, such as reactor mass and propellant varied by mission and vehicle type, bu t other components such as the Transhab had constant mass estimates.11 Using these spreadsheet vehicle m odels alone, parametric plots were generated that characterized th e performance of each power and propulsion system architecture.

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22 The final phase of the trade study was actually assessing each of the vehicle architectures for each of the mission scenarios. The tradespace matrix, which is seen in Figure 1-1 was defined from the very beginning of the study. It was only after building the vehicle models and assessing vehicle performance for the given mission scenarios, that vehicle attributes could be defined. These vehicle attributes then become th e parameters used to assess the MOEs that were established for all of the missions. Finally, the MOEs were rolled up into a single score for each mission architecture so that the architectures coul d be ranked in the tradespace matrix. This ranking procedure was then used to come up w ith a final best answer for three distinct categories: each of the 12 missi on profiles, each of the three plan etary destinations, and for one overall interplanetary mission. FastSlowFastSlowFastSlowFastSlowFastSlowFastSlow NTR (Graphite Fuel; Thermal Reactor) NTR (Composite Fuel; Thermal Reactor) NTR (Carbide Fuel; Thermal Reactor) NTR (CERMET Fuel; Fast Reactor) BNTR (CERMET Fuel; Fast Reactor) HYBRID (CERMET; Fast; MPD Thruster) HYBRID (CERMET; Fast; Ion Thruster) HYBRID (CERMET; Fast;Hall Thruster) NEP ( MPD Thruster) NEP ( Ion Thruster) NEP ( Hall Thruster) Jupiter Conjunc.Oppos. MARS Conjunc.Oppos. Saturn Conjunc.Oppos. Figure 1-1. Tradespace Matrix

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23 CHAPTER 2 MISSION PLANNING Astrodynamics Essential to a successful mission analysis study is the understa nding of astrodynamics fundamentals and how a few specific parameters determine stringent mission requirements. The definition of astrodynamics in the context of this project is the study of the motion of rockets, missiles, and space vehicles, as determined from Newtons laws of motion and universal gravitation. More specifically, it can be seen that astrodynamics deals with the trajectories spacecrafts will assume on interplanetary transfers. The most basic parameters that stream down from astrodynamics calculations will, for example, define the best time for a spacecraft to leave Earth to travel to another destin ation and how much energy the vehi cle will need in order to get there. Due to the highly importa nt yet complex nature of this fi eld of study, a brief overview of astrodynamics topics is discussed herein that will introduce basi c terminology and concepts that were integral in the early steps of the mission analysis process. Orbital Motion Understanding the transfer methods for interpla netary travel assumes an understanding of orbiting objects. Orbital motion can be describe d by a family of curves called conic sections which represent the only paths possible for the orb it of one body about anot her. The three types of conic sections include open, cl osed, and a borderline case. Open conics are those in which the orbiting body repeats its path, and th ey only consist of circles and el lipses. The orbits of planets around the sun, and satellites around the Earth are a ll elliptical, with one real and one imaginary focus. The circular orbit is a special case of th e elliptical orbit in which the distance from the orbiting body is constant, and the two foci overlap to create one central point of focus. Using the energy equation for all conics found by Equation 2-1, where v refers to the orbital velocity, r to

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24 the distance from the orbiting body, to the gravitation parameter (GM), and a to the semimajor axis, the velocity of both an ellip tical and circular orbit may be found. a r v 2 22 (2-1) For a circular orbit with radi us always equal to a, the ve locity is found by Equation 2-2. circ circr 12 (2-2) The borderline case between open and closed co nic sections is for a parabolic orbit. Parabolic orbits are rare in na ture and an object tr aveling on one would c ontinue traveling to infinity until eventually coming to rest when a ll of its kinetic energy was exhausted. The orbital speed required to do just this, overcoming the gravitational field of th e orbiting body, is called the escape speed. From the energy equation for a conic section, and the statement that the energy will be equal to zero at a distance of infinity, the escap e velocity is given by Equation 23. resc 2 (2-3) This equation reveals that the required velocity to escape a planet will be less the farther the spacecraft is from the planet that it rotates. This is representative of why spacecraft designed to leave Earths atmosphere will first be launc hed to LEO to be assembled, and then escape the Earth with a much lower velocity requirement.12 It is actually the hyperbolic or bit, categorized as an open c onic section, which a spacecraft would use to escape from Earth. This orbit differs from the closed orbits because the traveling body does not retrace its path, and from the parabolic orbit because it will have some speed left

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25 over after traveling an infinite distance. A hyper bolic trajectory will only be achieved with a spacecraft that leaves its orbit at a velocity gr eater than the escape sp eed. This will result in some residual velocity left over when the spacecra ft reaches infinity. This residual speed is referred to as the hyper bolic excess speed, and is found by Equation 2-4, where is the hyperbolic excess speed, bo is the burnout speed, and bor is the orbiting radius at burnout. This equation was again found via the energy equation for conic sections, where the energy is constant between the end of the burnout and when the spacecraft reaches an infinite distance.12 bo bor 22 2 (2-4) In the context of interplanetary transfers from Earth, the reaching of infinity is assumed to be the same as reaching the end of the Earths sphere of influence. Although a body never completely escapes the gravitational field of the Earth, it can be assumed to be nearly zero at some distance from the surface. The sphere of influence (SOI) is said to end when the gravitational influence on the spacecraft is larger due to the sun than the Earth or other orbiting planet. For the Earth, this dist ance has been approximated as 145 times the radius of the Earth. With respect to the solar system this distance is ne gligible, but with respect to the Earth it is very distant. It is in fact so distant, that the veloc ity at the edge of the sphere of influence is assumed mathematically to be th e velocity at infinity.12,13 Patched Conic Method A method called the patched conic method comb ines elements of elliptical, circular, and hyperbolic orbits to describe the orbital motion a spacecraft assumes in order to complete an interplanetary transfer. This method allows one to ignore the influence of the sun while the spacecraft is within the Earths SOI, to switch to a heliocentric (sun-centered) frame outside of

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26 the SOI, and then to reverse this process upon arrival at the destina tion planet. The first step in the trajectory design will be to determine the heliocentric transfer orbit. An ideal minimum energy transfer to the des tination planet would be a simple ellipse, commonly termed a Hohmann transfer. This he liocentric transfer assumes that the departure and arrival planets are in circul ar orbits around the sun with velo city increments tangent to the planetary orbits, and that the velocity changes o ccur instantaneously. Thes e high thrust velocity changes are commonly referred to as delta-Vs, a nd constitute the relative velocities between the respective circular planetary velocities and the perigee and apogee velo city which define the transfer ellipse.14 For example, the Earth departure delt a-V would be the difference between the Earths velocity relative to the sun, and the spacecrafts velocity relative to the s un as it exits the SOI. The delta-V increments necessary to transf er from the departure planet (perigee) to the arrival planet (apogee) for a Hohmann transf er are given by Equati on 2-5 and Equation 2-6, where vp is the velocity increment at perigee of the departure planet, va is the velocity increment at apogee of the arrival planet, r1 is the periapsis distance of the Hohmann transfer, and r2 is the apoapsis distance of the Hohmann transfer.15 Note that for an interplanetary transfer the large center circle in the picture would represent the sun, and the planets would be represented by the point masses at the out er edges of the bl ue and red arrows. 1 22 1 2 1 r r r r vp (2-5) 2 1 1 22 1 r r r r va (2-6) It is estimated that using a Hohmann transfer for an interplanetary voyage to Mars would take approximately 260 days and would requir e an outgoing delta-V of 2.98 (km/s) for an

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27 instantaneous thrust acceleration.13 Although the Hohmann transfer is ideal due to its minimum energy solution, it is not always the most practical transfer method. For instance, once a spacecraft reached Mars af ter such a transfer, it would have to linger for nearly 6 months before it could return to Earth by means of another Hohm ann transfer. Thus less optimal transfers are usually taken despite the corresponding higher energy and delta-V requirements. The delta-V of all heliocentric transfers will be determined based upon the departure date and the travel time. The future date of depart ure allows for the determination of the relative positions of the launch planet and the target plan et at time of launch. The time of transfer will then determine where the destination planet will be in its orbit when the spacecraft gets there. The path that the spacecraft must follow to successfully intercept the destination planet will determine the energy of the orbit and the delta-V at both departure and arrival. This delta-V will always be the difference between the planet ary orbit around the sun and the heliocentric spacecraft speed given by Equation 2-7, where pr is the radius of the planets orbit, and t is the specific mechanical energy of the transfer. ) ( 2, t p helio vr (2-7) Optimal departure dates based upon the transfer time and stay time at the planet can be determined based on ephemeris data that tracks a planets synodic periods. This is the time it takes to reappear at the same point in the sky as observed from Earth and relative to the Sun. For instance, the synodic period for Mars is 2.135 year s, thus the best launch opportunities that would have minimum delta-V requirements would occur approximately every 780 days.12,13 Once the heliocentric delta-V is known, th e patched-conic method continues with the determination of the velocities relative to the planets at departure and arrival. The major

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28 assumption that is made at this point is that the heliocentric delta-V is equal to the speed of the spacecraft relative to the planet at the SOI. Using the previously introduced term in the hyperbolic excess speed equation (2-4), the hyper bolic excess speed is ob tained by Equation 2-8, where ic heliocentr is the heliocentric transfer delta-V, helio v, is the velocity of the spacecraft vehicle at escape or capture into the planet, and helio p is the velocity of the planet around the sun. helio p helio v ic heliocentr , (2-8) Since the hyperbolic escape velocity can now be determined, equation (2-4) can be rewritten as Equation 2-9 to determine the speed after injection burn, given the altitude at which the burn takes place. This speed is essentially th e perigee burn on the Eart h escape hyperbola. bo bor 22 (2-9) Since the elliptical or circular orbital velocity before the bu rn can be determined given the burn altitude, the actual delta-V experienced by the spacecraft will be that given by Equation 210, where planet bo, was the speed after burn and planet orbv, was the orbiting speed before burn, both relative to the planetary frame of reference. planet orb planet bo planetv, (2-10) It is important to note that although the delta-Vs calculated in the planetary and heliocentric reference frames may be similar, they cannot be assumed equal as significant values may result depending on the specific transfer being used. The delta -V in the planetary reference is that which will be used to assess the amount of propellant needed onboard the spacecraft,

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29 where the thrust maneuver is assumed to o ccur instantaneously at the burn altitude.13 A diagram depicting the parameters introdu ced for the escape hyperbola rela tive to the Earth can be found in Figure 2-1. The main difference that will result between the hyperbolic escape from Earth and the subsequent planetary capture, is that the transfer orbit was assume d tangent to the Earths orbit at departure, but the capture orbit will most likely cro ss the target orbit at some angle. This angle will be taken into consideration along with the target planets speed and the spacecraft heliocentric speed to determine the speed at the ta rgets SOI. The speed at the periapsis radius from the target planet will then be calculat ed based on conservation of angular momentum. Special attention must also be paid to the mini mum distance from the orbital plane in which the spacecraft can enter the planets atmosphere, as entering below this distance will result in collision with the target planet.12 Although the delta-V requirements are essentia lly derived from pre-determined physical laws of nature involving the orbits of plan ets around the Sun, to a space mission planner they define the cost of the mission. Once the delta-V values are determined from calculations given departure date and transfer time details for a sp ecific mission, these values can then be used to calculate how much propellant w ould be required given a specifi c type of propulsion system. The rocket equation is used to perform this calculation, enabling one to determine not only propellant requirements, but essentially the entire spacecraft mass at the la unch altitude. In the space business, mass means money, thus the connection can be made between delta-V requirements that stem from orbital mechanics and final monetary cost of any space mission.14 More specifics related to calculating propellant and system mass based on delta-V requirements

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30 will be found in Chapter 3, which discusses th e actual sizing of the spacecraft systems under consideration in this trade study. Mission Profiles The difficulty in defining a mission profile is obtaining the seemingly best overall approach given the many variables that play into mission design. Mi ssion characterization involves defining a mission concept which will desc ribe how the mission will work in practice, the mission operations which will detail how peopl e will operate and control the mission, and the mission architecture which links the mission concept to the major mission components.16 This study was focused on developing mission concepts and mission architectur es such that the various power and propulsion systems being an alyzed could be compared under different mission scenarios. Mission operations was not a major focus in this study since the trade was essentially made on the capabi lities of the power and propulsi on architectures to transport humans, and was not concerne d with human activities dur ing the transfer periods. The mission concept was essentially broken down into different mission profiles, each of which characterized the specific parameters of departure date, destina tion planet, outbound and return transfer times, and planet stay times (o r class). The use of planetary swing-bys, and the re-entry method at Earth are two other considera tions that typically go in to a mission concept but they were not characterized in the mission profil es since all profiles used the same methods in this regards. The term 'mission architecture' was used in th is study to refer to one of the three vehicle architectures (NTR, NEP, and Hybrid) being coupled to a specific mission profile. The motivation behind generating a la rge number of mission architect ures is to obtain a better understanding of the performance of each vehicle unde r different scenarios. Since it would be inefficient to attempt to model every type of s cenario for every vehicle architecture, the mission

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31 profiles were configured in such a way that mo st of the subset of possible desirable missions would fall into one of the mission categories. This was achieved by setting upper and lower limits on quantifiable mission profile parameters, and then collectively optimizing these parameters for each mission architecture. The main driver behind assessing vehicle arch itectures given different mission profiles was to see if one system was decisive ly superior to the rest for all profiles examined, or if instead different systems were the optimal choice for spec ific mission profiles. These results would be significant as they would highlight the importan ce of either focusing development efforts on one vehicle architecture for all possible future manned missions, or focusing first on the mission profile decision and then following w ith the appropriate vehicle design. Considerations for Mission Planning Mission trajectory The trajectory chosen for each mission profile is dependent on parameters that determine the launch window. These include the departure date, outbound and return transfer times and the stay time on the planet. The latter parameter takes on a significant role in trajectory analysis, often being divided into the two categories of short-stay oppos ition-class trajectories, or longstay conjunction class trajectories. The opposition-class mission is generally charac terized by short stay times on the order of 40-60 days, and round-trip Mars missions that range from 365-660 days. Most opposition-class Mars missions take advantage of a Venus swi ngby on the return traj ectory to Earth, though discussion of this maneuver is br iefly withheld. Other characteristics of this type of trajectory include large propulsive energy requirements an d the combination of both a short and a long transit leg. Clear disadvantages to this method that have been noted include high variances in

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32 energy requirements given departur e date, and large spikes in es cape and capture delta-Vs due to decreased transit time.2 The conjunction-class mission typically has stay-times between 400 and 600 days, and total Mars mission trip times on the order of 900 days. These trips may be the more highly desired of the two trajectory cl asses, since the stay time on the planet is much longer without a significantly longer transfer time. In fact, relatively s hort transfer times on the order of 200 days have been examined in previous studies for thes e mission trajectories. In addition, compared to the high energy requirements of the opposition-cla ss missions, the conjunction class missions typically represent minimum energy so lutions for given launch opportunities.2 The use of a Venus swingby periapsis burn has been incorporated in to methodologies for opposition-class Mars missions, as mentioned previously. This swingby maneuver provides a change in the spacecrafts heliocentric energy, whic h diminishes the delta-V requirement to enter back into Earths atmosphere. The maneuver t ypically requires a very small propulsive burst, but will result in an overall reduction in propellant requirements. Past studies of this trajectory type for Mars missions have indicated that adding a small propulsive maneuver during the swingby increases mission flexibilit y, and that a transfer at the periapsis of the Venus orbit is close to the optimum transfer point. Note that the periapsis burn can be in either the direct or retrograde directions, such that the relative velocity may incr ease or decrease, respectively.17 Mission approach Two approaches are also recognized for mi ssion design: an all-up mission and a split mission. All-up missions are those in which cargo a nd crew leave the orbit of the Earth at the same time. In a split-mission design, cargo is flow n to the destination planet first, and then is followed by the crewed vehicle. One benefit of th is design is that the car go can be sent on a lowenergy trajectory to the destinati on planet and assure th at supplies will be there to greet the crew

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33 when they arrive. In addition, this method will reduce the payload of the crewed vehicle, which already requires more energy due to the need to be sent on a faster trajectory to reduce time in the radiation-filled vacuum of space. On-orbit asse mbly is a possibility for either approach, and allows for different vehicle components to be launched into LEO on separate heavy lift launch vehicles (HLLV). The assembly could ut ilize automated rendezvous and dock between subcomponents, or the use of a space station or other docking facility.2 The capture methods at both the destination plan et and following return to Earth are also a consideration in the mission approach. Both pr opulsive capture and aerocapture methods have been used in past studies for capture into th e destination planet or moon. Propulsive capture imposes a delta-V requirement on the propulsi on system, thus necessitating a propellant requirement for the propulsive burn. The aer ocapture method instead uses the planets atmosphere to slow down the vehicle. This f light maneuver uses the friction from the dense atmosphere of the destination planet or moon to slow down the spacecraft, thereby transferring the energy from the high spacecraft speed into he at. This method thus saves on propellant load, yet requires advanced heat sh ielding to protect the craft.7 An orbit elevator has also been proposed as a method by which the spacecraft can move in order to reach various orbits once it is already in its destination orbit, wi thout the use of a propulsion system.6 Propulsive capture or di rect Earth re-entry are the two methods suggested for return to Earth. The propulsive capture, as its name suggests, uses a propul sive burn to capture into an orbit around Earth. The other me thod, direct Earth re-entry, assu mes that the spacecraft skims the desired Earth orbit, without actually capturing in to it through any propulsive means. The Earth Crew Return Vehicle (ECRV) would then be released from the spacecraft allowing the crew to descend back to Earth Apollo-style, with the remaining spacecraft hardware being

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34 released into space.18 One of the considerations then in choosing between the methods is whether or not the interplanetary spacecraft should be expendable. Mission Profile Selection A primary objective in the selection of sp ecific mission profiles for this study was to choose missions that represent, as much as possibl e, the entire subset of interplanetary missions that may be desirable in the future. This wa s accomplished by incorporating different planetary destinations, using the same transfer and capture methodologies for all missions, and by allocating adequate ranges for the mission prof ile parameters. Through optimization of all mission profile quantifiable parameters, singular data points were generated for each profile, including an exact date of depa rture, number of outbound and retu rn transfer days, and specific number of stay-time days on the planet. Thus although seemingly specific missions will be run for the vehicle architectures, each mission will re present a larger subset of possible missions and will reveal the best performance data for each transportation system. The decisions that were made in the mission design process were influenced by two major sources. The first very influential source wa s that of the set of literature found on past interplanetary studies. Differences between methodologies arose, such as assuming delta-V values for a mission versus using ephemeris codes to solve for them based on a specified launch window. These literature sources also c ontained information on NASA DRMs that was particularly useful in the mi ssion planning process since they provided a comprehensive outline of various mission concepts. The second source that drove mission profile design was that of the ephemeris tools themselves. Since a low-thrust and high-thrust ephemeris tool were both being used, and each tool had some unique capabilities and disabilities, constraints were set from the very beginning on possible mission platforms. For example, either the propulsive capture or direct Earth re-

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35 entry methods could be designate d for the high-thrust ephemeris tool, yet the low-thrust tool assumes a capture into Earth orbit, thus only the propulsive capture opt ion could be used if commonality between mission profiles was to be maintained. Mission trajectory In order to increase the breadth of the mission profiles, and thereby increase the applicability of this study to future research interests, three planetary destinations were attempted, each with two different stay times and two transfer times. Many advanced propulsion mission studies have been done using Mars as the planetary destination. In terms of present-day space policy and thought, this is directly in-lin e with the Moon, Mars, and Beyond initiative proposed by George W. Bush in 2002. But ther e have also been various studies on mission analysis for farther-out missions with destinations including Jupiters moon Callisto, Plutos moon Charon, and objects in the Kuiper Belt.18,19 For this study, it was decided that Mars would be a logical choice for a planetary destination, and that Jupiter and Sa turn would also be investigated in order to test the vehicle architectures against higher-energy requirement s. The spacecraft orbit around Mars was a 250 (km) by 33793 (km) elliptical orbit, which is comparable to 1 solar day, and was taken from Borowskis Mars mission study ground rules and assumptions list.18 The orbits around Jupiter and Saturn will include some basic assumptions. The spacecraft itself will be assumed to be orbiting Jupiters moon Ganymede, and Saturns moon Titan, as the spacecraft would not easily find a safe orbit around the planet itself. Howe ver, the ephemeris tool input will include the planet as the destination orbiting body and the or biting distance will be that of the distance from the planet to the orbiting moon. This amounts to the spacecraft seemingl y taking on the orbit of the moon around the planet, although it will actually be in a low orbit around the moon, with the spacecraft and moon both being in orbit around th e planet. Although some studies model the

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36 three orbit insertion components (propulsive capture at radius of planets moon, plane change into moons orbit, and propulsive capture into circular orbit around moon) the results of this study are only able to account for the first component, and assume the other two components negligible.19 The orbits that will be used in this study include a circular orbit at 998600 (km) from Jupiter, and a 1162000 (km) orbit around Saturn, which again are synonymous with the orbit of the selected moons around their respective planets. Since both short and long stay times may be desired for future interplanetary missions depending on the operational nature of the mission, both conjunction and opposition-class missions were used for the mission profiles. A range was applied to the stay time so that an optimization of the stay time could be found as part of the overall effort to find the minimum delta-V requirements for each mission. It was decided that the opposition-class missions would have a stay time range of 40-60 days, and conj unction-class missions would have a range of 400600 days, as these values are c onsidered standard among mission analysis literature sources.5,12 Note once again that a singular data point will be found for th e stay time after optimization, despite the fact that the real CTV would actua lly be designed with sufficient propellant for a somewhat longer/shorter stay to account for a safety margin. Although high and low energy missions are already broken down as opposition and conjunction-class in the mission prof iles, it was decided that short and long transfer times should also be included due to their own effect on energy requirements and safety concerns. Four of the central issues for mission profile selection are said to be crew radiation exposure, crew time spent in zero g, the component of mission risk that increases with mission duration, and added cost of shortening trip time.20 Each of the first three of these components condones shortening transfer times, though the last highli ghts the great expense, both in fuel and in dollars, of shorter

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37 missions. It thus seems reasonable to want to analyze transfer times on both the low and high ends of the spectrum. Having two data-points for transfer time is al so practical from an energy consideration perspective. For instance, sin ce a decrease in transfer time is generally associated with an increase in delta-V requirement, the optimized point could likely be that of the longest transfer time. If this transfer time is longer than th at considered acceptable by future mission planners, the dataset for this particular DRM may be deemed useless. It was thus found both fruitful and necessary to include both short and long transfer categories in th e breakdown of design reference missions. The transfer time ranges that were determ ined for this studys mission profiles were decided by using past studies to determine t ypical transfer times and by trying to make simplifying assumptions. The first assumption that this study made was to assume that the outbound and return transf er times were equal. This was done to minimize time spent on mission optimization using the ephemeris tools. Secondly, since the transfer times were to be broken down into two groups, it was decided that the Mars transfer ranges would be 100-200 days for the short transfer range and 200-300 days for the long tr ansfer range, allowing for an even split of 100 days for each. In a study that compared 21 different crew and cargo mission studies, it was found that outbound transfer times ra nged from 80 days to 335 days, with return transfers being the same or somewhat quicker.2 It was thus felt adequate that the profiles covered the 100-300 day transfer range, which eq uates to nearly 80% of the range covered by numerous other studies. In regards to the Jupiter missi on, Ehricke has been cited as using a 640-day transfer for a Europa mission, but many sources show a more practical 1000-day transfer to another of

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38 Jupiters moons, Callisto.9,19 Using the 1000-day transfer as a mid-point, the short transfer range was decided to be 800-1000 days, while the l ong transfer range would be 1000-1200 days. These categories are applied to both the Jupiter and Saturn mission as no sources were found for crewed mission profiles to Saturn. In addition, research that was done previous to this study showed similar trends for delta-V vs. trip time for Jupiter and Saturn destinations. Mission approach The current study serves to assess multiple vehicle architectures under a spectrum of mission profiles for the crewed vehicle component of a split-mission architecture. It is assumed that an unmanned cargo vehicle departs Earth or bit on a minimum energy, one-way trip to the destination planet, carrying scien ce payloads and any other payloads that would be necessary for planetary exploration, but not ne cessary for the outbound and return trips on the CTV. Proposed additional functions of the pre-deployed cargo ve hicle noted in the lite rature include deploying an unmanned semiautonomous rover to explore the landing site region prior to crew arrival, and deploying navigation beacons to assist in landing the primary cargo payload.21 Beyond these general assumptions about the cargo vehicle, no asp ects of it are modeled or considered in this mission analysis study. The mission concept used herein assumed that the hardware components of the crewed vehicle were assembled in LEO prior to depart ure. The components would first have to be launched from Earth on HLLVs, and then assembled by means of an automated rendezvous and dock between elements, or a more complex me thod involving the space station or other construction facility.2 After assembly, the crew would be transferred from the International Space Station (ISS) or a similar post to the CTV. Thus missions assessed in this study began from an operational sense at the point were th e assembled CTV was ready to leave LEO. In

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39 staying consistent with orbital parameters found in Borowskis studies, the actual LEO altitude the vehicle was assumed to be in wa s a circular Earth orbit of 407 (km).18 Certain aspects of the mission concept that we re kept constant fo r all mission profiles included major CTV components, and planetary a nd Earth capture methods. It was determined that the CTV would only carry the mass needed for a full two-way trip to and from the destination planet. Although this did require carrying both outbound and return propellant from LEO, the propellant tanks could be dropped upo n completion of major burn segments. The propulsive capture method was used for all capture s, both at the destination and upon return to Earth to accommodate the capabilities of the ephe meris tools and still maintain the means for an apples-to-apples comparison between architect ures. The Earth orbit insertion maneuver occurred at the perigee of a 500 (km) by 71165 (km) orbit, again in accordance with Borowskis standard ground rules. The mission analysis did not consider anything past this point in the mission, though it may be assumed that the crew woul d depart the ship at this point and perform a r-eentry into Earths atmosphere in the ECRV capsule.19 Mission profile summary To provide a condensed account of the mission profile assumptions to be used for the reference missions analyzed in this study, a list of the basic mission trajectory and mission approach ground rules is given below. Delta-V requirements for each mission were determined through an optimization methodology using high-thrust and low-thrust ephemeris tools Outbound and inbound transfer durations were as sumed equalMars departure dates ranged between 2030-2035; Jupiter and Satu rn dates ranged from 2040-2045 The following orbits were used for Mars, Jupiter, and Saturn: 250 (km) by 33793 (km) elliptical orbit at Mars 998600 (km) circular orbit at Jupite r (assume orbiting moon Ganymede) 1162000 (km) circular orbit at Sa turn (assume orbiting Titan)

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40 A split-mission scenario was used for all missions; only CTV will be modeled Earth Orbit Rendezvous and Dock Vehicle A ssembly at 407 (km) o ccurred after HLLVs transport spacecraft components to orbit All CTV propellant for round-trip mission was carried onboard vehicle Propellant tanks were dropped after completion of major burn segments Propulsive capture method used to capture CTV into planetary orbit Earth Orbit Insertion of CTV occurred at perihelion of 500 (km) by 71165 (km) orbit TransHab was sized for a crew of 6, with c onsumables accounting for the total duration of mission The optimization of the trajectories using the ephemeris tools incorporated the transfer and stay time ranges discussed in the Mission traject ory section. These ranges and their associated design reference missions are given in Table 21. The first three categories of Planet Destination, Mission Class, and Transfer Type, all characterize the 12 basic mission profiles. The outbound and return transfer duration and stay time are both given as a range in days, and provide the bounds for the transf er type and mission cl ass, respectively. Ephemeris Tools In recognition of the high-thrust NTR ar chitecture and low-thrust Hybrid and NEP architectures modeled in this study, the IPRE P and CHEYBYTOP epheme ris tools were both needed in order to model the interplanetary transfers for the missions previously discussed. IPREP (Interplanetary PREProcessor) is a ra pid grid-search optimizer for launch and arrival windows, delta-V, and mass originally created by Martin Ma rietta Astronautics.22 It is a commonly used tool for estimation of high-thrust trajectories, thus it was the logical choice to model the trajectories for the NTR architecture in this study. Inputs to the program generally include the order of the planets to be encountered, maneuvers to be performed at each encounter, and the time-of-flight window for each mission se gment. IPREP then calculates the delta-V

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41 energy requirements using a patched-conic technique that assumes the planets to be point-masses and finds the position and velocity of each planet from one of the ephemerides. A transfer orbit for each leg of the trajectory is then found by solving Lamberts problem.23 CHEBYTOP (Chebyshev Trajectory Optimizatio n Program) is a program built in the late 1960s by the Boeing Company, which provides a tw o-body, sun-centered, low-thrust trajectory optimization and analysis. Its capabilities extend to preliminary mission feasibility studies such as the one undertaken. Its strengths include requiring a small number of simple inputs and having fast run times for quick interplane tary mission modeling. CHEBYTOP uses the CHEBYCHEV Optimization Method, which is a seri es of approximations to the control problem that breaks it down into a group of classical calculus optimizations.24,25 The algorithms used are only relevant to spacecraft with very low thrust-to-weight (T/W) ratios, which makes it an ideal modeler for NEP systems. This code has been used in numerous studies done in the past on lowthrust trajectory systems such as solar electr ic propulsion (SEP) and NEP for Mars missions and for other planetary destinations. NASA Marsha ll Space Flight Center (MSFC), for example, used the code to generate thrust-to-weight (T/W ) versus delta-V curves for an NEP mission to Pluto.25 Besides its use for the NEP architecture in this study, it was also used for the Hybrid architecture due to its capability to model high -thrust NTR burns at escape and capture, with an NEP mid-trajectory low-thrust burn. The transfer methodology used by each of th e ephemeris models developed with these codes was slightly different depending on wh ich propulsion system was being modeled. The IPREP code used to model the high-thrust NT R propulsion systems assu mes one burn at Earth Escape, a second burn at Planetary Capture, a thir d burn at Planetary Escape, and finally a fourth

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42 burn at Earth Capture. These burns can be thought of as in stantaneous burns for modeling purposes, as is commonly done in traditional delta-V calculations for high-thrust orbit transfers. The NEP system that was modeled with CHEB YTOP used a spiral trajectory out of Earths orbit and a spiral capture into orbit at the de stination planet. This is due to the fact that a low-thrust engine cannot achieve the necessary es cape speeds in a very short amount of time as the NTR systems do. Note that the transfer time designated by the user in the IPREP input file did not account for the number of days needed to spiral out of or into a planetary orbit, but only accounted for the days on the in terplanetary trajectories. The Hybrid system, which was also modeled with CHEBYTOP, assumed an NTR burn to escape Earths orbit, an NEP burn for the majority of the transfer duration, and an NTR burn to capture into the destination planet s orbit. This same method wa s then used in reverse to escape from the planet and return to Earth. Even t hough the Hybrid and NEP systems used the same ephemeris tool, they were indeed modeled in separate ways. Due to features inherent to both the CHEBYTOP and IPREP ephemeris codes and the ma nner in which the scripts were set up for the ephemeris models generated in this study, the de lta-V requirements given by the output files took into account the different trajecto ry methods used by each vehicle.

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43 Table 2-1. Design Refere nce Mission Categories Planet Destination Mission Class Transfer Ty peOut/Ret Tran. (days)Stay Time (days) Fast 100-200 40-60 Oppos. Slow 200-300 40-60 Fast 100-200 400-600 Mars Conjunc. Slow 200-300 400-600 Fast 800-1000 40-60 Oppos. Slow 1000-1200 40-60 Fast 800-1000 400-600 Jupiter Conjunc. Slow 1000-1200 400-600 Fast 800-1000 40-60 Oppos. Slow 1000-1200 40-60 Fast 800-1000 400-600 Saturn Conjunc. Slow 1000-1200 400-600 Figure 2-1. Escape Hyperbola

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44 CHAPTER 3 MODEL DEVELOPMENT Space Reactor Background Reactor Configuration All reactors rely on the principle of thermal energy production from the fission process of a fissionable atom such as 235U. This energy production results from the conversion of the kinetic energy of fission fragments and neutrons to heat after slowing down from collisions and interactions with other atoms. This thermal en ergy is then transferred to a coolant that flows through the reactor core. Space reac tors will typically use a lightweight gas such as hydrogen as both the coolant that extracts heat from the reactor, and in the case of a nuclear thermal rocket, the propellant that immediately th ereafter is shot out of the ro cket nozzle to create momentum.26 Propellants with low molecular weights are mo st effective for thermal propulsion as they produce the highest specific impulse.27 Space reactors that are sp ecifically used for propulsion are known for having high specific impulse and high thrust levels providing a clear advantage over alternatives such as chemical propulsion. The actual configuration of a sp ace NTR system is similar to that of a chemical system, except for the reactor heat source. The hardware consists of a reactor, propellant tank, radiation shielding, a feed system, and a nozzle. The major reactor com ponents consist of the radial reflector, reactor pressure vessel, moderator, fu el-element assembly, and control drums. The reflector surrounds the outside of the core and func tions to reflect neutrons produced in the chain reaction back into the core, helping to maintain a controlled chain reactio n. The pressure vessel is needed in order to maintain reactor pressure and must be made of an aluminum or composite material that will withstand th e high radiation, heat flux, and pr essures from the reactor. A moderator material is used in a thermal reactor to slow the neutrons produced from a nuclear-

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45 fission reaction to energies in which they are more apt to undergo another fission. The fuelelement assembly contains the ac tual heat-producing ur anium fuel, along with the flow channels for the coolant. Control rods ar e also found in the core and serve to absorb neutrons to decrease neutron population and maintain the ability to co ntrol the reaction rate or even shut down the reactor.26 Reactor Fission Spectrum A major choice in overall reactor configurati on concerns the type of fission spectrum it will operate under. It may operate using fast neutrons produced by fissi on reactions (a fast reactor), or neutrons slowed to thermal energi es and thus more likely to produce subsequent fission reactions (a thermal reactor). A fast reactor functions based on a chain of reactions propagated by high-energy fission ne utrons. The low probability of fast neutrons producing fission reactions results in a large fuel requirem ent for this type of reactor. However, highly concentrated fuel will allow a compact reactor, ty pically of smaller size and mass compared to a thermal reactor. The probability of neutron ca pture and consequent fission is much higher for thermal energy neutrons that have been slowed by interaction with moderator materials than for fast neutrons. Although thermal reactors are typi cally larger than fast reactors, they require much less reactor fuel than fast reactors and also a less complicated fuel element and core design.28 The reason that both fast and thermal reactors have been considered for space reactors is that they are inherently better equipped for different types of missions. The fast reactors are preferred for long-life, low power operations since the high fissile loading allows high total energy operation. Missions that re quire large bursts of power but small total lifetime energy may benefit more from a thermal reactor. The re latively long neutron lifetime and large delayed neutron fraction found in thermal reactors would he lp maintain precise control of burst power.27

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46 Nuclear Thermal Power and Propulsion History of Nuclear Thermal Rocket (NTR) Systems This history of the NTR engine began in 1953 when the ROVER program began at Los Alamos Scientific Laboratory in order to develop a reactor for th e operation of a nuclear rocket. The major reactor series that went through desi gn, build, and test phases during this program included KIWI, Phoebus, Peewee-1, and Nuclear Furnace-1.28 The KIWI reactor series holds the accolade of being the first NTR reactor built and tested. It allowed for advances to be made in the areas of instrumentation and control, fuel element design and fabrication, structural design, and testing techniques. The Phoebus reactors that followed had design specifications intended to meet the need of interplanetary propulsion systems, with special focus on manned missions to Mars. The major results from research on this reactor series included c ontrol of rocket parameters over a wide range of operating conditions, along with finding that large nozzles fo r NTR applications was feasible. The Peewee reactor series came next chronologically, and was meant to investigate performance characteristics of a smaller reactor. This was followed by the last stage of ROVER, the Nuclear Furnace series, which was designed to test advan ced fuel elements containing composite fuel.28 The end of the research-focused ROVER program led to the start of the Nuclear Engine for Rocket Vehicle Applications (NERVA) progr am, which focused more heavily on concept development. During this 11-year program the NRX reactor series was developed, incorporating the non-nu clear system components (propulsive components) into the reactor designs developed during ROVER. The NRX-XE engine was the main focus as vertical downward firing tests of the engine in a simula ted space vacuum were conducted. This allowed for investigation of the engine start-up and s hutdown characteristics, along with the resulting engine performance parameters. Over the cour se of this program, more than 20 NTR reactors

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47 were built and tested at the Nucl ear Rocket Development Station at Nevadas Nuclear Test Site. Although the NERVA program was canceled in Ja nuary of 1973 due to a change in national priorities, the most poignant outcome of the work done during ROVER/NERVA was the confidence that an NTR engine could be developed to meet the objectives of structural integrity, restart capability, predictability, control, and reliability.28 Fundamental Research Two of the primary areas of current research for NTR systems are reactor fuels and reactor operational capabilities. Research of space reactor fuels is of high importance due to the fact that the reactor is heating a coolant for propulsive pu rposes. Increasing the operating temperature of a space reactor fuel thus has implications for ove rall propulsive performance. Reactor operations are also being analyzed in terms of providing du al-mode functions instead of a single mode, the implications of which may have profound e ffects on overall spacecraft capabilities. Improvements in the area of fuels research have been consistent since the ROVER years, while dual-mode bimodal reactor configurations have only been studied in recent years. It is important to note that for both areas of research, however, no full-scal e hardware testing has been undertaken for nearly 35 years. The four main types of reactor fuels that ha ve undergone serious consideration for space application include graphite, com posite, and carbide fuels for ther mal reactors, and fast reactor CERMET fuel. Graphite fuel is the oldest and most mature, as it was studied and used during the ROVER/NERVA programs. The original ROVE R engine had rods 54 in ches in length, with a mixture of uranium, zirconium, a nd carbide in a graphite matrix.29 Both the composite and carbide fuels were developed based upon expe rience gained with graphite fuels, yet they only underwent minimal testing near the end of the ROVER/NERVA era. The composite fuel composition differed from the gra phite fuel in that it was a composite of the

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48 UC2 and ZrC used to make up the fuel pellet an d coat in the early generation coated-particle matrix graphite form. It was found that the ZrC coating increased both the lifetime and the integrity of the fuel as it protected the fuel fr om the hot hydrogen propellant. Carbide fuels, on the other hand, eliminated the protective carbi de coating required for matrix fuels. Improvements were made with carbide fuels sinc e it was found that the composite fuel coating severely limited the endurance and te mperature performance of the fuels. Since the termination of the ROVER/NERVA pr ograms, many advances in fuels research have been made. Some stem from improvement s in other related areas such as materials research, and include an incr ease in hydrogen turbopump effici ency, improvements in titanium pressure vessel manufacturing techniques, a nd improvements in nozzle cooling. The term applied to updated NERVA NTR en gines that use these latest fu el technologies available is NERVA Derivative Reactors (NDR). NDRs typically have carbonbased matrix fuel elements with graphite moderator and ZrH moderator slee ves in the support struct ure. Typical chamber temperatures for NDR graphite, composite and carbide fuels are 2500, 2700, and 3100 (K) respectively, while ISP values are 885, 921, and 1020 (s) respectively.28 Bimodal reactor designs, which have been under development only more recently, have consistently been designed based on the use of CERMET fuel. This is primarily a result of both CERMET fuels and bimodal designs being based upon fast fission reactors. CERMET fuel is termed according to its ceramic metallic form ulation, and is primarily configured of UO2 fuel encased in tungsten and tungsten-rhenium alloys.30 Properties of CERMET fuel include high strength, thermal conductivity, temperature capabili ty and burnup, in addition to giving reactors a long operating life and the ability to restart. It is important to note that use of CERMET fuel and fission reactors is not limited to bimodal reactor configurations.

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49 Bimodal reactors, named appropriately for th eir two modes of operation, have been kept behind their unimodal counterparts due to the signif icant costs associated with their testing and production. Their discussion, however, reaches back to the ROVER/NERVA program, during which the potential benefits of such reactors we re recognized. The basic reactor design used for that program was assessed for possible modifications that would allow electric power generation. In more recent years, the Air Force Phillips La boratory has conducted studies on bimodal reactor designs for possible military applications. The DOE Office of Nuclear Energy has since collaborated with the Phillips Laboratory in order to develop bimodal bus designs along with initial performance requirements.30 Bimodal reactors have increased complexity in both design and function due to their dualmode operation. During the power and propulsion modes, the reactor must operate under very different sets of conditions and must perform extremely different functions. In addition to these capabilities, a space reacto r is expected to have both high re liability and a long lifetime. The engine must function in cooperation with more hardware than a unimoda l engine, given that a power conversion unit and heat radiator are requir ed to produce electric power and get rid of waste heat. While the amount of electric pow er generated is the important performance parameter in power mode, the thru st-force and specific impulse are the driving parameters in the propulsion mode. In a typical bimodal reactor design, the core will consist of heat pipes and CERMET fuel with numerous propellant channels. In the propulsion mode, liquid hydrogen may be run through reactor components for cooling purposes, and will then run through the CERMET fuel element one time before expanding out the rocket no zzle. In the power mode, the heat pipes will serve as the energy transport medi um from the reactor fuel to the power conversion system. A

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50 small gap between the fuel elements and heat pipe s serves to allow for thermal radiation between them, creating the primary energy transfer mech anism for the power mode. A working fluid such as sodium or xenon then runs through the heat pipe, transferring the energy to the power conversion unit. A reactor design by the Phillips Laboratory us es 93% enriched CERMET fuel with finned heat pipes. The CERMET fuel elements are nine-s ided blocks with 52 axial propellant channels. They are 59.5% UO2 by volume and 40.5% tungsten. The heat pipes provide both energy transport and structural support, which relieves the need for tietubes. This reactor design was said to have a 10 (kW) electric power output cap ability with a 10-year lifetime, and produced 220 (N) of thrust with a specific impulse of 825 (s) for the propulsive mode.30 Development of the NTR Model The NTR architecture, which utilizes a nucl ear reactor for thermal propulsion, was one of the three main architectures studied in this proj ect. This architecture was broken down into five systems that all achieve NTR propulsion through diffe rent selections of fuel and reactor types. The five NTR systems that were assessed in the tradespace matrix are found in Table 3-3, which categorizes each system according to reactor mo de and energy spectrum along with fuel type. Although the NTR systems were broken down acco rding to characteri stics of the space reactor, it is important to unde rstand how the reactor functions w ithin the entire spacecraft. The term architecture itself was used in this study to describe not only the reactor, but also each of the subcomponents that makes up the entire CTV th at will carry humans to an interplanetary destination. Thus it should be noted that the NTR architectur es were all based upon the same basic principles and had almost identical vehicl e schematics. The exception to this generality within the NTR architecture group was the one syst em operating with a bimodal reactor. While unimodal reactors main functionality was to provide heat to the hydrogen coolant that is

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51 expelled through the rocket nozzle, the bimoda l reactor had a secondary coolant loop that provided electric power to the spacecraft. W ithin the group of unimodal reactor systems, the only real differences were in the physical makeup and operation of the reactor itself. The four unimodal reactor systems can all be described using the schematic found in Figure 3-1. The divisions of the entire CT V were broken-down into power, propulsion, and spacecraft categories. This schematic was in tended to provide an overview of how the components within these groups interacted, along with a very rough idea of where they were physically located respectively with in the space vehicle structure. The solid black lines depict flow of hydrogen propellant from the tanks all the way through its exit from the rocket nozzles. The dotted black lines depict tr ansfer of electricity from th e fuel cell power source to the components that require electric ity for functionality. The power and propulsion system boxes in the schematic are specific to the NTR unimodal systems, and will thus be discussed in the following section. The general spacecraft compon ents, indicated in yellow, will be discussed later in the text, as their mode ling was not dependent on the type of propulsion and power system architecture. The one NTR system that used a bimodal reac tor had a significantly altered architecture schematic as shown in Figure 3-2. As can be seen in the schematic, fuel cells utilized in the unimodal systems are no longer used for onboard spacecraft power needs. In addition, power conversion units, power management and distribu tion (PMAD) systems, and radiators were all added to accommodate the reactor power genera ted by the bimodal reactor, and a secondary propellant was added for power conversion purposes The solid black lines in the figure again indicate propellant flow, while the dashed lines indicate flow in the power distribution process.

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52 Power One of the ways in which an NTR reactor will differ from a purely NEP power reactor is that the required thermal power will be define d based on the desired temperature and mass flow of the propellant instead of the mission energy re quirements. The derivation of this power can be found through both Equation 3-1 and Equation 3-2, where NTR propm_ is the rocket mass flow in (kg/s), T is the thrust in (N), NTR ev_ is the exit velocity of the propellant in (m/s), prop sP_ is the thermal power in (MWt) for the propulsive mode, and 2 Ht is the temperature in (K). It is interesting to note that the thrust value, which is needed to determine th e power requirement, is a parameter that is chosen by the designer. In th is study it was chosen to be 15 (klbf), which is on the lower end of typically consider ed thrust ranges. Thus it can be reasoned that the power is determined by both the designer and th e properties of the fuel itself. NTR e NTR propv T m_ (3-1) ) 715417 5 018061 0 ( *2 _ H NTR prop prop st m P (3-2) After the power required of each type of reactor fuel core for a thrust of 15 (klbf) was calculated, the actual mass of the NTR power and propulsion system had to be determined. The basic components consisted of the reactor, pressure vessel, internal and external radiation shield, and propulsive hardware. The propulsive hardware was further broken down into the three main components of the nozzle, turbopump assembly, a nd nonnuclear support hardware such as lines, valves, actuators, and instrume ntation thrust structures. Th e data-points for these three propulsive hardware components were taken fr om an SAIC report for the SAIC ELES-NTR

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53 design. The mass estimates for the three com ponents were 421 (kg), 104 (kg), and 1264 (kg) respectively, which gave a tota l of 1789 (kg) to be added to the reactor and shielding masses.31 The mass of the reactor and pressure vessel we re estimated from SAIC plots, which gave mass as a function of power, pressure, and temp erature. Linear interpolation tools were developed in order to be able to confidently generate mass estimates based on the characteristic parameters of each fuel. The fuel parameter va lues that were used for the estimation of the reactor mass included chamber temperature, ISP, and calculated power, all of which are seen in Table 3-2. Although the interpolat ors required the three inputs of temperature, pressure, and power, the ISP was also needed for the calculation because the mass flow parameter that defined the power requirement was dependent on ISP. Als o, the pressure used for all cases was assumed constant at 1000 (psia). Once the reactor masses had been calculated, th e internal and external shield masses were calculated based on these values. It was decided to estimate the shielding masses by using ratios of shielding to reactor mass found from data-points in the literature. It was thus determined from data that characterized a 75 (k lbf) Westinghouse/NERVA design, that the internal shielding was to be 26.96% of total reactor mass, and ex ternal shielding 80.27% of reactor mass.28 Calculating the total NTR power and propulsion system mass was then accomplished using Equation 3-3, where shield Intf_ was set to 29.6% and shield Extf_ was 80.3%. ) ( *_ shield Ext shield Int reactor propul reactor NTRf f m m m m (3-3) The reactor and shielding masses for each type of fuel, along with the total NTR system mass for each fuel is given in Table 3-3. Note that although only one reactor was used to carry out power and propulsion requirement s, an extra reactor system e qual in weight to the primary reactor was carried on board for redundancy.

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54 A unimodal NTR architecture uses all of the h eat generated from the r eactor for heating of the rocket propellant. Any electric power n eeded by the spacecraft ha d to come from an alternate source. Due to proven performance on the Space Shuttle, fuel cells were chosen for electric power generation for the NTR unimodal sy stem. Fuel cells are self-contained generators that operate continuously without the need for s unlight. They work by converting the chemical energy of an oxidation reaction to el ectricity. An added benefit of their use is that they actually produce water as a byproduct, whic h can the be used for drinking.16 Two key parameters of fuel cells used in the calculation of their mass were the specific power of 275 (W/kg), and the lifetime estimate of 100 days. The equation used to calc ulate the mass of fuel cells based on electric power requirement and the mission duration is given by Equation 3-4, where eP is the electric power in (kWe). ) 100 / ( * ) 275 / 1 (_days mission P masse Cell Fuel (3-4) The bimodal reactor provided both the hydrogen propellant he ating and electric power requirement for the spacecraft. Thus, although the actual mass of the bimodal reactor was on average higher than that of a unimodal reactor, no extra mass was assumed for power generation. There was, however, mass associated with the po wer conversion that is required to convert the reactors thermal energy to electrical power. Th e three main components required to fulfill this function include the PMAD system, the Brayt on Conversion Unit, and the radiator. The equations defining these component masses can be found by Equation 3-5, Equation 3-6, and Equation 3-7, respectively, where spacecraft sP, is the thermal power requirement needed to satisfy spacecraft electrical power needs in (kW), S is thermal to electric conversion efficiency, and radSpMass is the specific mass of the radiator (kg/kW).

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55 e PMADP mass 635 (3-5) e CBCP mass 875 0 (3-6) rad S spacecraft s radSpMass P mass )) 1 )( ((, (3-7) These equations were derived from both li terature data-points and textbook material.16,32 Further details on these component s and their derivations is given in the NEP propulsion section, which focuses more predominantly on the power conversion process. Propulsion The NTR propulsion equations were developed based on the principles of the rocket equation, which take into account both delta-V requirements and the mass of the spacecraft. The ephemeris tool that was used to model th e NTR architecture assumed that there was an escape burn at Earth and the destin ation planet, and also a capture burn at the respective planets. These four high-thrust burns typically last on th e order of a few hours and are represented in the equations by four separate delta-V values. Th e four separate equations that were used to calculate the propellant mass at each of these bur ns are shown in Equation 3-8 through Equation 3-11, where the exit velocity is given by Equatio n 3-12, and the mass totals at the beginning of each of the four burn segments are gi ven by Equation 3-12 through Equation 3-16. ) 1 ( *_/ 1 1 1NTR ev v TOT NTR prope m m (3-8) ) 1 ( *_/ 2 2 2NTR ev v TOT NTR prope m m (3-9) ) 1 ( *_/ 3 3 3NTR ev v TOT NTR prope m m (3-10) ) 1 ( *_/ 4 4 4NTR ev v TOT NTR prope m m (3-11) g Isp vNTR e*_ (3-12)

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56 4 3 2 1 / 1 prop spacecraft inert prop power TOTm m m m (3-13) 1 tan 4 3 2 / 2 k prop prop spacecraft inert prop power TOTm m m m m (3-14) 2 1 tan 4 3 / 3 k prop prop spacecraft inert prop power TOTm m m m m (3-15) 3 2 1 tan 4 / 4 k prop prop spacecraft inert prop power TOTm m m m m (3-16) The v terms in the propellant mass equations signify the delta-V requirements determined by the ephemeris tools, and the inert prop powerm_ /, spacecraftm, propm, and k propmtan values are masses in (kg) calculated by the NTR model. Another important parameter in these equations th at is a function of fuel parameters is the ext velocity NTR ev_. It is seen from the exit velocity equation above that this parameter is a function of ISP, increasing for fuels with highe r operating temperatures, and thus decreasing the required propellant load. Note also that th e delta-V values from th e ephemeris model output were used with the assumption of no gravity lo sses. Therefore, any losses that were not accounted for in the original code, were not a ccounted for elsewhere in the propulsion models. The weight of the spacecraft at LEO was also an important parameter in the propulsion model. This can be seen by the dependence of prop ellant load on spacecraft mass as well. All of the spacecraft components, total propellant, and f our propellant tanks were included in the first mass estimate (1 TOTm). At the end of each stage, however, the fuel that was spent, along with the mass of the tank, had to be deducted from the mass estimated before the burn. The mass represented in the equations above by the notation mspacecraft comprises that of general components including the communication and navi gation systems, along with the TransHab and

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57 ECRV. The mpower/prop_inert notation refers to any component s required for power or propulsion including the reactor itself, fuel ce lls, and power conversion components. Nuclear Electric Power & Propulsion History of Electric Propulsion (EP) Systems The history of electric propulsi on (EP) dates all the way back to 1903 when the visionaries in the scientific community began exploring its potential for spaceflight. In that year Tsiolkovsky released his article Investiga tion of Universal Space by Means of Reactive Devices, which contained the Tsiolkovsky rock et equation, notably the most fundamental mathematical expression in the field of space prop ulsion. It was only eight years later that he published an article that mentione d the idea of electric propulsion, saying it is possible that in time we may use electricity to produce a large velo city for the particles ejected from a rocket device. The combined knowledge of cathode ray t ube development at that time, along with his appreciation for the importance of rocket exhaust velocity to space propulsion led to Tsiolkovskys anticipation of the future of electric propulsion.33 Around the same period, Robert Goddard was al so investigating elec tric propulsion ideas as a natural result of his physics work on elec tricity and his passion for propulsion. His initial musings on the electrostatic accel eration of electrons eventually led to thoughts on the reaction of ions in an electrostatic accelerator, and the ne utralization of the charge d exhaust with a stream of oppositely charged particles. By 1917 Goddard had developed the worlds first documented electrostatic ion accelerator, with the inte ntion that it could be used for propulsion.33 It was Hermann Julius Oberth, however, who firs t declared to the technical community that EP was a serious and worthy consideration for future astronautics. His 1929 text Ways to Spaceflight devoted an entire chapter entitled The Electric Spaceship to the capabilities and future role of EP in propulsion and attitude contro l. This book is also what brought EP into the

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58 spotlight for science fiction writers. Unfortunatel y, after the 1930s this is where EP stayed as scientific advancements halted and efforts eventu ally switched to chemical rockets that were needed for the second World War.33 The following decades after the war did see a rene wal in interest in EP research. In 1949 British physicists L.R. Shepherd and A.V. Cleaver declared that ion rocketry was impractical, and in 1951 American astrophysicist Lyman Spit zer found that ion propulsion was perfectly feasible. Ernst Stuhlinge r wrote his classic text Ion Propulsion on the same subject in 1964. The 1960s in particular saw a growth in coordinate d research and development programs addressing EP technology, mostly due to the pervasive upswing of U.S. space ambitions. Following experimental flight tests in the 1970s came the fi rst commercial applications of EP in the 1980s as attitude control thrusters on commercial spacecr aft. The early 1990s saw electrothermal arcjet use on many communication satellites, and by 1998 el ectrostatic ion thrusters had been used on a planetary mission for NASA.33 Current activities in electric propulsi on are being conducted by NASA and throughout academia and include basic research all the way to flight demonstration. A cooperative agreement between NASA and MIT has allowed fo r research in the area of Hall thruster modeling to provide a theoretica l understanding of th e physical processes occurring in this electrostatic thruster type. In the flight ar ena, NASA currently has one active electrically propelled satellite called Earth Observing-1 (EO-1) that operates with pulsed plasma thrusters. In addition, NASA is currently developing the DAWN asteroid science mission that will use three NSTAR ion engines, and is providing a se t of colloid thrusters in support of the ESA Smart2 spacecraft as part of the New Millenni um Program. NASAs Project Prometheus, which was canceled due to budget constraints in 2004, aimed to combine space nuclear power with

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59 electric propulsion technologies to allow for sophisticated active /passive remote sensing, greater launch window flexibility, and increased science data rates. The technology areas that were under development for the projec ts Jupiter Icy Moons Orbite r (JIMO) mission included high power and high ISP gridded ion th rusters, increased thruster lif etime, high power uptake power processing units (PPUs), and radia tion-hardened components. Con tinued research in these areas is essential if advanced robotic and human missions to interplanetary destinations is ever to be accomplished.34 Fundamental Research and EP Description Within the field of electric propulsion, three main subdivisions have arisen: electrothermal electrostatic, and electromagnetic propulsion. In electrothermal propulsion, the propellant is heated by an electrical process and expanded through a suitable nozzl e. Three subclasses of this division include resistojets, ar cjets, and inductively and radi atively heated devices. In electrostatic propulsion, the propellan t is accelerated by direct applic ation of electrostatic forces to ionized particles. The most common thruster t ype in this class is th e ion thruster, though some development work has also been done on fiel d emission electric propulsion (FEEP), which produces minute amounts of thrust. In the th ird category of electromagnetic propulsion, the propellant is accelerated under the combined acti on of electric and magnetic fields. Subclasses in this electric propulsion category include magnetoplasmadynamic thrusters (MPDT), Hallcurrent accelerators, an d pulsed plasma devices.35 When considering potential EP thrusters for interplanetary, hi gh-energy requirement missions, it is currently only feasible to consider the electrostatic and el ectromagnetic classes of propulsion. The electrothermal method falls behi nd the other two classe s due to fundamental thermal limitations for exhaust speeds and lifetimes that primarily result from the heating and expansion processes required of electrothermal accelerators.35 Acceleration of the propellant by

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60 external forces, however, will allow for higher e fficiencies and specific impulse values. The specific thruster types within the electrostatic and electromagnetic categories that were considered in this study include ion, MPD, and Hall thrusters. Ion thrusters One of the electrostatic devi ces capable of satisfying the acceleration levels described above is the ion thruster. It functions by acceler ating a beam of atomic ions with a suitable electric field and then neutralizing it with a fl ux of free electrons. Positive atomic ions are liberated from the propellant source and accelerat ed by the electrostatic field created with strategically placed magnets. These atoms are then combined with the electron source outside of the grid to produce a net zero charge stream with speed determined by the net potential drop over the distance between atom releas e and the neutralization plate.35 Although the ion engine is one of the most co mplex within the group of EP engines, they also have undergone the most research and deve lopment over the years. They are appealing for space travel due to past demonstration of l ong lifetime, high efficiency, and high specific impulses. For the purposes of this study, the lifetime was assumed to be 20000 hours, the ISP was estimated at 6000 (s), the thruster efficiency at 90%, and the thrust per thruster at 1 (N) with xenon as the propellant choice.26,35 Disadvantages with this thruster include the indicated low thrust density, high system complexity, and high PPU specific masses. This technology has been demonstrated on over a dozen U.S. and Soviet flight tests dating back to 1962. The first operational use of ion engines was in 1994 for th e Japanese ETS-6 and CO METS satellites, and was followed by use on a commercial satellite bu s in 1997 and the interplanetary Deep Space NASA mission in 1998.35 Future research on these systems will most likely focus on developing higher power capabilities, decreasing the mass of high power PPUs, and developing extremely small propellant feed systems.36

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61 MPD thrusters Despite having considered only one electrostat ic thruster in this study, both the MPD and Hall thruster engines were considered from the electromagnetic group. This follows appropriately from the fact that electromagne tic acceleration presents many possibilities for implementation given that the applied fields a nd internal currents ma y be steady, pulsed, or alternating over a range of frequencies. There are also a variety of pr opellant types, electrode configurations, and means for injection and ioni zation of atoms that are available when using these electromagnetic thrusters.35 An MPD thruster is configured with a centr al coaxial geometry cathode, an annular anode, and an interelectrode insulator. In this thrust er type, gaseous propellant flows into the upstream part of the channel whereby the atoms are ionize d with a uniform electric arc, compressed into a hot plasma just beyond the cathode tip, and expanded out the thruster as plasma exhaust. These thrusters are capable of high specific impulses, moderate efficiencies and high thrust. The parameters assumed for this study include a li fetime of 2400 (h), an ISP of 5000 (s), an efficiency of 65%, and a thrust of 100 (N).9,26,35,37 It is to be noted however, that megawatt power levels and the subsequent propellant flow rates for MPD thrusters have not yet been tested and are both technologic al and economic problems.35 A major objective for MPDT research is the achievement of thrust effici encies greater than 50% with non-condensable propellants while operating the thrusters below the plasma instability threshold.35 Hall thrusters Electromagnetic Hall thrusters rely on use of th e same-named Hall effect, which refers to the potential difference on opposite sides of a thin sheet of conduction material through which electric current created by a magnetic field applie d perpendicular to the Hall element is flowing. In this type of thruster, propell ant is accelerated by an electric fi eld in a plasma discharge with a

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62 radial magnetic field. Channel and field geomet ries lock the plasma electrons into nearly collisionless cross-stream drifts, leaving th e positive ions to accelerate downstream under the applied electric field.35 This thruster device is sometime s thought of as a cross between an electrostatic and electromagnetic accelerator. Although Hall thrusters have a re latively high specific impulse a nd thruster efficiency, they have a low thrust output per thruster, and pa st testing has shown lif e-limiting erosion and unstable and oscillatory plasma discharges to be of concern.34 The assumed performance parameters used in this study include a lifetime of 8000 hours, a specific impulse of 400 (s), a thruster efficiency of 62 %, and a thrust of 0.4 (N).38 Research interests for Hall thrusters include using krypton as propellant, increasing the life time to greater than 8000 hours, and eliminating the erosion in the plasma discharge chamber.34 Despite knowledge of current performance parame ters including ISP, efficiency, and thrust for each of the EP thruster types under considera tion, the type of missions that the thrusters are being applied to requires some assumptions on the future state of the technologies. Advanced interplanetary missions will have very high delta-V and burn time requirements that may be feasible for these EP systems on paper, but have not actually been tested experimentally. As alluded to earlier, testing of these thrusters at high powers and for long-life durations will, in itself, be a demanding task. However, as the mi ssions being studied have departure dates on the order of 2030 for Mars and 2040 for Saturn and Jupite r, a leap of faith will have to be taken by assuming that the technologies will be availa ble to satisfy the mission requirements. The assumptions made in this study also included the thrusters being able to operate at very high power with PPUs that can pro cess power coming from megawatt -level nuclear reactors, and thrusters that can burn for considerably long time s. It is suspect that these capabilities will

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63 improve over the next 20-40 years and the performance parameters will most likely increase along with them. Development of the Nuclear Electric Propulsion (NEP) Model The NEP architecture was broken-down into three systems based on the type of NEP thruster being used. This was different than the breakdown for NTR systems, which varied by reactor fuel type. Since the propulsive perfor mance of NEP systems was primarily dependent on thruster type and this performance can vary grea tly between different thruster options, this was the primary focus. The reactor needed for an NE P system was unimodal in the sense that it only needed to generate power (in contrast to the usual connotation of unim odal as only producing propulsive heat). It was thus not extremely impor tant what type of reactor was used since the masses of all reactor types were only slightly di fferent and the reactor mass was merely a small fraction of the total mass of the CTV. The thruster types chosen for analysis in this trade study included MPD thrusters, ion thrusters, and Hall thrusters. Each of thes e thruster types would provide different ISP, efficiency, and overall performance capabilities to an NEP spacecraft, thus it was paramount to properly explore the breadth of NEP systems by assessing each of these thruster options. The schematic shown in Figure 3-3 depicts th e spacecraft components th at together makeup the NEP vehicle architecture. It is apparent in this diagram that the reactors served only as power generation for the spacecraft subsystems and electric propulsion mechanism. While the blue components are part of the power gene ration subsystem and the yellow components comprise the set of common vehicle components, the pink components now re present the electric propulsion subsystem. In an NEP system, electri c thrusters provide the sole means of propulsion and require both liquid propellan t and electric power for func tioning. The arrows in the

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64 schematic are divided between the solid black arro ws that depict a cool ant or propellant flow, and the dashed black lines that indi cate flow of power to each system. Power The considerations for determining the reactor mass for the NEP power reactor were much different than those for sizing the different NTR reactors. Since space reactor fuels provided the trade within the NTR architecture, the different mass sizes of each of the NTR reactors was pivotal to the results since mass was based upon fuel choice. In addition, the NTR reactors provided the propulsive energy for th e spacecraft, making fuel parameters such as temperature and ISP important variables in an overall comparison of NTR systems. In the case of the NEP power reactor, the only parameter that went into the NEP architecture model based on the reactor was that of the reactor mass. In addition, only one reactor was needed for all of the NEP systems, since the NEP trade was essentially on the type of electric thruster being used. The propulsive properties of the EP thruster were what differentiated these systems, and thus the mass of the reactor simply became another term to lump into the total mass carried along with the spacecraft. It was thus decided to label the NEP power r eactor as a fast reactor, without specification to its specific fuel form. This generalism was used so that various data-points from literature sources could be used to develop an equation to be used in calculating reactor mass based on the power requirements calculated by the NEP model. Three NEP power reactor descriptions were taken from previous NEP mission analysis studies and their main descriptive parameters were compiled in Table 3-4.32,37,39 Although each of the three reactors from these studies had slightly different fuel forms, they were all fast reactors, as expected for an NE P system. As discussed pr eviously, fast reactors are ideal for NEP systems sin ce they can provide long-duratio n steady power. The mass and

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65 power data-points corresponding to these three reactors included the total mass of the power reactor, pressure vessel, and inte rnal and external shielding. By plotting each of the points on the curve seen in Figure 3-4, a trendline was used to calculate Equation 3-17, where the mass in (kg) is dependent on the electric power requirements Pe in (MWe) of the CTV. 9 2814 12 976 939 532 Pe Pe MassNEP reactor (3-17) The reactor masses were thus calculated sepa rately for each mission since the calculated power requirements varied for each mission. It s hould also be noted that the ISP was assumed for the NEP reactor to be 925 (s), since this is the same ISP as the CERMET fast reactor used in the NTR model. In order to better understand the flow of power and the associated losses from the reactor to the thruster, a pictorial repr esentation of this flow, along with parameter descriptions, is found in Figure 3-5. Because the Brayton power conve rsion unit was indiscriminate of the final destination of its electric power, it provided the function of converting all thermal power generated by the reactor to electric power for spacecraft needs. The PMAD system, which consisted of cabling, fault protection, and switchi ng gear, then directed the power to the appropriate spacecraft loads.16 Losses were found in the Brayton Conversion Unit, PPU, and electric thrusters, as shown a nd described in the power and effi ciency diagram. Some of the major equations used to describe the power and efficiency calculations visualized in the flow diagram are given by the thermal power Equatio n 3-18, and the electric power Equation 3-19, where the power for propulsive needs is give n by Equation 3-20, and the overall system efficiency is given by Equation 3-21 spacecraft s prop s sP P P, (3-18) spacecraft e prop e s s eP P P P, ,* (3-19)

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66 th pp prop eP P P (3-20) th pp s T * (3-21) Propulsion The size of the propulsion system was de pendent on two main factors: the energy requirements for the mission, and the remaining spacecraft vehicle mass. The energy requirements for the NEP spacecraft consisted of a delta-V value for the outbound and inbound transfers and the burn time of the thrusters during these transfers. It is important to note that this was the only architecture that requi red spiral trajectories out of th e Earth and destination planets atmospheres. This was necessary due to the need for the spacecraft to pi ck up speed in order to escape from the atmosphere of each respective planet. Although the propellant was sized according to the total requirements of both the tr ansfer and spiral durations, the crew would only be on the CTV for the transfer duration as it wa s assumed that they were transferred to the spacecraft after it picked up the necessary speed for escape. This was important so that the requirements for transfer time given in the mission profiles would not have to include the rather long spiral times required by the NEP system. This drawback to the NEP system was not neglected as it was accounted for in the final tradespace analysis. In addition to energy requirements, the sizing of the system propellant was also dependent on the weight of all the components it had to pr opel through space. This weight included the communication and navigation components, TransHab and ECRV, truss structure, power reactor, Brayton Conversion Unit, radiator s, and the propulsive thruster s and PPUs. More simply put, enough propellant had to beonboard to carry the rest of the spacecraft weight to its destination. It must not be forgotten that the weight of the unus ed propellant itself also played into the weight equation since it too was carried onboard the spacecraft until it was used.

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67 The two main mass estimates needed to size the propellant mass for outbound and return journeys were the IMLEO (1 TOTm) and the initial mass at departur e from the destination planet (2 TOTm). The propellant loads required to carry these masses were divided into two separate equations (Equation 3-22 and Equation 3-23) since the outbound and re turn burns are two distinct entities. ) 1 ( */ 1 1 1ev v TOT prope m m (3-22) ) 1 ( */ 2 2 2ev v TOT prope m m (3-23) This allowed for the second propellant load eq uation to be calculated assuming a smaller mass due to the first propellant load being burned off, and from dropping its associated propellant tank. The initial mass quantities at in itiation of the two main burns are given in Equation 3-24 and Equation 3-25, where it can be seen that the propellant and total mass equations are interdependent. 2 1 1 prop inert NEP TOTm m m (3-24) 2 1 tan _ 2 prop k prop inert NEP TOTm m m m (3-25) The term inert NEPm_ used in the mass equations represen ts the total spacecraft mass minus the propellant, an equation for which is given by Equation 3-26, where th e terms represent the mass of the power reactor, propulsion hardware communication and navigation components, and TransHab, ECRV and structural truss masses, respectively. truss ECRV Transhab nav comm th pp power inert NEPm m m m m, , 15 (3-26) As the mass sizing of the power reactor and its associated pressure ve ssel, shield, and feed system have been previously discussed, and sizing of the spacecraft communication and

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68 navigation systems is to follow, only the sizing of the propulsion system will be discussed herein. The mass estimates for the PPU and thruster hardware were taken from both text and literature sources that had specific mass (kg/kWe) values for ion, MPD, and Hall thrusters. The MPD thruster and PPU combined specific mass was the lowest of the thr ee, at 0.6 (kg/kWe). This parameter was mostly due to the very high po wer levels processed in this simple and robust thruster device, and has made MPD thrusters highly desirable for interplanetary missions.26,35 The specific mass estimate for the Hall thruster system was based off of both an advanced thruster and advanced PPU design. Th e High Voltage Hall A ccelerator (HIVHAC) Development Program predicted a possible future thruster specific mass of 1.3 (kg/kWe), while PPU specific masses ranged from a progressi ve 5 (kg/kWe) to a very advanced 2 (kg/kWe).37,38,40 Given the distant departure times of the reference mission profiles under consideration in this study, the advanced PPU specific mass was assumed, for a total PPU and thruster specific mass of 3.3 (kg/kWe). Although ion thrusters are known to require very heavy PPUs (~10 (kg/kWe)), a similar technology advan cement was assumed for the ion system and a value of 4 (kg/kWe) was used for the PPU specific mass.35 Combined with a literature datapoint of 1.2 (kg/kWe) for the i on thruster itself, a total spec ific mass of 5.2 (kg/kWe) was assumed for the thruster and PPU combination.37 The equations for the combined PPU and thruster mass for the MPD system is thus gi ven by Equation 3-27, the Hall system by Equation 3-28, and the ion system by Equation 3-29, where the term Pe,prop is the electric power supplied to the EP thrusters. ) / 1 ( * 6 0,lifetime duration burn P massprop e th pp (3-27) ) / 1 ( * 3 3,lifetime duration burn P massprop e th pp (3-28)

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69 lifetime duration burn P massprop e th pp/ 1 ( * 2 5, ) (3-29) Although these specific mass estimates alone pr ovided the sizing for the thrusters based upon satisfying the mission energy requirements, th ey do not account for the functional lifetimes of the thrusters that may limit the duration of th eir use. Thus contingency weight had to be added for thruster hardware for missions that re quired burns greater than the lifetime of the thruster selected. The burn durations for the NE P system were defined as the spiral out/in times, along with burn times between planets. This duration is not the same as the crewed mission duration, which would include the tr ansfer times and stay time on the destination planet. Given this assumption, thrust correction maneuvers needed during crew stay at the destination planet were assumed supplied via alternate met hods such as reaction control systems. The mass estimates for the xenon and liquid hydrogen propellant tanks differed based on the propellant, not the thru ster, and were approximated with textbook equations.16 The tank estimates for the ion and Hall thrusters thus us ed the same equation since both required xenon propellant, while the MPD thruster used a tank equation based upon hydrogen propellant.26 The differences in the densities of the propellants led to a different required volume per unit mass, and thus different propellant tank volumes and masses. Text book sources gave Equation 3-30 for MPD tank mass, while the equation for the xen on propelled ion and Hall thrusters was given by Equation 3-31. prop H kmass mass 287 .2 tan (3-30) 3 / 2 tan* 154 075 0 52prop prop Xe kmass mass mass (3-31) For both equations, the term massprop represented the total amount of propellant used for the mission.16,32

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70 The power conversion, power management, and heat dissipation system masses were not dependent on the thruster used, thus common equations could be used for all thruster models. Due to the lack of mass approximations for the PMAD component in textbook material, three data-points from the literature were used to approximate the PMAD mass equation. Microsoft Excel was used to derive a logarithmic approxima tion to the data-points trend line, a plot of which can be seen in Figure 3-6. Issues aros e for power requirements less than about 1000 kWe due to the nature of the logari thmic function, thus a regressed lin ear function was used for lower power requirements. The formula for PMAD ma ss is given in Equation 3-32, which shows the calculation based on the elec tric power requirement eP in (kWe). ) 1000 ( 635 );. 1000 ( 5 1067 ) ( 7 7772 e e e e PMADforP P forP P Ln mass (3-32) In the calculation for the Brayton Unit, a specific mass of 0.875 wa s used and was taken from a literature data-point for a similar NEP crewed vehicle.32 The Brayton Unit, or CBC, mass is given by Equation 3-33. e CBCP mass 875 0 (3-33) The mass of the radiators was determined by a straightforward e quation depicting the radiator mass needed to rid the system of excess heat due to losses. The value for the specific mass of the radiators was estimated to be 1.5, whic h is slightly more conservative than a text estimate of 0.1-0.4 kg/kW, and in close approximation to the Brayton to radiator mass ratio from the previously mentioned literature source.16,32 The equation used for the radiator mass estimate is seen in Equation 3-34, where radSpMass is the specific mass value of 1.5 for the radiator. rad S prop s s S pp th prop s radSpMass P P P mass )) 1 )( ( ) * 1 ( (, (3-34)

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71 An additional equation that was used to estimate the mass for the truss structure for the entire spacecraft was Eq uation 3-35, where again, inertmass refers to the entire spacecraft mass not including propellant. inert trussmass mass % 10 (3-35) In addition to sizing the NEP system compone nts, the propulsion e quations also derived the required thermal and electric power for the spacecraft, along with the thrust required by the EP thrusters. One important poi nt to note is that the thermal power generated by the reactor was converted to electric power and then split be tween the propulsion system that operates the thrusters, and the componen ts requiring constant power. These components included communication and navigation systems along with th e TransHab module. This effect is shown through the thermal power equation, Equation 3-36. ) ( 2 ) ( *, , 2 Transhab NAV COMM s prop T prop sP g Isp m P (3-36) The total electric power requirements for these two separate entities combined is given by Equation 3-37. s s eP P (3-37) The thrust variable T is then based only on the therma l power devoted to the propulsion system, prop sP,, as seen by Equation 3-38. g Isp P Tprop s T* * 2, (3-38)

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72 Knowing the value of the thrust is only signi ficant for calculating th e number of thrusters that would be needed for a particular EP syst em given by the simple relation in Equation 3-39, where the term in the denominator indicates the characteristic thrust per EP thruster. thruster T T thrusters_ # (3-39) It should be clear from inspection of the an alysis performed to develop the NEP propulsion equations, that thruster-specifi c parameters played a major pa rt in determining spacecraft performance. The EP thruster efficiency, speci fic impulse, and thrust pe r thruster had a large impact on power and thrust requirements, overa ll system efficiencies, and wet and dry mass estimates for the spacecraft. These thruster charact eristic parameters were even used as inputs in the ephemeris models, which determined the energy requirements for the system. The specific mass, fuel, and lifetime parameters for each thruster were the main drivers for the mass estimates of the actual thruster and PPU hardware, along with the propellant. The parameters used for the ion, MPD, and Hall thrusters in this study are briefly su mmarized again in Table 3-5. Hybrid Power and Propulsion Background Up until this point, Hybrid vehicle architectures have only been minimally analyzed through paper studies. One such study indicates that an integrat ed nuclear thermal and nuclear electric propulsion system would grant the ability to obtain from two systems much more than one can obtain from a single technology.6 A NASA presentation on Bimodal Nuclear Thermal Rocket Propulsion for Future Human Mars Explor ation Missions gives pictures, schematics, and explanation of how a bimodal nuc lear thermal reactor could appl y electrical power production to a secondary electric propulsion system. The basic methodology is that the bimodal nuclear thermal rocket (BNTR) would produce short bu rsts of high thrust for escape mechanisms,

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73 followed by a long power generation phase wherein the electric thrusters would be operational.41 It is the hope of mission planners that a Hybr id spacecraft would combine the benefits of both NTR and NEP to increase overall performance for challenging interplanetary missions. Development of the Hybrid Model The Hybrid vehicle architect ure schematic was devised as a cross between the NTR bimodal system and the NEP system. As can be s een in Figure 3-7, the schematic for the Hybrid system maintained the power and propulsion subs ystems seen previously in the NTR bimodal schematic, along with the electric propulsion subs ystem seen in the NEP schematic. Dashed lines in this schematic refer to power generati on flow, while the solid black arrows refer to propellant or coolant flow. The unique vehicle design in the Hybrid sche matic allowed for the reactor to provide power for the spacecraft needs and the necessary thermal energy and electric power for two modes of propulsion. The location of the two propulsion systems re lative to the TransHab was indeed intentional, as the reactor system woul d most likely reside at the opposite end of the TransHab, while the EP system would be on the same end. The two thrust mechanisms would also therefore produ ce thrust in opposite dire ctions relative to the sp acecraft itself, requiring realignment of the vehicle be fore changing propulsive mode. Power The same bimodal NTR reactor that was us ed for thermal power generation for the NTR bimodal architecture was also used for the thr ee Hybrid architectures. A bimodal engine was required for this architecture in order to pr ovide thermal propulsion for the NTR engines and electric power for the EP thrusters and other spac ecraft functions. The bim odal reactor utilized a fast fission spectrum and CERMET reactor fuel. An assumption of two reactors onboard the spacecraft was used just as for th e other architectures. In order to estimate the total reactor mass,

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74 a dependence on the electric power requirement wa s used just as it was for the NEP system. This estimate was found by using the same equation for reactor mass as the NEP system, but adding an additional weight to account for NT R reactor fuel and propulsive hardware. The bimodal engine at 5272.3 (kg) wa s designed to have an electric power output of 25 (kWe). An NEP reactor operating at 25 (kWe) would have a calculated mass of 2839.3 (kg). This difference between the two of 2432.9 (kg) was thus assumed to be the additional NTR weight added to the calculated weight of the reactor based only on the NEP reactor mass. The NEP reactor mass equation for one reactor is shown below in Equation 3-40, and was updated to account for the additional mass needed for the Hybrid reactor. 9 2432 ) 9 2814 12 976 939 53 (2 Pe Pe MassHybrid reactor (3-40) Propulsion The equations that defined the EP thruster sizing, along with the parameters for the EP thrusters, were the same as those used for the hybr id architecture. The EP aspect of the Hybrid vehicle was identical to the NEP architecture wi th all three types of EP thrusters assessed. Although the vehicle had an NTR engine, all co nfigurations utilized the bimodal CERMET reactor, thus the trade was solely on EP thruster type, not on NTR fuel type as it was for the NTR architecture. The equations that changed compared to th e NEP system were those for the propellant sizing. The Hybrid propulsive equations represented operationa lly both NTR and NEP delta-V requirements, in which one NTR delta-V burn occured for planetary escape and was then followed by a long-duration NEP delta-V burn that allowed for capture. Th e first three of the four tanks associated with the four delta-V maneuvers was dropped af ter its respective burn segment. For instance, the mass of the propell ant required for the se cond burn used the IMLEO

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75 mass minus the propellant burned in the first NTR segment along with the tank that housed that propellant. These tanks were non-recovera ble and inevitably became space junk. The equations that defined the propul sive mass requirements for the H ybrid architectures four burns are given by Equation 3-41 through Equation 3-44, wh ere the mass equations at initiation of each burn are given by Equation 3-45 through Equation 3-48. ) 1 ( *_/ 1 1 1NTR ev v TOT NTR prope m m (3-41) ) 1 ( *_/ 2 2 2NEP ev v TOT NEP prope m m (3-42) ) 1 ( *_/ 3 3 3NTR ev v TOT NTR prope m m (3-43) ) 1 ( *_/ 4 4 4NEP ev v TOT NEP prope m m (3-44) 4 3 2 1 _ 1 prop inert NEP inert NTR TOTm m m m (3-45) 1 tan 4 3 2 _ 2 k prop prop inert NEP inert NTR TOTm m m m m (3-46) 2 1 tan 4 3 _ 3 k prop prop inert NEP inert NTR TOTm m m m m (3-47) 3 2 1 tan 4 _ 4 k prop prop inert NEP inert NTR TOTm m m m m (3-48) Note that the propellant mass subscripts indi cating NEP or NTR burns are not included in the second set of total mass equations, but are inst ead merely indicated by the number, or order, in which they occur. Common Vehicle Subsystems TransHab Every vehicle system that was analyzed for the advanced interplanetary missions in this study had a crew living module called the Trans Hab. This term, commonly found in mission

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76 analysis literature, was coined for the actual st ructure that was still in the design phase when research efforts ended in 2001. The TransHab wa s the first space inflatable module designed by NASA, conceived as a technology capable of s upporting a crew of six on an extended space journey. The space habitat had an endoskeletal de sign, with a light and re configurable central structure and deployable pressure shell. The shell was elaborate in that it was made of several layers, could contain water or hydrogen for radiation shielding, and had an outer layer a foot thick. Recommendations drawn from the rese arch completed during this NASA program included having a three-level internal layout with isolated crew quarter s at the center, and exercise and hygiene stations situated on different levels from the public functions in the kitchen, dining and conference areas. Unfortunately bud get cuts along with reported values of $100 million to build the structure led to the projects demise.41 After the four-year research period between 1997 and 2001, Bigelow Aerospace purchased the rights to the patents developed by NASA and is currently continuing development of a similar structure for future commercial space applications.11 The ECRV was an additional habitation environm ent designed to be attached externally to the TransHab for a majority of the mission dura tion. In a hypothetical mission, when the CTV is being assembled in LEO prior to flight, the crew will be transferred from the launch vehicle to the CTV via the ECRV. The TransH ab will then be inflated and flooring and partitions deployed so the crew may move from the ECRV into the TransHab for the remainder of the trip. The ECRV will also likely be the capsule the crew w ill use to aerobrake into the Earths atmosphere after the interplanetary voyage. This would follow the propulsi ve capture of the entire CTV back into LEO and disengagement of the ECRV from the main space vehicle.

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77 A vital consideration for the future design a nd sizing of the TransH ab will be radiation shielding. It has been noted by several literatu re sources that the h azards of space radiation create the primary restraint on crewed interplaneta ry space missions. This is a problem that has not yet been resolved by nature due to the many complexities space radiation presents. Future crewmembers will be exposed to many types of radiation (ionizing and non-ionizing), which fluctuate considerably over time and location. This implies that different crew procedures and levels of protection will be need ed over the course of a mission. More research will have to be done on the shielding properties used for prot ection, as human-rated structures must be approached with a very ra dical design perspective.42 For example, some studies have proposed surrounding the crew module with hydrogen propellant tanks for radiation protection. Although the hydrogen-containing materials would be goo d radiation absorbers, the aluminum tank structure would actually produce dangerous secondary radiation particles. There is also uncertainty in the biological response to hi gh-charge radiation. Currently, no exposure limitations have been set for deep space missions, and recommendations simply rely on exposure limits for operations in LEO.43 Some work has been done in the past to try and model how much and what type of radiation could be expected on interplanetary missions. Th e two main forms of ionizing radiation that are estimated include the constant but low intensity galactic cosmic radiation (GCR), and the low frequency, but high intensity sola r particle events (SPE). One study reported that with 1280 (kg) of polyethylene TransHab shielding for a Mars journey, 13.38 (cSv) of radiation dose equivalent was detected for SPE and 17.65 (cSv) for GCR for each one-way 180 day transfer. The surface habitation on Mars re sulted in 28.05 (cSv) for GCR and 2.23 (cSv) for SPE.44 The lower rate of exposure on Mars was a re sult of the natural ra diation shielding from

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78 the Martian atmosphere. Although this study fou nd a total radiation dose equivalent rate under the NASA LEO limit of 50 (cSv/year), the rate of exposure during the transfer period would indeed have been over the limit if it had been received for an entire year. A long duration transfer would have led to 63.5 (cSv/year) exposur e, which is a significant violation of the current limit. Although radiation shielding woul d be a significant factor in the design and development of the TransHab, it was deemed too difficult to a ccount for shielding mass for all of the reference missions run in this study. It was assumed ther efore that the 1900 (kg) th at was assumed for the TransHab shielding weight, in addition to other materials such as wastewater and propellant tanks, would provide the necessa ry shielding for the trip. The breakdown of the TransHabrelated masses thus became quite simple and was modeled from previous mission analysis sources. The module itself had a mass of 21000 ( kg), the 6 crewmembers and their belongings equated to 600 (kg), and the ECRV was set at 5100 (kg). The consumables were dependent on the length of the mission (including stay time at planet), and were sized at 2.45 (kg) per person per day for food and water.18 The breakdown of masses for the TransHab components can be found in Table 3-6. The power requirement for the TransHab was also an important parameter to consider when generating the total operating power require ments of the spacecraft. The actual value of constant power that must be provided for the crew module so that ever yday living and working needs can be satisfied was a difficult value to estimate, especially given the scarcity of information in the literature. In the lecture notes of G.L. Kulcinsky, however, a value of 100 (kWe) was given for a manned outpost on the surface of Mars.45 It was believed that this

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79 would be a reasonable estimate for the TranHab si nce it would require similar power needs to a planetary ground station. Spacecraft Functional Components The remaining spacecraft components that are common to each of the space architectures were those that provided fo r various communication and navi gation functions to get the spacecraft to its destination and back. These system components were broken down into the following groups: command and data handling (C&DH), tracking, telemetry and command (TT&C), attitude determination and control system (ADCS), the reaction control system (RCS), and truss structure used to link all spacecraf t hardware. The C&DH subsystem distributed commands and accumulated, stored, and formatte d data from the spacecraft. The TT&C subsystem linked spacecraft with ground or othe r spacecraft by receiving commands and sending status telemetry. The ADCS provided stability a nd orientation in the desired directions during a mission despite external disturbances. The pr opulsive RCS subcomponent was linked to the ADCS by performing the thrust actions that were needed by the spacecraft as determined by the ADCS. Although the general functions and equipment needed by these subcomponents was important to understand from the systems engin eering perspective, th e two most important parameters in the spacecraft model development were the mass and power estimates for the components. It was again difficult to estimate the size and power of these subcomponents given the limited number of papers found in the litera ture that went to similar depth-levels for spacecraft sizing, yet one paper provided values for a reasonable estimati on. For the purposes of this study, the ADCS was estimated at 70 (kg), th e C&DH at 35 (kg), and the TT&C at 205 (kg). The total power required was taken from th e same reference, which designated the communication loads as requiring a power of 5 (kWe).37 The RCS mass was taken from two

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80 data-points, one from an NTR study, and the ot her from an NEP study. The NTR study showed an RCS mass that was 5.7% times that of the propulsion prope llant mass, the NEP study gave a value of 6.9%, and one textbook suggested a range from 3-10%.16,18,37 It was thus decided to set the RCS mass to 6% of the propellant mass requi red by the main propulsion systems for a given architecture. Since the RCS would only requir e a minimal amount of power at various small time increments over the course of the duration of the f light, it was decided to neglect its power requirement when sizing the overall power of th e nuclear reactors. The structure of each spacecraft was based on the total mass of the sy stem, scaling with system mass. A common estimate which was used to size the truss stru cture mass was 10% of the overall spacecraft mass.16 A summary of the mass breakdown between common subcomponents is given in Table 3-7. Analysis Tools ModelCenter Software The main analysis tool used in this study wa s Phoenix Integration, Inc.s ModelCenter. This tool created a visual environment for pr ocess integration and employed complex design exploration techniques, making it ideal for pe rforming trade studies and exploring a design space.46 One major benefit to this tool was that it reduced both the time to run individual programs and the possibility for human errors th at lead to costly mistakes. In addition, ModelCenter offered powerful and timesaving de sign tools that included parametric plots, optimization tools and linking of models. Each of these assets together provided the ability to conduct a large trade study within a reas onable time period. The Analysis Server, a product by the same company, was used in conjunction with ModelCenter in order to be able to convert codes being used into reusable components that were directly accessible within ModelCenter. The Analysis Server was a flexible, Java-based software

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81 server that allowed applications to be wrappe d for integrated engineering processes. The wrapped components, which are termed FileWrappers, are built from the FileWrapper utility, which was designed to automate the execution of analyses based on file I/O programs. The FileWrapper components thus had the ability to au tomatically edit the appropr iate input files, run the codes executable, and parse the output file whenever it is executed. A template file, which is a copy of the input file in standard format, had to be creat ed for use with a FileWrapper component, as the FileWrapper loads the template file and substitutes any user-inputs into it.46 Incorporation of Models into ModelCenter Framework ModelCenter was used in this study by wra pping individual Excel worksheets and ephemeris codes into ModelCenter system mode ls that allowed for user inputs, automatic passing of data from one worksheet to another, optimization and converging of parameters, and finally parametric analyses. Ephemeris models The IPREP and CHEBYTOP ephemeris codes we re used to generate the NTR and the Hybrid and NEP ephemeris modeling tools, respectively. This was accomplished through the use of ModelCenter in conjunction with An alysis Server. The IPREP and CHEBYTOP executables were first input into the models by generating FileWrappers and input templates that the Analysis Server used to conne ct with the ModelCenter files. An example of the FileWrapper and templates created for the three propulsion syst em-specific ephemeris models is given by the CHEBYTOP Hybrid model files found in Appendix C: ModelCenter Template and Filewrapper. Creating these FileWrappers allowed the executio n of IPREP and CHEBYTOP to be done using components within their respectiv e ephemeris models, with input and output variables appearing in easy-to-use lists within the ModelCenter GUI.

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82 The IPREP component was able to run an en tire two-way mission in one sweep, but the CHEBYTOP component had to be run twice within the NEP and Hybrid models to simulate the outbound and return transfers, and was thus input as two separate compon ents in each of the models. The data for the apogee and perigee variables for each destination was held within an Excel worksheet in each of the th ree models so that selection of planetary destination would allow the correct orbit parameters to be passe d to the IPREP or CHEBYTOP components. A tool to aid in converting a date input by the user to a Julian date required by the executable was also incorporated as an Excel component. For the two models that used the CHEBYTOP components, an additional Excel component was added that uploaded the thruster-related para meter values upon user selection of a thruster type. These parameters included the assumed values for thruster efficiency, ISP, and thrust force per thruster for the selection of Hall, ion, or MPD thruster type Once all of these components were incorporated into the mode l, input variables that included departure date, stay time and transfer time were incorporated into the user interface and then linked to the appropriate variables in the model components. Key output parameters that were desired from the models included the delta-Vs, electric thru ster burn times, and NEP spiral times. A screenshot of the ModelCenter NEP ephemeris models graphical us er interface (GUI) is given in Figure 3-8. ModelCenters Darwin optimization tool was al so incorporated into each of the three ephemeris models. This genetic algorithm-based t ool with fully featured GUI was integrated as a trade study plug-in within ModelCenter. The genetic algorithm method is based on probabilistic rules with evolution achieved through application of probabilistic operators that mimic biological genetics.46 The goal that was achieved by using this optimizer within the ephemeris model was to minimize the overall missi on delta-V for a given mission profile. Since

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83 each mission profile was designed with ranges on th e parameters for departure date, stay time, and transfer time, these variables became the design space for the optimization. The lower and upper bounds for these parameters became inputs that could be manually set by the user depending on the mission profile under review. Execution of this t ool enabled the energy requirements to be minimized based on manipulati ng each of the four va riables in the design space to find the very best answer given the specified mission profile requirements. Attempting to perform such an optimization manually would have undoubtedly required many hours of running the ephemeris model trying to make improve ments one variable at a time. A screenshot of the Darwin Optimization tool GUI is shown in Figure 3-9. Vehicle architecture models Incorporating the vehicle ar chitecture models into ModelCenter was much more straightforward than the ephemeris models si nce only Excel models that were fully selfgenerated were used. This involved using the Ex cel component tool to insert the various Excel models needed for each architecture, and then usi ng the linking tool to link parameters within the model. The NEP and Hybrid architectures both had two main Excel worksheets incorporated into their ModelCen ter models. The first was the main propulsion and power model that sized the reactor, while the second was a th ruster selector tool. This la tter worksheet allowed the user to select the desired thrust er type from the ModelCente r GUI, automatically updating the appropriate propulsive parameters characteristic to that th ruster in the sizing model. The NTR model was slightly different in that its secondary Excel worksheet was a reactor fuel selector, which would input the fuel parameters for the fuel selected into the main NTR propulsion and power model. A screenshot of the NEP model showing th e input parameter and output parameter lists, along with a block comp onent input and output diagram is shown in Figure 3-10.

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84 Mission architecture models The term mission architecture was used in this study to refer to a vehicle architecture performing a mission profile. From the modeling perspective this would equate to combining the architectures ephemeris model and its power and propulsion model. The mission architectures were then developed into their respective ModelCenter models, and were created by combining both the ephemeris and main architectur e model. This allowed a user to then run an entire mission, deriving the IMLEO and othe r output parameters by simply inputting the mission profile scenario and fuel or thruster type. Some difficulties were encounter ed, however, when trying to in corporate certain levels of automation into the models, which would have be en timesaving mechanisms for the user. For instance, the ephemeris models Darwin optimizer tool that manipulated de parture date, transfer and stay time in order to mini mize delta-V energy requirement s would not function properly within the larger mission architecture ModelCenter model. This was due to the fact that the term it was trying to optimize was being passed to the power and propulsion model, and was no longer an output parameter of the ephemeris model as required. This finding affected the planned methodology for the data analysis process. It was decided to maintain use of the ephemeris ModelCenter models for optimization purposes, reco rd all optimized depart ure dates, transfer and stay times in a database, and manually ente r these parameters into the mission architecture models to run the full-scale models. When the mission architecture models were run, the optimized delta-V energy requirements were au tomatically generated in the ephemeris model and passed to the power and propulsion model. Figure 3-11 provides a visual of a mission architecture model GUI that shows these user inputs, the two models incorporated into the larger model, and the desired output parameters.

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85 Validation of models Following completion of the mission architecture models it was decided to attempt to validate the models by comparing th e IMLEO results from alternate sources to the results of the mission architecture models developed in this stu dy. One literature source for each of the three main architectures was used for comparison. Atte mpts were made in each case to try and model a very similar mission to that presented in each study, though lack of specific mission analysis data such as the energy requi rements used for sizing was recognized as a weakness in the validation methodology. The NTR architecture model was validated using the results from a very detailed Pratt & Whitney Rocketdyne NTR model. This model was built upon numerous smaller models, highlevel analysis and very thorough attention to reactor design and development details.47 All of the fuels analyzed in the NTR model studied in this paper were also anal yzed by the P&W model, thus it was decided to compare the IMLEO results fo r both models at the 15 (k lbf) thrust setting. One of the main differences between the models which may have some effect on the validation results, is that the P&W model assumes 1, 2 or 3 engine confi gurations wherein each engine provides the indicated thrust. The NTR architect ure model developed in this study assumed only one engine was providing the thrust yet an extra engines weight was included in the model for redundancy purposes. The mission that was analyzed for the NTR va lidation was a round-trip Mars mission with these specific delta-V energy requirements: TMI (Trans-Mars Injection)=3.60 (km/s) MOC (Mars Orbit Capture)=1.44 (km/s) TEI (Trans-Earth Injection)=1.09 (km/s) EOC (Earth Orbit Capture)=0 (km/s)

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86 By running the NTR architecture model with th e above energy requirements for each of the fuel selections (graphite, composite, carbide, and CERMET), a comparison of IMLEO estimates was made with the P&W model. The results from both models, along with the percent error between the two sets of results, can be found in Table 3-8. The percent error data-points ranged from 0.4% for the graphite model up to 8.63% fo r the CERMET model, with the lowest error cases pertaining to the more researched graphite and composite fuels. Considering the differing assumptions that most likely went into developm ent of each of the two m odels, along with a vast difference in development time for the models, these comparison results provided strong support for validation of the NTR model. The NEP literature source used in validating the NEP architecture model analyzed a 2033 Mars crewed round-trip mission th at used MPD thrusters and a fa st power reactor. The exact delta-V values used in the study were not given, thus a missi on was set up in the NEP ephemeris model to generate approximate values for the energy requirements. A known difference between the paper-study analysis a nd the approximations used herein wa s that an Earth flyby was used in the paper-study versus the direct Earth re-entry assumed in the ephemeris model. To account for this difference, it was decided to match the prop ellant mass of 276 (kg) given in the literature study in order to determine the correct delta-V fo r the return trip. Using an outbound delta-V of 19.3 (km/s) and return delta-V of 19 (km/s), a transfer time of 275 days, and MPD thruster selection, the resulting IMLEO from the NEP architecture model was found to be 598.8 (MT). The literature source quoted an IMLEO of 560.7 (MT).32 The NEP model thus gave a 6.8% higher estimate of the IMLEO than did the so urce, a rather reasonable value given the assumptions and approximations that had to be made to compare the two estimates.

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87 A study on hybrid vehicles was also used in an attempt to validate th e Hybrid architecture model. The mission was again a ro und-trip crewed Mars mission w ith a short stay. Similar to the previous validation methodology, a longer st ay-time was assumed. This assumption was made since energy requirements were not explici tly cited in the paper-study, and the ephemeris model gave far less demanding energy requireme nts for a longer stay-time on the surface of Mars. The outbound and inbound NTR delta-V requi rements were found to be 3.17 (km/s) and 0.22 (km/s), respectively, while the NEP requirements were 14 and 11.8 (km/s). The literature source cited an IMLEO of 298 (MT) for the hybrid vehicle with MPD thrust ers, and the Hybrid architecture gave a result of 275 (MT), a difference of only 7.7%.6 The validation results presented for each of the three architecture models satisfied the objective of comparing the results seen in a l iterature study for each vehicle type, to those generated by the models developed in this study. The outcome was highly sa tisfactory in that the models created herein did a re latively good job of predicting the same results as professional mission analysis results done by e xperts in each particular power and propulsion field. These results were important since they provided cred ibility to the predictive nature of each of the individual architecture models. This was crucial for conducting a meaningful trade study of all three architectures, since they we re compared directly against each other to determine the best performing vehicle architecture.

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88 Table 3-1. NTR Architecture Systems NTR System Reactor Mode Energy Spectrum Fuel Type 1 Unimodal Thermal Graphite 2 Unimodal Thermal Composite 3 Unimodal Thermal Carbide 4 Unimodal Fast CERMET 5 Bimodal Fast CERMET Table 3-2. Fuel Operating Parameters Chamber Temp. (K) ISP (s) Power (MWt) Graphite 2500 875 306.6 Composite 2700 925 316.6 Carbide 3100 1000 342.0 Cermet 2700 925 316.6 Table 3-3. NTR Reactor Mass Sizing prop. hardware added to total Reactor (kg)Inter n. Shield (kg) Extern. Shield (kg) Total (kg) Graphite 3665.0 988.4 2,942.0 9384.4 Composite 3062.3 825.9 2,458.2 8135.3 Carbide 3404.0 918.0 2,732.5 8843.5 Cermet-uni 1059.1 285.6 850.2 3983.9 Cermet-bi 1680.8 453.3 1,349.2 5272.3 Table 3-4. NEP Power Reactor Descriptions Power (MWe) Mass (kg)Type Spectrum 5 9044 Liquid-metal Fast 1 3845 Particle bed reactorFast 10 17970 Li-cooled pin typeFast Table 3-5. EP Thruster Characteristic Parameters Type Efficiency ISP (s) Thrust (N)S pec_Mass (kg/kWe) Fuel Lifetime (days) Ion 0.90 6000 1.0 5.2 Xenon 850 MPD 0.65 5000 100.0 0.6 Hydrogen 350 Hall 0.60 2500 0.4 3.3 Xenon 100 Table 3-6. TransHab Mass Breakdown Component Mass (kg) Module 21000 ECRV 5100 Crew/Suits 600 Consumables 14.7/day

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89 Table 3-7. Common Subcomponent Mass Breakdown Component Mass (kg) ADCS 70 C&DH 35 TT&C 205 RCS 6%*mass_prop Structure 10%* total mass Table 3-8. NTR Model Validation Mass Estimates P&W Mass (kg)Models Mass (kg)% Error Graphite 239000 238042 0.40% Composite 200000 211738 5.87% Carbide 180000 195854 8.81% Cermet 170000 184666 8.63%

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90 Figure 3-1. NTR Unimodal Vehi cle Architecture Schematic Figure 3-2. NTR Bimodal Vehi cle Architecture Schematic

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91 Figure 3-3. NEP Vehicle Architecture Schematic 0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000 01234567891011 Power (MWe)Reactor System Mass (kg) Figure 3-4. NEP Reactor Mass vs. Power

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92 Figure 3-5. Power and Efficiency Flow Diagram y = 7772.7Ln(x) + 1067.5 0 5000 10000 15000 20000 25000 30000 051015202530 Power (MWe)PMAD Mass (kg) Figure 3-6. PMAD Mass Sizing Equation

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93 Figure 3-7. Hybrid Vehicl e Architecture Schematic Figure 3-8. NEP Ephemeris Tool GUI

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94 Figure 3-9. Darwin Optimization Tool GUI Figure 3-10. ModelCenter Screenshot

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95 Figure 3-11. ModelCenter Mission Architecture GUI

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96 CHAPTER 4 TRADE STUDY RESULTS Due to the breadth and depth of this mission an alysis trade study, the data analysis phase of the investigation was divided be tween three main areas: ephemeri s analysis, vehicle architecture analysis, and tradespace assessment. The epheme ris analysis utilized the ephemeris tools in order to model the energy requirements for the th ree major vehicle systems based upon departure dates and mission duration. The vehicle arch itecture analysis determined performance parameters of the power and propulsion systems for these missions. The tradespace assessment then used these results to assess the entire tradespace matrix based on pre-determined MOEs. The ability to create the parametric and carpet plots that were needed for the data analysis and tradespace analysis to follow were done usin g the time-saving functionality inherent to ModelCenter. ModelCenters data analysis t oolbar was used to set various parameters for plotting, producing multiple charts for different types of interpretation. Figure 4-1 shows the actual ModelCenter Toolbar with intuitive buttons for each data analysis option, along with four examples of plots that can be generated, incl uding carpet, contour, su rface, and statistical dependence plots. Ephemeris Analysis One of the key steps in performing mission anal ysis is using ephemeris tools to determine system energy requirements for a desired missi on. Although the energy requirements for the specific reference mission profiles were needed to assess the tradespace matrix, it was desired to first derive some more general ephemeris results from the models created in this study for the NTR, NEP, and Hybrid vehicles. The single most important parameter to be ob tained from the ephemeris models was the delta-V value, which is essentially a requireme nt on how much energy will be needed by the

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97 propulsion system. This requirement will determ ine how much propellant must be carried on board in order to get the useful hardware to its destination in space. The analysis to follow will allow one to see how this delta-V parameter is dependent on other variables such as stay time, transfer time, and departure date, the effects of which are determined by celestial mechanics. The following plots attest to the capabilities of both the ephemeris models developed for this study using legacy NASA codes, and the Mo delCenter software tool used to run the ephemeris models. Their presentation is separa ted between departure and duration categories, which refer to whether the delta-V parameter wa s plotted against the departure year and month, or versus transfer and stay time. NTR Departure The LEO departure date is one of the major factors in determining delta-V values with ephemeris models. This is why the departure ye ar, month, and day were all used in the process of trying to find the optimal time to leave the Earth. The idea of the synodic periods for each planet can be visualized in Figure 4-2 as a wavelike pattern is seen for the delta-V values over a range of 15 years. The peaks and valleys of eac h curve represent the cy clical nature of the location of the Earth relative to each destination planet over time. It can be seen that the number of cycles is directly proportional to the distance from Earth, giving more minimum-energy opportunities to Mars, and less to Saturn. Note that the data seen in this plot, and similar ones to follow, assumes a transfer time of 200 days fo r Mars, and 1000 days for both Jupiter and Saturn, with a 500-day stay at all planets. Contour plots were also created to asse ss how the delta-V energy requirements were affected by both departure month and year. Fi gure 4-3 shows the dependence of the energy requirement on both of these parameters, where th e color legend in the upper right-hand corner

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98 relates delta-V values in units of (km/s) to the co lored lines seen on the plot This plot indicates that any year and month between the two green curves in the upper left -hand area would be the optimal time of departure. The range of values for delta-V based on all results is between 39.1 and 109.8 (km/s), though changing the stay and transf er time would certainly provide flexibility in further optimizing these results. The Saturn and Jupiter contour plots could not be created for the NTR model using the IPREP ephemeris code. Consistent with prior knowledge of this code, the simulations seemed to break down at these more distant planets and years, specifically with Saturn, and at years after 2042. This also meant that the optimizer tool could not be used to optimize the Jupiter and Saturn reference missions, thus they had to be estimated using a manual optimization process. Duration In addition to departure date, the transfer time and stay time play a major part in determining energy requirements. Delta-V versus transfer time curves were plotted for all destination planets across a range of 200 to 1200 days, fully encompassing the ranges allowed for the reference missions. It can be seen in Figure 4-4 that the requirements for the Mars missions actually get more difficult after 300 tr ansfer days, and the Jupiter and Saturn requirements get less difficult after approximate ly 400 days. This observation can again be related to the distance of each plan et from the Earth and/or the sun. Note that the departure year used for the Mars analysis was 2031, wh ile that for Jupiter and Saturn was 2041. The delta-V values were also plotted against stay time to see how this parameter affects requirements. It can be seen from Figure 4-5 that each planets curve has two valleys and one peak. This seems to support the idea of us ing conjunction and opposition-class missions since the valleys correspond relatively to the 40 to 60-day stay opposition-class missions, and the longer 400 to 600-day stay of the conjunction-class missions.

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99 Next, the energy requirements for each planet we re assessed separately using a surface plot to see the affects of both transf er time and stay time on delta-V Figure 4-6 shows the results given a transfer time range of 100-300 days, and a 400-600 day stay time range. It appears the NTR system would have a very difficult time comp leting this mission at low transfer times that also have a short stay time, but will have a much simpler mission at all other transfer and stay times. Jupiters surface plot shows a delta-V asse ssment given transfer times of 1000 to 1200 days and stay times between 400 and 600 days. Th is plot, seen in Figure 4-7, shows much more variation in energy requirements across the transf er and stay time ranges along with a more even influence of either independent parameter on delta-V. Saturns contour plot in Figur e 4-8 reveals much smoother curves than that of Jupiters when plotted over the same transfer and stay tim e ranges. However, similar patterns are seen between the areas of maximum and minimum delta-V s. Jupiters optimum point for the specific departure date used is around 900 days for the tran sfer and 500 days for the stay time. Saturns optimal mission profile is around 1150 transfer da ys and 500 stay days for the same departure date as Jupiter. NEP Departure The NEP mission profile analysis plots were obtained just as they were for the NTR ephemeris model, with expected variations to be seen due to the different type of propulsion system. The delta-V versus departure year plot seen in Figure 4-9 shows both similarities and differences compared to the comparable NTR plot A similar synodical pa ttern is seen in the Mars curve, which has exactly six peaks, just as in the NTR curve. The Jupiter and Saturn curves, however, appear somewhat different, as only one peak and valley can be seen for each.

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100 This may be due to the fact that the NEP sy stem has less propulsive power and will be accelerating much slower to its de stination, thus increasing the pe riods of the curves of these more distant planets. The contour plot of delta-V vs departure month and year fo r Mars is shown in Figure 410. It can be seen that the optimal departure is in the early months of 2 031 and a bit later in the year 2033, representing the Earth-Mars 2.13 year s ynodic period. The delta-V values range from 40.2 to 99.7 (km/s) at these departure dates. The departure contour plot for Jupiter can be seen in Figure 4-11. Th is plot shows a much larger relative area of optimal energy requirement as compared to the Mars plot. This is in agreement to the earlier finding that there are more cycles over time in the Jupiter and Saturn curves versus Mars for the same range in years. This plot serves to portray the 1.09 year EarthJupiter synodic period, as similar delta-V minimu ms and maximums are repeated just over every 12 months. However, it is apparent from the high delta-V results ranging from 73.6 to 133.4 (km/s), that further optimization of mission parame ters would be needed to make this flight practical. The Saturn contour plot in Figure 4-12 shows th e trend seen in the J upiter plot even more clearly as much less variation is seen for delta-V requirements ove r the range in dates plotted. Most of the plot reveals an area of optimal depart ure dates, while only the lower left-hand corner shows departures dates of high delta-V. Again, the synodic cycle of 1.03 years can be seen as the delta-V values repeat just a bout every year. The delta-V valu es for the Saturn mission are the highest, ranging from 126.4 to 169.6 (km/s). Duration The dependence of delta-V on transfer time for the NEP missions can be seen in Figure 413. In comparison to the NTR plot, this plot shows fewer peaks in delta-V across the spectrum

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101 of transfer days. The Mars delta-V again sees a large increase after 400 transfer days, but the Saturn and Jupiter missions dont seem to be optimal until after their first peaks around 650 transfer days. Figure 4-14 shows a plot of delta-V versus st ay time. Although the NTR plot showed one prominent peak for each of the three curves, the NEP plot shows somewhat different profiles. Mars still has a central peak, but a much lower delta-V valley in the upper stay time range. Jupiter has two separate peaks in the NEP plot, and Saturn has a flattened spectrum with one low peak and one shallow valley. The surface plot in Figure 4-15 shows the Mars delta-V prof ile for the 100 to 300day transfer time range and 400 to 600-day stay ti me range. A more gradual slope downward as transfer time increases can be seen in th is plot versus that for the NTR system. Jupiters contour plot, which is seen in Fi gure 4-16 has more variation than found in the Mars plot. It only shows a small area of mini mum delta-V, primarily for missions with both long transfer and stay times. Saturns contour plot, seen in Figure 4-17, is similar to Jupite rs in that it has a peak area protruding close to the center of the 3-D mapping. It does, ho wever, show a more sharply negative slope with larger regions of bot h optimal and poor delta-V requirements. Hybrid Departure Although the Hybrid vehicle does utilize NTR bur sts at the beginning of each of its escape and capture maneuvers, the longer EP burns result in Hybrid ephemeris re sults similar to those seen in the previous NEP analysis section. Fo r instance, Figure 4-18 s hows a cyclical delta-V pattern for Mars with six peaks, just as in the NE P plot. Jupiters curve is also familiar, showing

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102 mostly a low valley region between two peaks, wh ile the Saturn curve is primarily flat, not having any apparent cyclical pattern. Looking at Figure 4-19, the cont our plot of delta-V vs. depa rture month and year for a Mars mission shows optimal departures accord ing to the 2.13 year synodic period near the beginning of 2031 and 2033. This was the same tr end found for the NEP vehicles, as could be expected. The range in delta-V values for the Hybrid system is between 37.9 and 127.5 (km/s), which has both higher and lower limits than the smaller NEP range, which was between 42.6 and 102.6 (km/s). The contour plot for a Jupiter mission, shown in Figure 4-20, also gives more desirable delta-V values than the NEP vehi cle. The delta-V range for the Hybrid system is between 52.8 and 146.8 (km/s), while that for the NEP system was between 113.6 and 171.1 (km/s). Just as in the NEP case, the Hybrid ephemeris plot shows th at an optimal departure can be obtained for at least one month out of every year in the period between 2040 and 2045. Saturns contour plot, shown in Figure 4-21, reiterates the fact that there is little fluctuation in delta-V over the five-year time period analyz ed. The optimal departure time for the Saturn missions is actually around mid-year each year between 2040 and 2045, being at a minimum delta-V between the 4th and 8th months. The range given for delta-V does not vary much from the NEP results, fluctuating between 120.0 and 158.2 (km/s) compared to the NEP range of 126.4 to 169.6 (km/s). Duration The Hybrid results for delta-V vs. transfer time, found in Figure 4-22, seem to provide the same general results as the NEP duration plot s. The Mars curve shows a minimum energy region before 400 transfer days, and only return s to these optimal delta-V values around 650 transfer days. Despite small disturbances in sl ope, both the Jupiter and Saturn curves show a

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103 continually downward trend in delta-V values until around 900 days, at which point the slopes begin to flatten out. Comparing delta-V vs. stay time results f ound in Figure 4-23 with the comparable NEP plot also shows very similar results between the two. The same patterns are seen for each curve, with the Mars delta-V curve dipping between 400 and 600 days, the Jupiter curve having two peaks at either end, and the Satu rn curve being relatively flat. The effects of transfer and stay time on delta -V for a Hybrid mission to Mars can be seen in Figure 4-24. A more variable dependence on stay time can be seen for this Hybrid vehicle compared to the NEP vehicle. This is shown in the red delta-V area found in the first half of the stay time range. The NEP and Hybrid vehicle pl ots also have the similarity that both show a minimum delta-V at high tran sfer and short stay times. Figure 4-25 shows much more variation in delt a-V with changing stay and transfer times as compared to the Mars mission. This Jupiter mission shows high delta-V values at low transfer and high stay times, but seems to find optimal delta -V values above transfer times of 1000 days. In comparison to the NEP Jupiter mission, this delta-V plot again has much more variation and also has a wider range of opportunities for optimal mission delta-Vs. The Saturn plot shown in Figure 4-26 reveals a surface plot with only minor folds. It is generally of a downward slope from low to high tr ansfer times, and is much more dependent on transfer time than stay time. This is shown as the optimal delta-V areas are at higher transfer times given any stay duration. Vehicle Architecture Analysis Parametric performance plots of the vehicle models over a range of delta-V requirements were also found in this trade study in order to assess the overall capabiliti es of the vehicles. IMLEO sizing plots based on delta-V values were used to reveal the response of each system

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104 model across a range of energy requirements. In this way, the system performance was revealed for any possible mission that would fall into the range of the ener gy parameter plotted, dismissing temporarily the focus on the specific design reference missions. In addition, these parametric plots may serve as points of validation for the system models to be compared against other models in previously completed or future studies. Parametric plots tailored to the design of each specific vehicle architecture were also generated in order to draw upon parameters that best characterize each system. For example, trades within the NTR category were on fuel type, while the trade was on electric thruster type for the NEP and Hybrid vehicles. The NTR results also include a plot of burn time, as this is a calculated value in the NTR vehicle model. Burn time, however, is an input to the NEP model, and was thus included as an independent variable in an IMLEO plot. The plot that was unique to the Hybrid model was an IMLEO surface plot that had both NT R and NEP delta-V requirements as independent variables. It was hoped that the results provided through the fuel and thruster trades, along with the vehicle-spec ific plots, help further reveal the capabilities of each of the vehicle configurations. NTR The main point of departure between the 5 NT R configurations was the type of nuclear fuel used in the reactor. The NT R vehicle architecture model was used in this analysis to assess any difference the different fuel types actually ma de on the overall size of the vehicle. Figure 427 reveals that there is little difference in the IMLEO between the fuel types for a total mission delta-V less than approximately 6 (km/s). This value, however, would still be less than the minimum required for a typical Mars mission. Above this delta-V value, it appears the CERM ET fuels perform the best until the carbide fuel outperforms the CERMET around a delta-V of about 10 (km/s). This finding is most likely

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105 due to the heavier weight of th e carbide fuel negating its perfor mance at low delta-Vs, but then its superior propulsive propertie s showing its benefits given hi gher mission energy requirements. The relatively low propulsive performance paramete rs of the graphite and composite fuels make these vehicles the heaviest for all delta-V values above 6 (km/s) A sharply increasing slope in the IMLEO curve for graphite is evident ar ound 11 (km/s) and for composite fuel around 12 (km/s), thus it is expected that they w ould reach their delta-V limits soon after. The IMLEO values for each fuel were also found for a singular mission to Mars that had a total delta-V of 8 (km/s). This allowed for the opportunity to view the IMLEO of each fuel type side-by-side, as seen in Figure 4-28. The IM LEO is broken down between the propellant mass and the inert mass to give an idea of the spl it between the two, and to compare the relative amount of propellant to total IMLEO between fuel t ypes. For instance, it can be seen that the graphite vehicle has a higher relative amount of propellant mass to total IMLEO than does the carbide fuel, evidence of the better fuel performance of the carbide reactor. Since the NTR model also calculated the burn ti me of the NTR at each of the escape and capture maneuvers, these values were plotted ve rsus fuel type in Figure 4-29. The delta-V values used to estimate these burn segments we re for an optimal Mars mission, with an Earth escape delta-V of 3.6 (km/s), plan etary capture of 1.4 (km/s), plan etary escape of 0.85 (km/s), and Earth capture of 1.46 (km/s). These burn times were directly calculated from the value of propellant for each burn, which is a function of de lta-V, thus the relative heights of the bars are consistent with those seen in the previous figur e of IMLEO versus fuel type. Although trajectory equations often use the assumption of instantaneous velocity changes at each burn, these results reveal that the NTR systems actually take betw een 0.5 and 3 hours for the completion of a single burn, with total mission burn times under 5 hours.

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106 One of the other utilit ies of the vehicle models was in assessing the breakdown of IMLEO between all of the major vehicl e components. As seen in Figure 4-30, the breakdown of the NTR graphite vehicle model for a singular Mars mission is in the following proportions: the reactor mass is 6% of IMLEO, fuel cell mass is 1%, TransHab and consumables are 16%, propellant is 55%, truss structure is 4%, RCS is 3%, and propellant tanks are 15%. Note that the reactor mass is actually a small percentage of the total IMLEO, and that the real effect of changing fuel type is in how th e fuel performance affects the si zing of the propellant mass. In addition, this percentage actually represents the weight of two reactors, where only one is being used for propulsion, with th e second for mere redundancy. NEP One of the dependent parameters unique to the NEP vehicle architecture model was the required thermal power. In the NTR and Hybr id models the maximum power was decided by the parameters of the fuel for the thermal rock et propulsive mode, and a ll were on the order of 320 (MWt). The actual thermal power required fo r the NEP system was much lower since it was only needed to provide electric power that drove the much lower-thrust electric thrusters. Figure 4-31 shows the required thermal power in (kWt) that is converted from the electric power requirement, and plots this power versus total mi ssion delta-V. A spread in power requirements between the thruster types can be seen above approximately 15 (km/s). The ion thruster maintains a modest positively sloped curve, whil e the Hall and MPD thrusters show a sharp rise in delta-V beginning around 30 (km/s). The MPD a nd Hall thruster curves continue to overlap until about 37 (km/s) at which point the Hall curve shows an even more drastic increase in slope. An examination of Figure 4-32 shows agreemen t between the effect of delta-V on IMLEO and on the power requirement, as similar trends can be seen. Above 30 (km/s), the IMLEO also begins to increase dramatically for the MPD and Hall thrusters. This se rves to show how the

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107 mission energy requirements based on trajectory not only affect the mass of the system, but also the power reactor requirements. It also asserts th at the NEP vehicle model begins to fail first for the Hall and MPD thrusters, followed by the ion thruster system, as mission demands increase. A specific Mars mission was also assessed for the NEP system in order to compare mass estimates based on the thruster selected. The opposition-class mission used for these data-points was based on transfer delta-Vs of 20 (km/s) each, and a burn time of 500 days in each direction. Figure 4-33 shows the very large difference in IM LEO between thruster types for a Mars mission of total delta-V equal to 40 (km/s) This is consistent with the previous plot, which asserts that the IMLEO spikes upwards for the Hall and MPD syst ems at large delta-V requirements. It also provides a glimpse of the results to be found fo r the tradespace matrix, which had missions to Jupiter and Saturn that both required much higher delta-Vs than the case shown here. In addition to the IMLEO dependence on delta-V, it was also desired to assess the level of dependence the IMLEO parameter had on the thru ster burn time. For the NEP system, less thrusters and less propellant we re needed to complete the sa me delta-V burn if it could be accomplished over a longer period of time. ModelC enter was thus used to generate a surface plot of the IMLEO versus burn time and delta-V, which can be seen in Figure 4-34. For the specific mission modeled in the run, it is found that the IMLEO has a 68% dependence on deltaV and a 32% dependence on burn time. These results are portrayed through th e plot and attest to the significant dependence on burn du ration for success of an NEP mission. Lastly, a vehicle mass breakdown based on a Mars mission was generated for the NEP architecture using the more effici ent ion thrusters. From Figure 4-35 it is found that the reactor comprises 5% of the total IMLEO, the thrusters are 9 % of the total, thermal components are 9%,

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108 TransHab and consumables are 6%, propellant is 57%, truss structure is 5%, RCS is 3%, and propellant tanks are 6%. Hybrid The Hybrid vehicle architecture was expected to perform similarly to the NEP architecture due to the nature of the trade on electric thruster type. Confirmation of this is seen in Figure 436, which shows that the thruster curves for IMLEO versus deltaV have the same trends that they did for the comparable NEP plot. The ion thruster outperforms the MPD and Hall thrusters by a significant margin. The main difference, how ever, is that at the same delta-V values, the Hybrid plot gives IMLEO values around two to th ree times those of the NEP system. This is confirmation that, in order for the Hybrid system to be competitive, it must have significantly lower delta-V requirements placed on the electric thruster system. This relief may be granted by having the NTR engines onboard, taking responsibi lity for a portion of the total mission energy requirements. A Mars conjunction-class mission was set up for the Hybrid vehicle in order to compare the IMLEO values of each thruster for a spec ific mission. The outbound and return NEP deltaVs were 14.0 and 11.8 (km/s), while those for the NTR were 3.17 and 0.22 (km/s). Figure 4-37 reveals a gradual increase in mass for the mission when going from ion thrusters, to MPD, and then to Hall thrusters. Compared to the NE P results, it is found here that this mission was accomplished with less mass for all three thruster cases. It can be seen from both the NEP and Hybrid plots, however, that there is a relativel y even split between propellant and inert mass for the ion system, but there is a higher proportion of propellant ma ss for the Hall thrusters, and a smaller proportion for the MPD thrusters in the Hybrid cases. This effect of the NTR and NEP engine re quirements on Hybrid vehicle performance can also be seen through the surface pl ot found in Figure 4-38. This pl ot shows the IMLEO of an ion

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109 thruster Hybrid vehicle plotted against both NE P and NTR delta-V requirements. It was found by repeating this analysis for vehicles with MPD and Hall thrusters th at the dependence on the NTR system did decrease. For example, the mission analyzed gave a 39% dependence on NTR delta-V vs. 61% dependence on NEP delta-V, while giving a 34% dependence for MPD vehicles and 33% for Hall thruster vehicles. A breakdown of IMLEO mass for a typical Mars mission is given for the Hybrid system in Figure 4-39. This mission is seen to be less de manding for the Hybrid system than for the NEP system, as a lower overall IMLEO is found. The difference results in lo wer propellant mass and lower thruster and thermal mass percentages, bu t generates higher percen tages for the propellant tanks and TransHab. Tradespace Assessment The utility of a tradespace assessment is in bei ng able to take a set of systems that could all be used to satisfy a given scenario or objective, and determine which system will best achieve the desired end-result. The culminating phase of this study was thus to assess the mission architecture tradespace matrix in order to determine the best vehicle systems for the reference mission profiles. In order to accomplish this, al l of the mission architecture models had to be run. This entailed optimizing the delta-V for each mission profile using the ephemeris models, and then running the vehicle architecture models with these optimized energy requirements. A metric had to then be derived to draw a conn ection between the vehicle architecture results and the overall ability of the vehicle sy stem to satisfy the mission objectives. This metric can be seen as a way to wrap all of a vehi cles attributes and performance pa rameters into one package that more simply represents the ability of the system to achieve the desired goal. This package was represented via a score that was used to ra nk each system in the desired categories.

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110 In deriving a score to use for the tradespace assessment, it was important to first establish the measures of effectiveness. These MOEs serve as the standard for identifying successful solutions, as they can only be measured against the accomplishment of a mission.48 Since they are often determined by the end-user of the proposed systems, MOEs can be seen as a wish-list of the desired system attributes. For instance, two different systems may be suggested to best fulfill a mission given two different sets of MOEs, since a different type of system performance may be desirable based on the attributes establis hed in each set. Deciding which system will best support a mission thus depends to some degree on th e priorities of those who will use the system. Although the MOEs are seen from the stakeholde rs perspective as qualitative attributes, a quantifiable figure of merit was still needed in order to provide an assessment of the MOEs. This is where the real systems engineering task ar ose, as separate metrics had to be created using quantitative Measures of Performa nce (MOPs) to assess each MOE. The score associated with each MOE was then used in a metric to derive on e score that determined the overall effectiveness of each system in completeing the assigned mi ssions. These numerical scores were then translated into a relative assessment of the sy stems, relieving an outsiders concern for the internal details of the solution.48 This allowed for a straight-forward assessment of the power and propulsion architectures based upon pre-determined needs, givi ng big-picture results of how well each system performed. Methodology Mission architecture analysis The term mission architecture has been previ ously defined as the application of a vehicle architecture configuration to the completion of a reference mission profile. Although the capabilities of both the epheme ris models and the vehicle ar chitecture models has been previously explored, a discussion of their use in modeling specific missions has been saved until

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111 this point. It was the assessment of the mission architectures that was needed to generate the performance parameters required for assi gning scores in the tradespace matrix. The first step in analyzing each mission arch itecture was to gene rate optimal delta-V energy requirements from the ephemeris models This optimization was achieved through the use of the Darwin optimizer t ool, which optimized a mission give n a range on departure year and month, transfer time, and stay time at the planet These ranges came directly from the reference mission profiles found in the tradespace matrix. It was pivotal that the optimal mission be found for each propulsion and power system so that on e system would not have an unfair advantage over the other when comparing their results in the tradespace analysis. Th is is in essence why three ephemeris models were created for each of the three main vehicle architectures, since the models had to be able to handle th e specifications of each unique system. Due to the inclusion of the Darwin optimiza tion tool in the epheme ris models, the mission optimization procedures became somewhat automa ted, reducing overall analysis time. The first step after opening one of the ephemeris models and inputting the destin ation was to update the stay time, transfer time, and departure month and date upper and lower bounds. These appeared as inputs in the ModelCenter GUI that employed simple pull-down selection menus so as to warrant off careless mistakes in inputs. The next step was to run the Darwin utility component that appeared as an icon in the ModelCenter viewing window. Once the utility stopped running after an optimization time of approximately 3-4 minutes, the output section in the GUI contained the resultant delta-V value of th e optimal mission given the paramete rs that were set by the user. The snapshot tool was then select ed and a listing in a cell-structur ed format similar to Excel was brought up on the screen allowing simple highlighti ng of the desired cells. The parameters that were selected to be included in the missi on architecture database included the optimized

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112 departure month and year, stay tim e, transfer time, delta-V, and also the burn time and spiral time for the electric thruster systems. This proc edure then had to be repeated for the remaining mission profiles. Note that the same optimized mission profile parameters were used for all configurations within the NTR, NEP, and Hybrid systems since energy requirements did not vary based on fuel and thruster type. The set of optimized mission requirements fo r each of the 12 missions corresponding to each of the 3 vehicles were then put into the appropriate vehicle model, and were run for all configurations. Since the optimization steps had been previously completed, each of the missions took only seconds to run. The immediate result to be noted was whether or not the vehicle was able to complete the mission under the 1000 (t) IMLEO limit discussed in the next section. If it did not, all other results were discarded. Missions that did satisfy this requirement resulted in recording of th e vital vehicle parameters. Determining MOE scores In order to be able to assess which system best satisfied each mission, a metric able to handle the mission architectur e results had to be defined. The MOE scores were used to rank the vehicles for each mission, for each planet, and fo r overall interplanetary travel. The 132 MOE scores appearing in the tradespace matrix were thus used to find the best system for each of the 12 mission profiles, for each of the 3 planetary de stinations, and for one general interplanetary mission. These categories were broken down in orde r to represent different levels of flexibility in system design based on the required missions that engineers may need to design for in the future. Although multiple missions were analyzed in th e tradespace matrix, the same five MOEs were used to assess the systems for each refe rence mission profile. Based upon the capabilities

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113 of the engineering models developed for this study, along with important measures of success found in past mission analysis studie s, the MOEs were established as: 1. Mission completion 2. Launch costs 3. Crew safety 4. Vehicle complexity 5. Mission control operational costs By designating these five attributes as the MO Es, they alone were used to determine the relative effectiveness of each sy stem in satisfying the requir ements of each mission scenario. The first MOE, Mission Completion, was a simple assessment of whether or not the vehicle architecture model was able to run give n the ephemeris energy requirements. For some of the more energy-intensive missions, it was ex pected that the less pow erful propulsion systems, notably the NEP systems, would not perform ad equately. This could be seen by running the model and seeing if the IMLEO blo ws up, or stated mathematically, goes towards infinity. As most interplanetary mission analysis done in the past has reported IMLEO estimates in the range of 100 to 500 (t), it was decided to set a mode st cutoff limit of 1000 (t) for a vehicle to be considered to have completed the mission. A lthough the other MOE scores were based upon a percentage out of 100%, the score for MOE1 was simply either a for a system weighing less than 1000 (t), or a for a system weighing greate r than 1000 (t). Note that a would nullify the results from the remaining four MOEs for th at specific mission architecture, giving a 0% as the total MOE score. The second MOE was Launch Cost, an attr ibute that was entirely based upon the IMLEO of the system to complete a mission. This was d ecided given the fact that most of the cost of

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114 launching a spacecraft into orbit is based upon th e mass of the useful sp acecraft hardware and propellant that must get to orbit. Total la unch cost estimates have been given around $2000/kg, where the actual launch vehicle launch co st will be about 60 % of this value.48 In order to generate a MOE2 metric that gave a score as a percenta ge, a ratio of the IMLEO to the maximum possible IMLEO of 1000 (t) was used. This metric can be seen by Equation 4-1, where IMLEO is given in units of kilograms and is thus di vided by 1000 in order to provide a ratio to the 1000 (t) limit. 1000 1000 / 12IMLEO MOE (4-1) The third MOE of Crew Safety was focused on how long the crew would remain in the transfer orbit within the tight constraints of the TransHab and subject to large radiation doses. A metric for crew safety was derived based upon th e actual number of transfer days the crew was subject to (days transfer _), the minimum number of days the model allowed for (min transfer), and the range of transfer days the model considered (range transfer _). This metric is given by Equation 4-2, where the minimu m transfer was set as 100 days for Mars and 800 days for Jupiter and Saturn, and the range was 200 days for Mars and 400 days for Jupiter and Saturn. range transfer transfer days transfer MOE min / 13 (4-2) The least quantifiable MOE was the fourth in the list: Vehicle Complexity. The complexity of the vehicle really came down to th e choice of fuel for the NTR configurations and the electric thruster type for the NTR and Hybrid configuratio ns. The complexity of these systems was thus given by the rela tive technology readiness level (T RL) of the different fuel and thruster types, as perceived by this re searcher. The TRL was assessed based upon a

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115 comprehensive understanding of past research an d development, and current versus expected performance parameters of all the fuel and thrust er options. A qualitative assessment of the TRL for each of the 11 systems was then equated with a percentage, resulting in each of the scores found in Table 4-1. Note that the Hybrid c onfiguration scores were based upon the product of the BNTR CERMET fuel score and the electric thruster type score. The final MOE established was Mission Control Operational Costs. Just as the Launch Cost MOE had to be related to something othe r than a dollar amount, the mission control costs also had to be related to a figure of merit from da ta analysis results that would relate to cost in the real world. It was decided to use the duration (in days) of the mission for this figure of merit. For each day that the spacecraft was being operated, mission control had to be available to control this operation and assess spacecraft hea lth. The number of mission days included the days of interplanetary transfer, the stay time in which the spacecraft was in orbit at the destination planet, and for the case of the NE P vehicles, the duration of spiral escape and capture. This MOE is also significant to mi ssion planners because although lower IMLEOs may be achieved from higher efficiency engines, thes e same engines may have to be in operation longer, equating to increased hardware and maintena nce costs. The metric for this MOE is given by Equation 4-3, where max mission is an estimated value of the maximum possible duration of a mission, and was determined to be 2000 days for a Mars mission and 4000 days for a Jupiter or Saturn mission. max _ 15mission days mission MOE (4-3) Tradespace Mission Architecture Results In order to be able to perform the analysis of the tradespace matrix previously discussed, all of the parameters that were needed for the MOE metric had to be a ppropriately book-kept.

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116 Ephemeris parameters that were included in the database for each of the 132 mission architectures included delta-V, transfer time, stay time, burn time, and spiral time. The major performance parameters from the vehicle architect ure models that were also included were the IMLEO, propellant mass and operating power. All of these results can be found in Appendix B: Tradespace Mission Architecture Results. The only vehicle performance parameter that was plotted based on the mission architecture analysis results was the IMLEO. It was felt that the IMLEO was the one true performance parameter that could be directly compared between systems and would provide useful information. For instance, the thrust to weight parameter that is typically used to define propulsion system performance could not be used to determine the best system out of the three main architectures since it was already known that electric thru sters provided a much smaller thrust to weight than thermal rockets. On the other hand, the ISP, or measure of thrust efficiency, is much higher for the electric propuls ion systems than for the thermal rockets. Thus, on a parameter -to-parameter level, it is only trul y useful to get a glimpse of the overall system performance by comparing the direct weight differences of the systems by their IMLEO values. The results of running each of the mission archit ectures in the tradesp ace matrix revealed the result that none of th e three architectures could complete the Jupiter or Saturn missions under the designated IMLEO limitation of 1 million (kg). Within the Mars reference missions, only the Hybrid vehicle with ion thrust ers could complete an opposition-class mission, doing so with an IMLEO of 472 (t). Nine of the eleven vehicle configurations were able to complete the fast transfer conjunction-class Mars mission, and all eleven vehicles completed the slow transfer conjunction class-mission. Sin ce the only significant results come from a comparison of

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117 completed Mars missions, the results from these missions are presented si de-by-side in Figure 440. In examining this figure, it is seen that there is very little difference between the slow and fast transfer conjunction IMLEO re sults for the NTR configurations. This results from the very similar delta-V energy requirements for both missi ons, despite the longer transfer times allowed for the slow transfer missions. The Hybrid and NEP configurations are much more affected by the transfer durations, as shown by the much higher IMLEO values for the fast transfer conjunction missions. It can be s een in the figure that two of the NEP configurations couldnt even complete these missions with a reasonabl e mass. Only the Hybrid ion vehicle could complete the opposition slow transfer mission. These results reinforce the idea that NEP and Hybrid vehicles will be much more affected by tran sfer time as compared to NTR vehicles due to their much lower thrust capabilities. Tradespace MOE Results After the mission architectures were run, the MOE scores were compiled for the tradespace matrix. These scores all appear in Table 4-2. The mission architectures that have a 0% MOE score are those that could not be co mpleted, as previously mentioned. The MOE scores for the completed Mars missi ons, which represent the overall capability of each vehicle configuration to co mplete the desired mission, can be seen plotted in Figure 4-41. The scores for these missions range from 40. 67% for a Hybrid ion vehicle opposition-class mission, to a high of 68.67% for a BNTR CERMET vehicle conjunc tion-class mission. Although the IMLEO results for the NTR configurations were similar, the MOE scores are much higher for the faster transfer missions, as compared to the slower missions. This contests that other important concerns beside s IMLEO, such as the amount of time astronauts spend in space, become important factors when assessing mission architectures. The Hybrid and NEP results

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118 reveal that the fast transfers are better fo r the ion thruster systems, but MPD and Hall configurations would perform better for slow transfers. It can also be seen that the Hybrid ion configuration is robust, in that it completes the most missions, but the NTR vehicles have higher scores for specific missions. Based upon these MOE score results, the syst ems were ranked in order of decreasing performance from 1 to 11. Table 4-3 gives th e rankings for each vehi cle based on each of the four Mars missions. The Hybrid ion configuration has been gi ven the rank of for the opposition-class mission since it was the only vehicle that successfully completed this mission. For the conjunction-class missions, it is seen that the NTR CERMET, carbide, and composite configurations rank higher than the Hybrid ion configuration. Since there were no Jupiter or Saturn missions completed, the planet-specific and overall rankings were exactly the same, with the overall rankings being the same as those for the Mars missions. Table 4-4 gives these results, wherein the MOE scores across the board were averaged and then ranked. It can be seen that the Hybr id ion vehicle takes the highest ranking, followed closely by the CERMET, carbide, composite, and gr aphite NTR configurations, respectively. The NEP configuration with ion thrusters obtained the next hi ghest ranking, followed by the Hybrid and then the NEP MPD and Hall thruster configurations. By comparing the results from the mission-b ased and overall rankings, it can be gleaned that the same vehicle configura tion will not necessarily excel for all types of missions. Although the Hybrid ion configuration received the top ranking for overall Mars missions, it would probably not be desirable to focus on this one vehi cle for all types of Mars missions. It is clear that most of the NTR configurations actually outperform the Hybrid ion vehicle for the Mars conjunction-class missions, with the fast-trans fer missions revealing much higher MOE scores

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119 for the NTR systems. This seems to provide motivation for focusing more on specific missions to be performed, versus a specific vehicle to be used, when trying to determine the best power and propulsion system for interplanetary travel. When examining these results it should be remembered that many of the vehicle configurations, most notably thos e with EP thrusters, were de signed based on projected future operating parameters. These vehicles were co mpared to NTR configurations given current operating parameters that have already been proven in full-scale testing. In addition, all of the vehicle models in this study were very sensitive to ISP or efficien cy values due to the nature of sizing propellant mass based on the rocket equation. It can thus be seen that the highest-ranking Hybrid ion vehicle may have appeared elsewhere in the ranking list if its ion thrusters were simply assigned less optimistic parameter values. This highlights one of the major difficulties in giving an honest comparison between power a nd propulsion systems, as preliminary system definition and vehicle design play a major role in how the results will stack up.

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120 Table 4-1. MOE4 TRL Score System/Technology MOE4 Score NTR/Graphite fuel 0.90 NTR/CERMET fuel 0.80 BNTR/CERMET fuel0.80 NTR/Composite fuel 0.85 NTR/Carbide fuel 0.80 NEP/Ion 0.70 NEP/MPD 0.60 NEP/Hall 0.50 Hybrid/Cermet+Ion 0.8*0.7= 0.56 Hybrid/Cermet+MPD0.8*0.6= 0.48 Hybrid/Cermet +Hall0.8*0.5= 0.40 Table 4-2. Tradespace MOE Scores Destination MARS Jupiter Saturn Stay Conjunc. Oppos. Conjunc. Oppos. Conjunc. Oppos. Transfer Fast Slow FastSlow Fa stSlowFastSlowFast Slow FastSlow NTR (Graphite) 66.65% 54.38% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NTR (Composite) 67.63% 55.40% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NTR (Carbide) 68.01% 55.81% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NTR (CERMET) 68.45% 56.24% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% BNTR (CERMET) 68.67% 56.49% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% HYBRID (ION) 60.06% 54.81% 0% 40.67% 0% 0% 0% 0% 0% 0% 0% 0% HYBRID (MPD) 52.69% 54.79% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% HYBRID (Hall) 44.78% 52.29% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NEP (ION) 56.78% 52.01% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NEP (MPD) 0.00% 47.27% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0% NEP (Hall) 0.00% 42.17% 0% 0% 0% 0% 0% 0% 0% 0% 0% 0%

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121 Table 4-3. Tradespace Mission-Based Rankings Destination MARS Stay Conjunc. Oppos. Transfer FastSlowFastSlow NTR (Graphite) 5 7 NTR (Composite)4 4 NTR (Carbide) 3 3 NTR (CERMET) 2 2 BNTR (CERMET)1 1 HYBRID (ION) 6 5 1 HYBRID (MPD) 8 6 HYBRID (Hall) 9 8 NEP (ION) 7 9 NEP (MPD) 10 NEP (Hall) 11 Table 4-4. Tradespace Planet -Based and Overall Rankings Destination Mars NTR (Graphite) 6 NTR (Composite) 5 NTR (Carbide) 4 NTR (CERMET) 3 BNTR (CERMET)2 HYBRID (ION) 1 HYBRID (MPD) 8 HYBRID (Hall) 9 NEP (ION) 7 NEP (MPD) 10 NEP (Hall) 11

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122 Figure 4-1. ModelCenter Parametric Tools 0 10 20 30 40 50 60 70 80 90 100 20302032203420362038204020422044 Departure YearDelta-V (km/s) Mars Jupiter Saturn Figure 4-2. NTR Delta-V vs. Departure Year

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123 Figure 4-3. NTR Mars Delta-V vs. Departure Year and Month

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124 0 25 50 75 100 125 150 175 200 225 250 275 300 200300400500600700800900100011001200 Transfer Duration (days)Delta-V (km/s) Mars Jupiter Saturn Figure 4-4. NTR Delta-V vs. Transfer Time 0 10 20 30 40 50 60 70 80 90 40100160220280340400460520580 Stay Time (days)Delta-V (km/s) Mars Jupiter Saturn Figure 4-5. NTR Delta-V vs. Stay Time

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125 Figure 4-6. NTR Mars Delta-V vs. Transfer and Stay Time

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126 Figure 4-7. NTR Jupiter Delta-V vs. Transfer and Stay Time

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127 Figure 4-8. NTR Saturn Delta-V vs. Transfer and Stay Time

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128 40 50 60 70 80 90 100 110 120 130 140 20302032203420362038204020422044 Departure Year Delta-V (km/s) Mars Jupiter Saturn Figure 4-9. NEP Delta-V vs. Departure Year

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129 Figure 4-10. NEP Mars Delta-V vs. Departure Year and Month

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130 Figure 4-11. NEP Jupiter Delta-V vs. Departure Year and Month

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131 Figure 4-12. NEP Saturn Delta-V vs. Departure Year and Month

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132 25 50 75 100 125 150 175 200 225 250 275 200300400500600700800900100011001200 Transfer Time (days)Delta-V Mars Jupiter Saturn Figure 4-13. NEP Delta-V vs. Transfer Time 40 60 80 100 120 140 160 180 40100160220280340400460520580 Stay Time (days)Delta-V (km/s) Mars Jupiter Saturn Figure 4-14. NEP Delta-V vs. Stay Time

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133 Figure 4-15. NEP Mars Delta-V vs. Transfer and Stay Time

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134 Figure 4-16. NEP Jupiter Delta-V vs. Transfer and Stay Time

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135 Figure 4-17. NEP Saturn Delta-V vs. Transfer and Stay Time

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136 20 40 60 80 100 120 140 160 20302032203420362038204020422044 Departure YearDelta-V (km/s) Mars Jupiter Saturn Figure 4-18. Hybrid Delta-V vs. Departure Year

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137 Figure 4-19. Hybrid Mars Delta-V vs. Departure Year and Month

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138 Figure 4-20. Hybrid Jupiter DeltaV vs. Departure Year and Month

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139 Figure 4-21. Hybrid Saturn Delta-V vs. Departure Year and Month

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140 0 50 100 150 200 250 300 350 20040060080010001200 Transfer time (days)Delta-V (km/s) Mars Jupiter Saturn Figure 4-22. Hybrid Delta-V vs. Transfer Time 40 60 80 100 120 140 160 180 40100160220280340400460520580 Stay Time (days)Delta-V (km/s) Mars Jupiter Saturn Figure 4-23. Hybrid Delta-V vs. Stay Time

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141 Figure 4-24. Hybrid Mars Delta-V vs. Transfer and Stay Time

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142 Figure 4-25. Hybrid Jupiter DeltaV vs. Transfer and Stay Time

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143 Figure 4-26. Hybrid Saturn Delta-V vs. Transfer and Stay Time

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144 0 100 200 300 400 500 600 700 800 900 0123456789101112 Total Delta-V (km/s)IMLEO (tons) Graphite Composite Cermet-unimodal Cermet-bimodal Carbide Figure 4-27. NTR IMLEO vs. Delta-V 0 50 100 150 200 250 300 350graphcompcarbcerm-unicerm-biMass (tons) Inert Mass Propellant Mass Figure 4-28. NTR Mass Breakdown vs. Fuel Type

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145 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 graphcompcarbcerm-unicerm-biBurn Time (hrs) Earth Capture Planet Escape Planet Capture Earth_Escape Figure 4-29. NTR Burn Time vs. Fuel Type 6% 1% 16% 55% 4% 3% 15% mass_reactor mass_fuel_cells mass_transhab+consum mass_propellant mass_truss mass_RCS mass_prop_tanks Figure 4-30. NTR Vehicle Mass Breakdown

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146 0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000 0510152025303540 Delta-V (km/s)Power_req (kWt) MPD ION HALL Figure 4-31. NEP Power vs. Delta-V 0 100 200 300 400 500 600 700 800 900 1000 05101520253035404550Delta-V (km/s)IMLEO (tons) MPD_IMLEO ION_IMLEO Hall_IMLEO Figure 4-32. NEP IMLEO vs. Delta-V

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147 0 100 200 300 400 500 600 700 800 IONMPDHallMass (tons) Inert Mass Propellant Mass Figure 4-33. NEP Mass Brea kdown vs. Thruster Type

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148 Figure 4-34. NEP IMLEO vs. Burn Time and Delta-V

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149 5% 9% 9% 6% 57% 5% 3% 6% mass_reactor mass_thrusters mass_thermal mass_transhab+consum mass_propellant mass_truss mass_RCS mass_prop_tanks Figure 4-35. NEP Vehicle Mass Breakdown 0 100 200 300 400 500 600 700 800 900 1000 05101520253035Delta-V (km/s)IMLEO (tons) MPD_IMLEO ION_IMLEO Hall_IMLEO Figure 4-36. Hybrid IMLEO vs. Delta-V

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150 0 50 100 150 200 250 IONMPDHallMass (tons) Inert Mass Propellant Mass Figure 4-37. Hybrid Mass Breakdown vs. Thruster Type

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151 Figure 4-38. Hybrid IMLEO vs. NEP and NTR Delta-V

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152 5% 4% 5% 15% 49% 5% 3% 14% mass_reactor mass_thrusters mass_thermal mass_transhab+consum mass_propellant mass_truss mass_RCS mass_prop_tanks Figure 4-39. Hybrid Vehicle Mass Breakdown N TR ( G r a p h i t e ) N T R ( C o m p o s i t e ) N T R ( C a r b i d e ) N TR ( C E R M E T ) B N TR ( C E R M E T) H Y B R I D ( I O N ) H Y B R I D ( M P D ) H Y B R I D ( H a l l ) NE P ( I O N ) N E P ( M P D ) N E P ( H a l l ) Opp_Slow Conj_Fast 100 150 200 250 300 350 400 450 500 550 600 650 700IMLEO (tons) Opp_Slow Conj_Slow Conj_Fast Figure 4-40. Mars Mission IMLEOs

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153 N T R ( G r a p hi t e ) N TR ( C o m p os i t e ) N T R ( C a r b i d e ) N T R ( C E R M E T ) B N T R ( C E R M E T ) H Y B R I D ( I O N ) H Y B R I D ( M P D ) H Y B R I D ( H a l l )N E P ( I O N ) N E P ( M P D ) N E P ( H a l l ) Opp_Slow Conj_Fast 30% 35% 40% 45% 50% 55% 60% 65% 70% 75% 80%MOE Score Opp_Slow Conj_Slow Conj_Fast Figure 4-41. Tradespace MOE Scores for Mars Missions

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154 CHAPTER 5 CONCLUSION Study Contributions One of the primary contributions of this th esis was to produce a mission analysis trade study with results that were on an order of ma gnitude not seen before in the short list of interplanetary mission analysis studies. This trade study comp ared three different vehicle architectures that were all based upon nucle ar reactor power sources and broke these architectures down further into configurations ba sed upon reactor fuel or el ectric thruster type. In addition to the eleven vehicl e configurations that resulted from this framework, missions beyond Mars, to Jupiter and even Saturn were at tempted. These destinations were aimed for while maintaining a sense of awareness for the duration that astronauts may safely travel in space. A secondary contribution was to examine NEP, NTR, and Hybrid vehicles in a way that was unbiased, essentially analyzing each system si de-by-side when tasked to perform the same type of mission. This contribu tion is significant as it is difficult to find any literature source that compares one architecture with another. This is a real challenge in industr y because an expert in one area will no doubt be an advocate for the type of vehicle that falls into his/her area of expertise. Creating a trade study from a mission analysis study and using a tradespace matrix was in itself another significant contributio n to this area of research. The tradespace matrix served to represent high-level requirements that may be given in the future to interplanetary mission planners. More often than not, the decision on which technology to c hoose to satisfy these requirements will not lie in the hands of an engi neer with the technical knowledge of how these systems work. Therefore, it is paramount that the engineer be able to analyze the systems

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155 capabilities in meeting specific requirements, and th en present this analysis in a simple and often qualitative manner. The addition of a systems engineering tradesp ace methodology was thus used in this thesis to help draw straightforward, unbiased, and mean ingful results as to which type of power and propulsion system may be most suitable for manned interplanetary missions. Unfortunately, the results fail to give any vehicle c onfiguration that could complete a Saturn or Jupiter mission. It is not believed that these results signify a failure in the models developed for this study. They are simply a reflection of the great demands in en ergy required to allow hu manity to explore our vast solar system. It is thus believed at this time that new analysis cannot solve this problem without the development of new physical c oncepts for reaching these destinations. It is important to note that the results of th is study do not provide a veritable best answer to the question of which propulsi on and power system should be us ed for crewed interplanetary missions. This analysis instead serves to prov ide a comparison of specific system designs based on estimated operating parameters, applied to a ra nge of reference missions. One must be aware of the fact that the IMLEO and MOE scores foun d for these systems were very sensitive to the selected operating parameters. Changing the value of an important system parameter such as ISP could entirely change the results, revealing a much different ranking matrix. This is an especially vital point, since the selected operating parameters were a mix of current and projected values. It is thus reit erated that the results of this trade study are meant only to reveal the performance of specific vehicle models, and do not asse rt that a specific power and propulsion system is better or worse than another. Areas of Future Work The results found in this study only open the door further for new work to be done in both the areas of mission analysis and in advanced nuclear and aerospace technologies. Although this

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156 study allowed for increased breadth and depth of mission analysis as compared to any study found through an in-depth literature search, it is important that mission designs be even further optimized based on the unique capabilities of each power and propulsion system. Although optimization of mission parameters was a key utili ty in this study, it was done under the umbrella of maintaining homogeneity in mission design be tween the vehicle architectures. Thus, energysaving maneuvers such as planetary swingbys and Earth flybys upon return were not used in this study. For instance, the NTR architecture ha d trouble completing the Mars opposition-class missions, yet it is common in mission analysis st udies to lower delta-V through a Venus swingby gravity assist. It will be importa nt, however, to incorporate such things into future analysis while stemming any bias towards one power and propulsion system. The technology readiness level of many of the technologies needed for interplanetary travel must also see a sharp increase before thes e missions become a reality. In the EP field, special attention should be paid to power conve rsion units to allow for higher electric power draw to electric thrusters while minimizing weight of these units. The thruster power processing units must also be improved, as they are the h eaviest part of the electric thruster hardware component. In the NTR field, reactor fuel testi ng must also continue in order to increase ISP while remaining below temperature limits of the materials being used. Nuclear thermal rocket testing must be reinstated before the resu lts and know-how from the 1960s and 1970s are completely lost upon new engineers. Lastly, spec ial attention should be paid to developing the complex, but achievable bimodal reactors, which performed strongly in this trade study, and must be advanced before Hybrid vehicles may come to fruition. Final Comments Despite the findings in this st udy that nearly half of the mi ssions to Mars and no missions beyond Mars could be completed with the advanced technologies analyzed herein, there is no

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157 doubt in the mind of this author that any of these missions coul d be achieved in the future. Reasonable results for some Mars opportunitie s were found given power and propulsion system hardware that was being tested nearly 40 years a go. Now, due to limited financial resources and political motivation, preparation for these Mars missions has been constrained to mostly paper studies. These studies do, however, keep alive the motivation to make these interplanetary missions a reality. This research was performe d with that same spir it of hope that someday scientists and engineers will be given the opportunity to make the dream of interplanetary travel come true.

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158 APPENDIX A LIST OF ABBREVIATIONS ADCS Attitude Determination and Control System BNTR Bimodal Nuclear Thermal Rocket C&DH Command and Data Handling CERMET CERamic-METallic; Uranium-fuel ed/refractory metal matrix reactor core CHEBYTOP Chebyshev Trajectory Optimization Program cSv Centi-Sievert CTV Crew Transfer Vehicle DOE Department of Energy DRM Design Reference Mission ECRV Earth Crew Return Vehicle EP Electric Propulsion FEEP Field Emission Electric Propulsion GCR Galactic Cosmic Radiation GUI Graphical User Interface HIVHAV High Voltage Hall Accelerator HLLV Heavy Lift Launch Vehicles IMLEO Initial Mass in Low Earth Orbit IPREP Interplanetary PREProcessor IPS Integrated Propulsion Systems ISP Specific Impulse ISS International Space Station JIMO Jupiter Icy Moons Orbiter

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159 K Kelvin kg Kilogram klbf Kilo-Pound-Force km Kilometer kWe Kilo-Watt Electric LEO Low Earth Orbit m Meter MOE Measure of Effectiveness MOP Measure of Performance Mpa MegaPascal MPD Magneto-Plasma-Dynamic MPDT Magneto-Plasma-Dynamic Thruster MSFC Marshall Space Flight Center MW Mega-Watt (Thermal) N Newton NASA National Aeronautics a nd Space Administration NDR NERVA Derivative Reactor NEP Nuclear Electric Propulsion NERVA Nuclear Engine for Rocket Vehicle Applications NTR Nuclear Thermal Rocket PMAD Power Management and Distribution PPU Power Processing Unit psia Pounds per Square Inch, Absolute (referenced to a vacuum)

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160 RCS Reaction Control System s Second SEP Solar Electric Propulsion SOI Sphere of Influence SPE Solar Particle Events t Ton TRL Technology Readiness Level TT&C Tracking, Telemetry, and Command T/W Thrust to Weight

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161 APPENDIX B TRADESPACE MISSION ARCHITECTURE RESULTS Table B-1. NTR Optimized Ephemeris Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) Departure Day 3 2 2 2 Departure Month 4 1 9 1 Departure_Year 2033 2031 2034 2033 Transfer (days) 200 294 199 243 Stay (Days) 600 443 46 46 Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000) S(1000-1200) Departure Day 1 24 14 4 Departure Month 1 12 1 1 Departure_Year 2041 2040 2041 2041 Transfer (days) 1000 1200 1000 1200 Stay (Days) 480 461 60 60 Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000) S1000-1200) Departure Day 1 1 14 2 Departure Month 1 1 1 1 Departure_Year 2041 2041 2041 2041 Transfer (days) 1000 1200 1000 1200 Stay (Days) 475 461 60 60

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162 Table B-2. NEP Optimized Ephemeris Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) Departure Day 24 6 25 15 Departure Month 3 5 3 3 Departure_Year 2033 2035 2033 2033 Transfer (days) 198 296 200 300 Stay (Days) 500 464 40 52 Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) Departure Day 16 3 27 23 Departure Month 8 9 10 11 Departure_Year 2040 2042 2043 1044 Transfer (days) 1000 1200 984 1191 Stay (Days) 400 520 41 54 Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) Departure Day 27 12 15 25 Departure Month 9 8 9 8 Departure_Year 2040 2041 2041 2040 Transfer (days) 992 1197 1000 1097 Stay (Days) 445 519 44 40

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163 Table B-3. Hybrid Optimized Ephemeris Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) Departure Day 5 21 27 13 Departure Month 4 1 1 12 Departure_Year 2033 2033 2031 2030 Transfer (days) 182 297 200 298 Stay (Days) 587 488 400 40 Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) Departure Day 15 15 10 21 Departure Month 9 8 11 10 Departure_Year 2041 2041 2042 2043 Transfer (days) 993 1197 987 1042 Stay (Days) 490 543 51 56 Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) Departure Day 22 24 14 29 Departure Month 5 10 8 9 Departure_Year 2040 2040 2040 2040 Transfer (days) 1000 1000 948 1000 Stay (Days) 400 596 404 599

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164 Table B-4. NTR Graphite Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 287057.8 294324.4 NA NA Mass_Inert 114294.6 116823.1 NA NA Mass_Propellant 152117.5 156855.7 NA NA delta-V1 3.5 3.6 10.8 4.2 delta-V2 1.4 1.4 7.7 2 delta-V3 1.2 0.9 1.1 3.9 delta-V4 0.9 1.5 0.9 4.7 Crewed_days 1000 1031 NA NA MOE_Score 66.65% 54.38% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S(1000-1200) delta-V1 6.2 6.4 22.8 7.9 delta-V2 5.5 5.6 20.6 5.8 delta-V3 7.3 8.3 9.8 7.7 delta-V4 22.7 25.6 6.7 23 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S1000-1200) delta-V1 9.2 8.9 11.1 9.1 delta-V2 9.3 7 9.1 7 delta-V3 10 7.5 9.6 7.3 delta-V4 6.2 5.4 8.9 6.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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165 Table B-5. NTR Compos ite Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 252334.7 258895.6 NA NA Mass_Inert 105397 107729.1 NA NA Mass_Propellant 129039.9 133268.8 NA NA delta-V1 3.5 3.6 10.8 4.2 delta-V2 1.4 1.4 7.7 2 delta-V3 1.2 0.9 1.1 3.9 delta-V4 0.9 1.5 0.9 4.7 Crewed_days 1000 1031 NA NA MOE_Score 67.63% 55.40% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 6.2 6.4 22.8 7.9 delta-V2 5.5 5.6 20.6 5.8 delta-V3 7.3 8.3 9.8 7.7 delta-V4 22.7 25.6 6.7 23 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 9.2 8.9 11.1 9.1 delta-V2 9.3 7 9.1 7 delta-V3 10 7.5 9.6 7.3 delta-V4 6.2 5.4 8.9 6.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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166 Table B-6. NTR Carbide Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 229760 235744.4 NA NA Mass_Inert 98681.7 100853.3 NA NA Mass_Propellant 111622.7 115435.4 NA NA delta-V1 3.5 3.6 10.8 4.2 delta-V2 1.4 1.4 7.7 2 delta-V3 1.2 0.9 1.1 3.9 delta-V4 0.9 1.5 0.9 4.7 Crewed_days 1000 1031 NA NA MOE_Score 68.01% 55.81% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 6.2 6.4 22.8 7.9 delta-V2 5.5 5.6 20.6 5.8 delta-V3 7.3 8.3 9.8 7.7 delta-V4 22.7 25.6 6.7 23 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 9.2 8.9 11.1 9.1 delta-V2 9.3 7 9.1 7 delta-V3 10 7.5 9.6 7.3 delta-V4 6.2 5.4 8.9 6.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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167 Table B-7. NTR CERMET-Unim odal Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 220997.8 227047.7 NA NA Mass_Inert 99218.5 101408.4 NA NA Mass_Propellant 113014.7 116874.8 NA NA delta-V1 3.5 3.6 10.8 4.2 delta-V2 1.4 1.4 7.7 2 delta-V3 1.2 0.9 1.1 3.9 delta-V4 0.9 1.5 0.9 4.7 Crewed_days 1000 1031 NA NA MOE_Score 68.45% 56.24% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S(1000-1200) delta-V1 6.2 6.4 22.8 7.9 delta-V2 5.5 5.6 20.6 5.8 delta-V3 7.3 8.3 9.8 7.7 delta-V4 22.7 25.6 6.7 23 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S1000-1200) delta-V1 9.2 8.9 11.1 9.1 delta-V2 9.3 7 9.1 7 delta-V3 10 7.5 9.6 7.3 delta-V4 6.2 5.4 8.9 6.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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168 Table B-8. NTR CERMET-Bim odal Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 216602.9 222076.8 NA NA Mass_Inert 94236.6 96161.7 NA NA Mass_Propellant 110767.3 114316 NA NA delta-V1 3.5 3.6 10.8 4.2 delta-V2 1.4 1.4 7.7 2 delta-V3 1.2 0.9 1.1 3.9 delta-V4 0.9 1.5 0.9 4.7 Crewed_days 1000 1031 NA NA MOE_Score 68.67% 56.49% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600)Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S(1000-1200) delta-V1 6.2 6.4 22.8 7.9 delta-V2 5.5 5.6 20.6 5.8 delta-V3 7.3 8.3 9.8 7.7 delta-V4 22.7 25.6 6.7 23 MOE_Score 0 0 0 0 Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000) S(1000-1200)F(800-1000)S1000-1200) delta-V1 9.2 8.9 11.1 9.1 delta-V2 9.3 7 9.1 7 delta-V3 10 7.5 9.6 7.3 delta-V4 6.2 5.4 8.9 6.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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169 Table B-9. Hybrid I on Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 342639.3 147146.7 NA 471985.9 Mass_Inert 161842 78653.5 NA 190950.1 Mass_Propellant 156310.6 55095.9 NA 251892.3 delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 13.5 7.7 15.7 9.4 delta-V3 (NTR) 0.2 0.2 0.2 0.2 delta-V4 (NEP) 12.8 6.5 44.7 28.9 Crewed_days 951 1082 NA 636 Burn_Days 363.8 593.7 NA 595.8 Power_Elec 4664.4 719.8 NA 6018.1 No. Thrusters 128.5 19.8 NA 165.8 MOE_Score 60.06% 54.81% 0.00% 40.67% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 23.3 20.9 23.4 22 delta-V3 (NTR) 4.5 4.5 4.5 4.5 delta-V4 (NEP) 22 20.3 22.5 22.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 58.7 57 56.2 44.4 delta-V3 (NTR) 2.3 2.3 2.3 2.3 delta-V4 (NEP) 51.8 52.4 56.1 41.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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170 Table B-10. Hybrid MPD Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 514107.3 171516.1 NA NA Mass_Inert 229630.7 89795.6 NA NA Mass_Propellant 249265.5 67807.8 NA NA delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 13.5 7.7 15.7 9.4 delta-V3 (NTR) 0.2 0.2 0.2 0.2 delta-V4 (NEP) 12.8 6.5 44.7 28.9 Crewed_days 951 1082 NA NA Burn_Days 363.9 593.7 NA NA Power_Elec 7618.8 937.4 NA NA No. Thrusters 1.8 0.2 NA NA MOE_Score 52.69% 54.79% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 23.3 20.9 23.4 22 delta-V3 (NTR) 4.5 4.5 4.5 4.5 delta-V4 (NEP) 22 20.3 22.5 22.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 58.7 57 56.2 44.4 delta-V3 (NTR) 2.3 2.3 2.3 2.3 delta-V4 (NEP) 51.8 52.4 56.1 41.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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171 Table B-11. Hybrid Hall Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 624351.9 173573.2 NA NA Mass_Inert 256258.6 85559.4 NA NA Mass_Propellant 333371.2 74474.5 NA NA delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 13.4 7.7 15.7 9.4 delta-V3 (NTR) 0.2 0.2 0.2 0.2 delta-V4 (NEP) 12.7 6.5 44.7 28.9 Crewed_days 951 1082 NA NA Burn_Days 363.9 593.7 NA NA Power_Elec 7495.5 780.2 NA NA No. Thrusters 533.5 55.5 NA NA MOE_Score 44.78% 52.29% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 23.3 20.9 23.4 22 delta-V3 (NTR) 4.5 4.5 4.5 4.5 delta-V4 (NEP) 22 20.3 22.5 22.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 (NTR) 3.2 3.2 3.2 3.2 delta-V2 (NEP) 58.7 57 56.2 44.4 delta-V3 (NTR) 2.3 2.3 2.3 2.3 delta-V4 (NEP) 51.8 52.4 56.1 41.9 MOE_Score 0.00% 0.00% 0.00% 0.00%

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172 Table B-12. NEP Ion Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO 379199.6 161640.8 NA NA Mass_Inert 165437.9 84860.5 NA NA Mass_Propellant 189758.9 67888.2 NA NA delta-V1 21.2 16.6 21 18.1 delta-V2 20.7 16 63.2 40.6 Crewed_days 896 1056 NA NA Burn_Days 701.2 1418.1 NA NA Power_Elec 6076.8 1075 NA NA No. Thrusters 167.4 29.6 NA NA MOE_Score 56.78% 52.01% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 35.2 32.4 34.5 32.1 delta-V2 36 32.5 36.9 38.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 51.1 48.7 59.1 52.3 delta-V2 45.3 45.3 61.7 62 MOE_Score 0.00% 0.00% 0.00% 0.00%

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173 Table B-13. NEP MPD Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO NA 286161.4 NA NA Mass_Inert NA 143767.1 NA NA Mass_Propellant NA 131168.7 NA NA delta-V1 21.2 16.6 21 18.1 delta-V2 20.7 16 63.2 40.6 Crewed_days NA 1056 NA NA Burn_Days NA 1419.7 NA NA Power_Elec NA 1994.9 NA NA No. Thrusters NA 0.5 NA NA MOE_Score 0.00% 47.27% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 35.2 32.4 34.5 32.1 delta-V2 36 32.5 36.9 38.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 51.1 48.7 59.1 52.3 delta-V2 45.3 45.3 61.7 62 MOE_Score 0.00% 0.00% 0.00% 0.00%

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174 Table B-14. NEP Hall Tradespace Results Dest. Mars (2030-2035) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(100-200) S(200300) F(100-200) S(200-300) IMLEO NA 327884 NA NA Mass_Inert NA 135316.6 NA NA Mass_Propellant NA 181721.9 NA NA delta-V1 21.2 16.6 21 18.1 delta-V2 20.7 16 63.2 40.6 Crewed_days NA 1056 NA NA Burn_Days NA 1422.9 NA NA Power_Elec NA 1850.3 NA NA No. Thrusters NA 131.7 NA NA MOE_Score 0.00% 42.17% 0.00% 0.00% Dest. Jupiter (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(10001200)F(800-1000)S(1000-1200) delta-V1 35.2 32.4 34.5 32.1 delta-V2 36 32.5 36.9 38.1 MOE_Score 0.00% 0.00% 0.00% 0.00% Dest. Saturn (2040-2045) Stay Conjunc.(400-600) Oppos.(40-60) Transfer F(800-1000)S(1000-1200)F(800-1000)S1000-1200) delta-V1 51.1 48.7 59.1 52.3 delta-V2 45.3 45.3 61.7 62 MOE_Score 0.00% 0.00% 0.00% 0.00%

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175 APPENDIX C MODELCENTER TEMPLA TE AND FILEWRAPPER CHEBYTOP Hybrid Ephemeris Model Template: $input HEAD='Mars Bimodal Hybrid propulsion system 1' SHOTA='Earth', ; departure planet BULSI='Mars', ; arrival planet jdl=2033,4,1, ; Julian date (year, month, day) iprnt=2, ;forces printing of intermediate ChebyTOP results jdate=0, ;Julian departure date bias from epoch of jdl nv1=2, ;Departure velocity bias flag (2=hyp) nv2=2, ; Arrival velocity bias flag (2=hyp) npow=0, ; constant power source (0=nucl) p0=5000,500,8000, ; initial powernot used if nmuop > 0, units kW is=6000, ;elect_prop_specific_impulse bb=.315, ; elect_prop_thruster efficiency dd=0., ; elect_prop_thruster_eff_isp kt=.05, ; electric_prop_tankage fraction ALFA=15,0.0, ; prop_system_specific_mass nb1=2, ;departure date flag (use jdl given) nb2=2, ; arrival date flag (use jdl given) radep=6785., ; elliptical departure apo tend=300.,-0,300.0, ; mission duration relative to jdl vhl=0, 0.0, 0., ;departure excess velocity vhp=0,0.0,0., ;arrival excess velocity nlv=0, ;launch vehicle number m0d=180000., ; initial mass_tot_system depart=t, ; if true depart from Earth orbit jetisd=t, ; high thrust departure model, jettison mass ispd=915, ;ISP of nuclear prop system glossd=1.015, ;high thrust departure model gravity loss kd=.25,0, rpdep=6787., ; departure parking orbit perigee (km) rarr=37190, ; arrival apo retro=t, ; simulate high thrust retro stage at arrival jetis=t, ; simulate jettison of high thrust retro stage cisp=915, kr=.25,15.606, ; retro stage tankage fraction rparr=3640., ;arrival peri flyby=f, rn=0, ;heliocentric revolution count keep=f, ;if 't' save intermediate values for next run ncop=0, ;optimal Isp flag,1 keeps it constant (keep=0) nmuop=1, ;optimal power flag (keep=1) $nctopt=2 ; constant thrust $end CHEBYTOP Hybrid Ephemeris Model FileWrapper: # ChebyTOP Earth Mars File Wrappe98765 #RunCommands { generate inputFile run "chebytop cheby_Hybrid.inp" parse outputFile templateFile: cheby_Hybrid.template initializationFile: cheby_Hybrid.template fileToGenerate: cheby_Hybrid.inp setDelimiters "=,'"

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176 removeMissingVariables: true # name type row field #---------------------------------------------markAsBeginning "$input" #variable: Title string 2 2 description="Trajectory Title/Description" #variable: rp integer 9 2 description="must be 0 when using alta" enumValues="0" #variable: nctopt integer 23 2 description="1=compute constant thrust solution" enumValues="0,1" default="1" #variable: copla string 28 2 description="Coplanar Solution Flag" enumValues="True,False" #variable: DepartureYear integer 5 2 units="year (jdl(1))" #variable: DepartureMonth integer 5 3 units="month (jdl(2))" #variable: DepartureDay integer 5 4 units="day (jdl(3))" #variable: DepDtFlag integer 17 2 description="Departure Date Flag:0=optimal travel angle solution,1=optimal date min J,2=use date and TOF supplied (nb1)" enumValues="1,2" default="2" #variable: ArrDtFlag integer 18 2 description="Arrival Date Flag:0=optimal travel angle solution,1=optimal date min J,2=use date and TOF supplied (nb2)" enumValues="1,2" default="2" #variable: OptPwrFlag integer 40 2 description="1=Optimal Power Flag (nmuop)" enumValues="0,1" default="1" #variable: SaveIntValues string 39 2 description="If true, save intermediate values for next run (keep)" enumValues="True,False" default="False" #10/25/2001 modification: add to delimiters group and remove ' from planet names setGroup "MissionInfo" variable: DepartureYear integer 5 2 units="year (jdl(1))" variable: DepartureMonth integer 5 3 units="month (jdl(2))" variable: DepartureDay integer 5 4 units="day (jdl(3))" variable: DeparturePlanet string 3 2 description="Departure Body (shota)" variable: ArrivalPlanet string 4 2 description="Arrival Body (bulsi)" variable: HelioTripTime_max double 20 2 description="Heliocentric transfer time-max (tend)" units="days" variable: HelioTripTime_min double 20 4 description="Heliocentric transfer time-min (tend)" units="days" variable: Power_Source integer 10 2 enumValues="0,1" description="Nuclear=0,Solar=1 (npow)" variable: Flyby string 37 2 enumValues="True,False" variable: DepartureApo double 19 2 description="Departure planet apo (radep)" units="kilometers" variable: DeparturePeri double 30 2 description="Departure planet perigee(rpdep)" variable: ArrivalApo double 31 2 description="Arrival Apogee (rarr)" units="kilometers" variable: ArrivalPeri double 36 2 description="Arrival Perigee" #variable: DepVelBias integer 8 2 description="Departure velocity bias flag:0=none,1=asym,2=hyp,3=spiral (nv1)" enumValues="0,1,2,3" default="3" #variable: ArrVelBias integer 9 2 description="Arrival velocity bias flag:0=none,1=asym,2=hyp,3=spiral (nv2)" enumValues="0,1,2,3" default="3" variable: DepHypVel double 21 2 description="Departure excess velocity for use when nv1=1 or 2 (vhl)" units="kilometers/second" default="0"

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177 variable: ArrHypVel double 22 2 description="Arrival excess velocity for use when nv2=1 or 2 (vhp)" units="kilometers/second" default="0" #variable: Helio_Rev_Count double 38 2 description="Heliocentric revolution count, default=0" variable: Outbound_v_bias double 8 2 description="Velocity bias, 1=asymp,2=hyper,3=tangen" variable: Inbound_v_bias double 9 2 description="Velocity bias, 1=asymp,2=hyper,3=tangen" setGroup "Prop_systems" variable: InitialMass_tot double 24 2 description="Initial Mass (m0d)" units="kilograms" variable: InitialPower double 11 2 description="Initial Power at 1 AU (p0)" units="kilowatts" variable: Alpha double 16 2 description="power system alpha kw/kg" variable: Isp_elec double 12 2 description="Specific Impulse (is)" units="seconds" #variable: AlphaSlope double 16 2 description="Propulsion System Specific Mass (alpha(1))" units="kilograms/kilowatts" #variable: AlphaIntercept double 16 3 description="Propulsion System Fixed Mass (alpha(2))" units="kilograms" default="0" variable: Efficiency_elec double 13 2 description="Thruster Efficiency (bb)" default="0.6" #variable: EfficiencyIsp double 14 2 description="Thruster Efficiency (Isp) (dd)" units="seconds" default="0" variable: TankageFrac_elec double 15 2 description="Tankage Fraction (kt)" default="0" #variable: StructuralFrac double 31 2 description="Structural Fraction (ks)" #variable: StationKeepingPower double 15 2 description="Station Keeping Power, subtracts from p0 (sap)" units="kilowatts" default="0" variable: Isp_nucl double 27 2 description="Specific Impulse_nuclear_prop (ispd)" units="seconds" #variable: Isp_flag double 41 2 description="Optimal Isp flag (when non-zero, does not work) variable: Depart_nucl string 25 2 description=="If 't', depart from Earth orbit" enumValues="True,False" variable: Retro_tank_frac double 35 2 description="Retro Stage Tankage Fraction" variable: Retro_nucl string 32 2 description="If 't', simulate high thrust retro stage at arrival" enumValues="True,False" variable: Jettison string 33 2 description="If 't', simulate jettison of high thrust retro stage" enumValues="True,False" } RowFieldOutputFile outputFile { fileToParse: cheby_Hybrid.out setDelimiters whitespace removeMissingVariables: true #resize=true # name type row field #---------------------------------------------clearMarks markAsBeginning "Chebytop 3" setGroup "Dates" variable: DepDtJulian double 4 7 ignoreConversionErrors=true description="Julian departure date"

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178 variable: DepMonth string 4 4 description="departure month" variable: DepDay string 4 5 description="departure day" variable: DepYear integer 4 6 ignoreConversionErrors=true description="departure year" variable: ArrDtJulian double 5 7 ignoreConversionErrors=true description="Julian arrival date" variable: ArrMonth string 5 4 description="arrival month" variable: ArrDay string 5 5 description="arrival day" variable: ArrYear integer 5 6 ignoreConversionErrors=true description="arrival year" markAsBeginning "Solution time=" setGroup "NTR_DeltaV" variable: TMI_Burn double 3 8 description="init.delta V from main engine (nuclear)system" markAsBeginning "Spacecraft parameters" setGroup "NEP_1st_Iter" variable: Elec_Delta_V double 6 7 description="total del-V for elec_prop sys" variable: Burn_time1 string 3 16 #variable: Burn_time_highpe string 3 15 description="total thruster on time (solar is approximated at 1AU) (tp)" units="days"ignoreConversionErrors=false markAsEnd "Solution time=" setGroup "NEP_2nd_Iter" variable: Elec_Delta_V double -4 7 #variable: Burn_time_highpe string -7 15 variable: Burn_time2 string -7 16 #markAsBeginning "Spacecraft parameters" #setGroup "Trajectory_values" #array: x double 24:-26 2 description="plotting data" ignoreConversionErrors=true #array: y double 24:-26 3 description="plotting data" ignoreConversionErrors=true #array: rdata double 24:-26 5 description="plotting data" ignoreConversionErrors=true #array: thetadata double 24:-26 6 description="plotting data" ignoreConversionErrors=true # Unused Variables #markAsBeginning "Constant specific impulse solutions" #variable: jv double 3 12 ignoreConversionErrors=false #variable: jc double 7 14 ignoreConversionErrors=false #variable: jc double 7 -3 ignoreConversionErrors=false #variable: initialaccel double 7 -11 description="a0" units="mm/day^2" ignoreConversionErrors=true #variable: initialaccel2 string 7 4 description="used when accel# runs into a0 variable name" #array: timedata double 10 1 description="plotting data" ignoreConversionErrors=true #array: phidata double 10:-25 7 description="plotting data" ignoreConversionErrors=true #array: magadata double 10:-25 8 description="plotting data" ignoreConversionErrors=true #array: pwrratiodata double 10:-25 8 description="plotting data" ignoreConversionErrors=true #array: conedata double 10:-25 10 description="plotting data" ignoreConversionErrors=true #array: clockdata double 10:-25 11 description="plotting data" ignoreConversionErrors=true #markAsEnd "Solution time="

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179 #setGroup "Summary_Parameters" #variable: tot_TOF double 4 6 description="total flight time (tf)" units="days" #array: ThrusterSwitching string -3 1:-1 description="Thruster switching information" #variable: EqiDv double 9 -11 ignoreConversionErrors=false #variable: PowerTxt string 7 1 description="used when power so large it runs into preceding text" #variable: Power double 7 2 description="Optimal Power" units="kiloWatts" ignoreConversionErrors=false #variable: JetPower double 9 2 description="Jet Power = pwr*efficiency" units="kiloWatts" ignoreConversionErrors=false #variable: MOC_burn double 5 11 description="retro delta V from main engine (nuclear)system" #variable: Power_elec double 6 2 description="total power" #variable: PayloadMass double 4 10 description="Payload Mass = FinalMass-PropSys-Tankage-Structure (mn)" units="kilograms" ignoreConversionErrors=false #variable: FinalMass double 4 4 description="Final Mass = Moprop (mf)" units="kilograms" ignoreConversionErrors=false #variable: PropMass double 4 6 description="Propellant Mass (mp)" units="kilograms" ignoreConversionErrors=false #variable: PropSysMass double 4 8 description="Propulsion System Mass = Pwr*alpha (mps)" units="kilograms" ignoreConversionErrors=false #variable: m0_pe double 6 11 description="Initial Mass/Power" ignoreConversionErrors=true }

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180 LIST OF REFERENCES 1Wade, M., 2007, Von Braun Mars Expedition-1952, Encyclopedia Astronautica, http://www.astronautix.com/craft/vonn1952.htm Jan 2007. 2Griffin, B., Thomas, B., Vaughan, D., A Comparison of Transportation Systems for Human Missions to Mars, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 11-14 July 2004, Fort Lauderdale, FL. 3Stafford, T.P., 1991, America at the Thre shold, Report of the Synthesis Group on Americas Space Exploration Initiative, http://history.nasa.gov/staffordrep/main_toc.PDF Jan 2007. 4Space Transfer Concepts and Analysis for Exploration Missions, Boeing Aerospace and Electronics, Fourth Quarterly Review, NASA Contract NAS8-37857, Oct. 17, 1990. 5Borowski, S.K., Corban, R.R., McGuire, M.L., Beke, E.G., Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars, NASA Lewis Research Center. 6Roy, S., 2006, For Fuel Conservation in Space, NASA Engineers Prescribe Aerocapture, Marshall Space Flight Center study, http://www.nasa.gov/vision/univers e/features/aerocapture.html Jan 2007. 7Borowski, S.K., McGuire, M.L., Mason, L. M., Gilland, J., High Power MPD Nuclear Electric Propulsion (NEP) for Artificial Grav ity HOPE Missions to Callisto, NASA/TM--2003212349. 8Accettura, A.G., Bruno, C., Casotto, S., Marzari, F., Mission to Mars Using Integrated Propulsion Concepts: Considerations, Opportunitie s, and Strategies, Acta Astronautica 54 (2004) 471-486. 9Borowski, S.K., Dudzinski, L.A., Bimoda l Nuclear Electric Propulsion: Enabling Advanced Performance with Near-Term Technologies, AIAA 2002-3653, 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Indianapolis, Indiana, July 710, 2002. 10Beaty, James S., Glossary of Optical Communication Terms, DOT/FAA/CT-TN91-9, U.S. Department of Transportati on, Federal Aviation Administrati on, Technical Center, Atlantic City International Airport, NJ 08405, April 1991. 11Belfiore, M., The Five-Billion-Star Ho tel, Popular Science Magazine, General Technology Section, 2005. 12Bates, R.R., Mueller, D.D., White, J.E., "Fundamentals of Astrodynamics," Dover Public ations, INC., New York.

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181 13Conway, B.A., Prussing, J.E., "Orbital Mechanics," Oxford University Press, 1993. 14Sutton, G.P, Biblarz, O., Rocket Propul sion Elements, John Wiley & Sons, Inc., 2001. 15Eagle, C.D., 2007, Orbital Mechanics with Numerit: The Hohmann Orbit Transfer, Science Software, http://www.cdeagle.com/omnum/hohmann.pdf Jan 2007. 16Wertz, J.R., Larson, W.J, Space Mission Analysis and Design, Microcosm Press & Kluwer Academic Publishers, 3rd edition, 2003. 17Striepe, S.A., Braun, R.D., Effects of A Venus Swingby Periapsis Burn During an Earth-Mars Trajectory, The Jour nal of Astronautical Sciences, Vol. 39, No. 3, July-September 1991. 18Borowski, S.K., Dudzineki, L.A., McGuire, M.L, Bimodal Nuclear Thermal Rocket (NTR) Propulsion for Power-Rich Artificial Grav ity Human Exploration Missions to Mars, IAA-01-IAA.13.3.05, NASA Gle nn Research Center, 2001. 19Benton, M.G., Conceptual Design for Interplanetary Spaceship Discovery, American Institute of Physics 0-7354-0305-8, 2006. 20Space Transfer Concepts and Analysis for Exploration Missions, Fourth Quarterly Review, Boeing Aerospace and Electronics, October 17, 1990. 21Exploration Studies Technical Report; Volume 1: Mission and Integrated Systems, The Office of Exploration, FY 1989 Annual Report. 22Sakai, T., 2004, A Study of Variable Thru st, Variable Specific Impulse Trajectories for Solar System Exploration, Georgia Inst itute of Technology, Aerospace Engineering Dissertation, http://etd.gatech.edu/thes es/available /etd-11212004-151326/ Jan 2007. 23Navagh, J., 1993, Optimizing Interplane tary Trajectories with Deep Space Maneuvers, NASA Cont ractor Report 4546, http://ntrs.nasa.gov/archive/nasa /casi.ntrs.nasa.gov Dec. 2006. 24Riehl, J.P., 2007 Tools used by the Analys is & Integration Group: Chebytop, NASA Glenn Research Center: Space Propuls ion & Mission Analysis Office, http://trajectory.grc.nasa .gov/tools/chebytop.shtml Feb 2007. 25Polsgrove, CHEBYTOP Users Manual, NASA internal document, April 2006. 26Humble, R.W., Henry, G.N., Larson, W.J ., Space Propulsion Analysis and Design, McGraw-Hill Companies, 1995. 27Pike, J., 2007, Nuclear Resources: React or Concepts, Federation of American Scientists, http://www.fas.org/nuke/space/c02early.htm Jan 2007.

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182 28Cinnamon, J.D., 2005, Nuclear Thermal Rocket Propulsion: Design Issues and Concepts, Texas Space Grant Consortium, Univ ersity of Texas at Austin, Department of Aerospace Engineering, http://www.tsgc.utexas.e du/archive/fulltext/nuke.pdf Jan 2007. 29Howe, S., Assessment of the Advantages a nd Feasibility of a Nuclear Rocket for a Manned Mars Mission, Los Alamos Nationa l Laboratory LA-UR-85-2442, Los Alamos, New Mexico, 1985. 30Bixler, C.H., Gunther, N.A., Rochow, R.F ., Polansky, G.F., A Bimodal Power and Propulsion System Based on Cermet Fuel and H eat Pipe Energy Transport, Sandia National Laboratories, US. DOE c ontract DE-AC04-96AL85000. 31Nuclear Thermal Rocket (N TR) Engine System Assessm ent: Status Review, NASA Contract No. NA53-25809, Science Applications International Corpor ation, Nov.7, 1990. 32McGuire, M.L., Martini, M.C., Packard, T. W., Weglian, J.E., Gilland, J.H., Use of High-Power Brayton Nuclear Electric Propulsi on (NEP) for a 2033 Mars Round-Trip Mission, NASA/TM-2006-214106. 33Choueiri, E.Y., A Critical History of El ectric Propulsion: The First 50 Years (19061956), Journal of Propulsion and Power, Vol. 20, No. 2, 2004. 34Dunning, J.W., Hamley, J.A., Jankovsky, R.S., Oleson, S.R., An Overview of Electric Propulsion Activities at NASA, Glenn Research Center, Cleveland, Ohio, 2004. 35Jahn, R.G., Choueiri, E.Y., Electric Propulsi on. Encyclopedia of Physical Science and Technology, Third Edition, Volume 5, 2002. 36Peradzynski, Z., 2006, Electric Propulsion Research, Institute of Fundamental Technological Research, Polish Academy of Science, http://fluid.ippt.gov. pl/sbarral/index.html Jan 2007. 37Charania, A., St. Germain, B., Wallace, J.G., Olds, J.R., REACIONN: A Nuclear Electric Propulsion Mission Concept to the Outer Solar System, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Fort Lauderdale, FL, July 11-14, 2004. 38Jacobson, D.T., Manzella, D.H., Hofer, R.R., Peterson, P.Y., NASAs 2004 Hall Thruster Program, AIAA conference paper. 39MMWe NEP Charts, NASA Glenn Res earch Center Internal Document. 40Pinero, L.R., 2005, Multikilowatt Power Module Designed and Fabricated for HighPower Hall Thrusters, NASA Gl enn Research Center study, http://www.grc.nasa.gov/W WW/RT/2003/5000/5430pinero.html Feb. 2007.

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183 41Borowski, S.K., Bimodal Nuclear Therma l Rocket (BNTR) Propulsion for Future Human Mars Exploration Missions, presentation at the 2003 NASA Seal/Secondary Air System Workshop, November 5-6, 2003. 42Lindroos, M., 2007, TransHab Modul e, Encyclopedia Astronautica, http://www.astronautix.com/craft/traodule.htm Jan 2007. 43Guidelines and Capabilities for Designing Human Missions, Chapter 5: Radiation, NASA Technical Report, PB2003-101438. 44Wilson, J.W., et.al., Galactic and Solar Cosmic Ray Shielding in Deep Space, NASA Technical Paper 3682, Dec., 1997. 45G.L. Kulcinski, NEEP 423 Course Notes, Lecture 2, University of Wisconsin, 1996. 46Hodus, S., 2006, ModelCenter 7.0: Design Y our Success, Phoenix Integration, Inc., http://www.phoenix-int.com Jan 2007. 47Joyner, R., Lentati, A., Cichon, J., Missi on Analysis Trades Using Nuclear Propulsion Design Options Apllied to a Human Mars Mission , American Institute of Aeronautics and Astronautics, working paper, 2006. 48Sproles, N. Establishing Measures of Effec tiveness for Command and Control: A Systems Engineering Perspectivethe Journa l of The Internatio nal Council on Systems Engineering, Vol. 4, No.2, 2001.

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184 BIOGRAPHICAL SKETCH Jaclyn Cichon grew up in Tampa, FL and grad uated from the Academy of the Holy Names in May 2001. In fall 2001, her passion for aeros pace engineering led her to the Massachusetts Institute of Technology. Upon graduating with her Aeronautical/Astro nautical Engineering degree in June 2005, she moved back to Florida to begin graduate studies at the University of Florida. Here she entered the Nuclear Engineer ing Department, which en abled her to pursue her interests in propulsion through st udy in the area of nuclear space propulsion. She worked under Dr. Samim Anghaie with the Innovative Nucl ear Space Power and Propulsion Institute and focused her thesis work on missi on analysis studies for interpla netary crewed missions. Upon graduating with her Master of Science degree in nuclear engineering in May 2007, she will begin work as a Missile Defense Agency contractor with the Boeing Company in Washington DC.


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Permanent Link: http://ufdc.ufl.edu/UFE0020625/00001

Material Information

Title: Trade study of nuclear space power and propulsion system architectures for advanced interplanetary travel
Physical Description: Mixed Material
Language: English
Creator: Cichon, Jaclyn ( Dissertant )
Clarke, John-Paul ( Thesis advisor )
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2007
Copyright Date: 2007

Subjects

Subjects / Keywords: Nuclear Engineering Sciences thesis, M.S
Dissertations, Academic -- UF -- Nuclear and Radiological Engineering
Genre: bibliography   ( marcgt )
theses   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract: Past mission analysis studies for interplanetary missions have typically focused on assessing one power and propulsion system applied to a singular mission. This has led to a lack of comparable results for different power and propulsion systems, and limited systematic procedures for mission planning. We attempted to address these faults through a comprehensive trade study on NTR, NEP, and Hybrid power and propulsion systems, grouped by NTR fuel type or electric thruster type into eleven configurations. Ephemeris codes were used to investigate mission planning and departure options, and vehicle models were developed and then analyzed through parametric performance plots. Lastly, measures of effectiveness were defined and used to assess a tradespace matrix containing all vehicle configurations applied to twelve optimized reference missions. This assessment provided a simplified ranking of specific vehicle designs for each given interplanetary mission, along with a comparison of vehicle performance based on important figures of merit such as propellant mass and IMLEO.
Subject: electric, NEP, NTR, nuclear, propulsion, space
General Note: Title from title page of source document.
General Note: Document formatted into pages; contains 184 pages.
General Note: Includes vita.
Thesis: Thesis (M.S.)--University of Florida, 2007.
Bibliography: Includes bibliographical references.
General Note: Text (Electronic thesis) in PDF format.

Record Information

Source Institution: University of Florida
Holding Location: University of Florida
Rights Management: All rights reserved by the source institution and holding location.
System ID: UFE0020625:00001

Permanent Link: http://ufdc.ufl.edu/UFE0020625/00001

Material Information

Title: Trade study of nuclear space power and propulsion system architectures for advanced interplanetary travel
Physical Description: Mixed Material
Language: English
Creator: Cichon, Jaclyn ( Dissertant )
Clarke, John-Paul ( Thesis advisor )
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2007
Copyright Date: 2007

Subjects

Subjects / Keywords: Nuclear Engineering Sciences thesis, M.S
Dissertations, Academic -- UF -- Nuclear and Radiological Engineering
Genre: bibliography   ( marcgt )
theses   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract: Past mission analysis studies for interplanetary missions have typically focused on assessing one power and propulsion system applied to a singular mission. This has led to a lack of comparable results for different power and propulsion systems, and limited systematic procedures for mission planning. We attempted to address these faults through a comprehensive trade study on NTR, NEP, and Hybrid power and propulsion systems, grouped by NTR fuel type or electric thruster type into eleven configurations. Ephemeris codes were used to investigate mission planning and departure options, and vehicle models were developed and then analyzed through parametric performance plots. Lastly, measures of effectiveness were defined and used to assess a tradespace matrix containing all vehicle configurations applied to twelve optimized reference missions. This assessment provided a simplified ranking of specific vehicle designs for each given interplanetary mission, along with a comparison of vehicle performance based on important figures of merit such as propellant mass and IMLEO.
Subject: electric, NEP, NTR, nuclear, propulsion, space
General Note: Title from title page of source document.
General Note: Document formatted into pages; contains 184 pages.
General Note: Includes vita.
Thesis: Thesis (M.S.)--University of Florida, 2007.
Bibliography: Includes bibliographical references.
General Note: Text (Electronic thesis) in PDF format.

Record Information

Source Institution: University of Florida
Holding Location: University of Florida
Rights Management: All rights reserved by the source institution and holding location.
System ID: UFE0020625:00001


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TRADE STUDY OF NUCLEAR SPACE POWER AND PROPULSION SYSTEM
ARCHITECTURES FOR ADVANCED INTERPLANETARY TRAVEL



















By

JACLYN CICHON


A THESIS PRESENTED TO THE GRADUATE SCHOOL
OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT
OF THE REQUIREMENTS FOR THE DEGREE OF
MASTER OF SCIENCE

UNIVERSITY OF FLORIDA

2007


































2007 Jaclyn Cichon

































I would like to dedicate this work to my parents.









ACKNOWLEDGMENTS

The material presented in this thesis is the culmination of both hard work and dedicated

studies in the fields of aerospace and nuclear engineering. Presenting this thesis in its final form

has only been done through the support and assistance of many people.

I would first like to thank my parents for their support throughout my life, and most

notably their encouragement of my intellectual pursuits. I would like to thank Dr. John-Paul

Clarke, my UROP advisor at MIT who assured me that I would get through college and someday

be a successful graduate student. I would like to thank Dr. Samim Anghaie for convincing me to

further my studies in graduate school and allowing me the opportunity to pursue a Nuclear

Engineering master's degree. I owe much gratitude to my Pratt & Whitney mentor, Russ Joyner,

who gave me the time and the tools I needed to begin and complete this thesis. Lastly, I would

like to thank all of my friends and peers both at MIT and UF for their constant words of

encouragement, especially those who truly believed I could someday become the "President of

NASA."









TABLE OF CONTENTS

page

A CK N O W LED G M EN T S ........................................................ ....................................... 4

L IST O F T A B L E S ...................................... ..................................... ................ .. 8

LIST OF FIGURES .................................. .. .... ..... ................. 10

A B S T R A C T ......... ....................... ............................................................ 13

CHAPTER

1 INTRODUCTION ............... .......................................................... 14

M motivation .......... ............................... ................................................14
P roje ct S tatem en t ...............................................................................................1 5
Previous W ork and Contributions ................................................ .............................. 16
S tu dy O v erv iew .................................................................................................. ........... 19
Study M methodology .................................................... ........ .. ........ .... 19

2 M IS SIO N P L A N N IN G .............................................................................. ......................23

A strody n am ics ................................................................2 3
O rbital M option .............................................................................................................23
Patched Conic M ethod .................................. .. .. .. ...... .. ............25
M mission P profiles ....................................... .......................................................... ............... 3 0
Considerations for M mission Planning ........................................ ......................... 31
M mission trajectory ................................................................................ .. 3 1
Mission approach .................................. .............. ............... 32
M mission P rofi le Selection ........................................................................ .................. 34
M mission trajectory ............... ................. ............. ........... ........... .. 35
M mission approach ............................... .. ..... ......... ............. ..... 38
M mission profit le sum m ary .............................................................. .....................39
E p h em e ris T o o ls ........................................................................................................4 0

3 M O D EL D E V E L O PM E N T .......................................................................... ....................44

Space R actor B background ............................................................................. ............... 44
R actor C onfi guration .......................................................................... .....................44
R actor F mission Spectrum .......................................................................... ............. 45
Nuclear Thermal Power and Propulsion......................... ....... .......................... 46
History of Nuclear Thermal Rocket (NTR) Systems ............................................... 46
F undam ental R research ............ ...... ........................................................ ......... ....... 47
D evelopm ent of the N TR M odel ................... ......... ............................ ............... 50
P o w e r ................................ ......................................................5 2
P ro p u lsio n .......................................................................... 5 5









Nuclear Electric Pow er & Propulsion ................................................... .... ...........57
History of Electric Propulsion (EP) System s ...................................... ............... 57
Fundam mental Research and EP D description ........................................ .....................59
Ion thrusters ................................................................................................... ....... 60
M P D th ru sters ................................................................6 1
H all thrusters ............... .......... ..... .......... .........................61
Development of the Nuclear Electric Propulsion (NEP) Model .............................. 63
P o w e r ..........................................................................................6 4
P ro p u lsio n ........................................................................................................... 6 6
H ybrid P ow er and P ropulsion ........................................................................ ...................72
B background ................... .........................................................................72
D evelopm ent of the Hybrid M odel........................................................ ................73
P o w e r ..........................................................................................7 3
P propulsion ......................................................................................................74
C om m on V vehicle Subsystem s ........................................................................ ..................75
T ra n sH ab ...............................................................................7 5
Spacecraft Functional Components............................ ........ .................. 79
A n aly sis T o o ls ............................................................................... 8 0
M odelC enter Softw are............. .................. .............. .......................... .............. 80
Incorporation of Models into ModelCenter Framework ...............................................81
E p h em eris m o d els ................ .... .............. ...................... ................. .. .. .. ..8 1
Vehicle architecture models ................................. ........................ ....83
M mission architecture m odels........................................................... ............... 84
V alidation of m models .............. ........ ......... ... .................. .......... .... 85

4 TR A D E STU D Y R E SU L T S ......................................................................... ...................96

Ephemeris Analysis .............. ..... ................ .............. 96
N T R ................................ ............................................................................ 9 7
D ep artu re ................................ ................................................................... 9 7
D u ra tio n .............................................................................................................. 9 8
NEP ............. ......... ........... ................ ...............99
D ep artu re .................................................................................................... 9 9
D u ratio n ................................ .......................................... 10 0
H y b rid ....................... .................... .. .............................................................................1 0 1
Departure ........................ ....................... 101
D u ratio n ........................................................................................1 0 2
Vehicle Architecture Analysis.............................................. 103
N T R ................... ...................1...................0.........4
NEP ............. ......... ............................................106
Hybrid ............... ................................. ...............108
Tradespace Assessment ........................................................................... ... ......... ................... 109
M methodology ............................................................................................. ........ 110
M mission architecture analysis ............. ................................ ............... 110
Determining M OE scores ................................................................................. 112
Tradespace Mission Architecture Results ............. ......... ........................115
Tradespace M O E R results ........................ .. .............. ..............................117


6









5 C O N CLU SIO N ................... .................................................................. .... 154

Study Contributions ....................................................... .. .......... ......... ... .. 154
Areas of Future W ork ............... ................. ........... ................. ......... 155
F in al C om m ents.............................. ............................................................. ............... 156

APPENDIX

A L IST O F A B B R E V IA TIO N S.................................................................... ..................... 158

B TRADESPACE MISSION ARCHITECTURE RESULTS .............................................161

C MODELCENTER TEMPLATE AND FILEWRAPPER....... .........................................175

L IST O F R E F E R E N C E S ......... .. ............... ................ .......................................................... 180

B IO G R A PH IC A L SK ETCH ......................................................................................... 184





































7









LIST OF TABLES

Table page

2-1 Design Reference Mission Categories .................... ................................ 43

3-1 N T R A architecture Sy stem s ........................................................................ .................. 88

3-2 F uel O operating P aram eters........................................................................ ...................88

3-3 N T R R actor M ass Sizing ........................................................................ ...................88

3-4 NEP Pow er Reactor D descriptions ............................. ....... .............................88

3-5 EP Thruster Characteristic Param eters ........................................ ......................... 88

3-6 TransH ab M ass B reakdow n ............................................... ........................... ...............88

3-7 Common Subcomponent M ass Breakdown................................... ........................ 89

3-8 N TR M odel Validation M ass Estim ates ........................................ ........................ 89

4-1 M OE4 TRL Score ........ ..... .. .... .......................... ... .... 120

4-2 Tradespace M OE Scores..................................................................... ............... 120

4-3 Tradespace M ission-Based Rankings ....... .. ........................... ..... ................... 121

4-4 Tradespace Planet-Based and Overall Rankings .................................. ...............121

B -l N TR O ptim ized Ephem eris R esults..................................................................... ...... 161

B-2 NEP Optimized Ephemeris Results ............. ............ ............... 162

B-3 Hybrid Optim ized Ephem eris Results ........................................ ........................ 163

B -4 N TR G raphite Tradespace R esults....................................................................... ...... 164

B-5 N TR Com posite Tradespace Results .......................................................... .... ........... 165

B -6 N TR C arbide Tradespace R esults........................................................................ ... ... 166

B-7 NTR CERMET-Unimodal Tradespace Results........................................................167

B-8 NTR CERMET-Bimodal Tradespace Results ........................ ...............................168

B-9 Hybrid Ion Tradespace Results.............. .............................................. ............... 169

B-10 Hybrid M PD Tradespace Results........... .......................................... ............... 170



8









B -11 H ybrid H all Tradespace R results ............................................. ............................. 171

B -12 N EP Ion Tradespace R esults......... ................. ................... .................. ............... 172

B-13 N EP M PD Tradespace Results ............................................... ............................. 173

B -14 N EP H all Tradespace R results ......... .................... ......... .................................174









LIST OF FIGURES

Figure page

1-1 Tradespace M atrix ..................................... ................. ...... ....... 22

2-1 Escape Hyperbola ............................... ... ...... ... .............. ... 43

3-1 NTR Unimodal Vehicle Architecture Schematic ................................... .................90

3-2 NTR Bimodal Vehicle Architecture Schematic...................................... ............... 90

3-3 NEP Vehicle Architecture Schematic..... .. .............................................. ............... 91

3-4 N EP R actor M ass vs. Pow er .................................................. .............................. 91

3-5 Pow er and Efficiency Flow D iagram ........................................... .......................... 92

3-6 PM A D M ass Sizing E quation ......................................... .............................................92

3-7 Hybrid Vehicle Architecture Schematic......... ............................................... 93

3-8 N E P E phem eris T ool G U I ........................................................................ ...................93

3-9 D arw in O ptim ization Tool G U I........................................................................... .... ... 94

3-10 M odelC enter Screenshot.......................................................................... ....................94

3-11 M odelCenter M mission Architecture GUI ........................................ ........ ............... 95

4-1 M odelCenter Param etric Tools............................................... ............................. 122

4-2 N TR D elta-V vs. D departure Year ....................................................... ............... 122

4-3 NTR Mars Delta-V vs. Departure Year and Month................................... ...............123

4-4 N TR D elta-V vs. Transfer Tim e ......................................................................... ...... 124

4-5 N TR D elta-V vs. Stay Tim e.......................................................... ............................. 124

4-6 NTR Mars Delta-V vs. Transfer and Stay Time ......... ................... .... ............125

4-7 NTR Jupiter Delta-V vs. Transfer and Stay Time ................................. ...............126

4-8 NTR Saturn Delta-V vs. Transfer and Stay Time.........................................................127

4-9 N EP D elta-V vs. D departure Y ear............................................. ............................ 128

4-10 NEP Mars Delta-V vs. Departure Year and Month............... .................................... 129



10









4-11 NEP Jupiter Delta-V vs. Departure Year and Month ............................................... 130

4-12 NEP Saturn Delta-V vs. Departure Year and Month.............. ..... .................131

4-13 N E P D elta-V vs. Transfer Tim e............................................................ .....................132

4-14 N EP D elta-V vs. Stay Tim e ........................... ...................................... ............... 132

4-15 NEP Mars Delta-V vs. Transfer and Stay Time ................................... ............... 133

4-16 NEP Jupiter Delta-V vs. Transfer and Stay Time......... ......... .................... 134

4-17 NEP Saturn Delta-V vs. Transfer and Stay Time ....................................... ...............135

4-18 Hybrid Delta-V vs. Departure Year ....................................... 136

4-19 Hybrid Mars Delta-V vs. Departure Year and Month ............................................. 137

4-20 Hybrid Jupiter Delta-V vs. Departure Year and Month................................................138

4-21 Hybrid Saturn Delta-V vs. Departure Year and Month.............................139

4-22 H ybrid D elta-V vs. Transfer Tim e....................................................................... ...... 140

4-23 H ybrid D elta-V vs. Stay Tim e ................................................................................... 140

4-24 Hybrid M ars Delta-V vs. Transfer and Stay Time.................................... ............... 141

4-25 Hybrid Jupiter Delta-V vs. Transfer and Stay Time....................................................... 142

4-26 Hybrid Saturn Delta-V vs. Transfer and Stay Time ................................ ...............143

4-27 N TR IM L E O vs. D elta-V ........................................................................ .................. 144

4-28 NTR M ass Breakdown vs. Fuel Type ...................................................... ............. 144

4-29 N TR Bum Tim e vs. Fuel Type .................................................................................. 145

4-30 NTR Vehicle M ass Breakdown ....................................................................... 145

4-3 1 N E P P ow er v s. D elta-V ........................................ ........... ............. .............................146

4-32 N E P IM LE O vs. D elta-V ............................ ............................................................. 146

4-33 NEP Mass Breakdown vs. Thruster Type................................. ...............147

4-34 NEP IMLEO vs. Bum Time and Delta-V.............. ................. ....... ............... 148

4-35 NEP V vehicle M ass Breakdow n............. ............................................... ............... 149









4-36 H ybrid IM LE O vs. D elta-V ............................................................................. ............149

4-37 Hybrid Mass Breakdown vs. Thruster Type............ .............. .................... 150

4-38 Hybrid IMLEO vs. NEP and NTR Delta-V............................. .................151

4-39 H ybrid V vehicle M ass B reakdow n ........................................................................ ... ... 152

4-40 M ars M mission IM LE O s ........................................................................................... 152

4-41 Tradespace MOE Scores for Mars Missions ....................................... ............... 153









Abstract of Dissertation Presented to the Graduate School
of the University of Florida in Partial Fulfillment of the
Requirements for the Degree of Master of Science

TRADE STUDY OF NUCLEAR SPACE POWER AND PROPULSION SYSTEM
ARCHITECTURES FOR ADVANCED INTERPLANETARY TRAVEL

By

Jaclyn Cichon

May 2007

Chair: Samim Anghaie
Major: Nuclear Engineering Sciences

Past mission analysis studies for interplanetary missions have typically focused on

assessing one power and propulsion system applied to a singular mission. This has led to a lack

of comparable results for different power and propulsion systems, and limited systematic

procedures for mission planning. We attempted to address these faults through a comprehensive

trade study on NTR, NEP, and Hybrid power and propulsion systems, grouped by NTR fuel type

or electric thruster type into eleven configurations. Ephemeris codes were used to investigate

mission planning and departure options, and vehicle models were developed and then analyzed

through parametric performance plots. Lastly, measures of effectiveness were defined and used

to assess a tradespace matrix containing all vehicle configurations applied to twelve optimized

reference missions. This assessment provided a simplified ranking of specific vehicle designs

for each given interplanetary mission, along with a comparison of vehicle performance based on

important figures of merit such as propellant mass and IMLEO.









CHAPTER 1
INTRODUCTION

Motivation

The desire to travel though space and reach distant planets has no doubt existed for ages,

yet the actual planning of interplanetary missions truly began in the United States approximately

60 years ago. This is when Werner Von Braun, former aerospace engineer and National

Aeronautics and Space Administration (NASA) administrator, conducted the first engineering

analysis of a manned mission to Mars. By the late 1960s Von Braun had become a proponent for

nuclear thermal rocket-powered spacecraft, while his Soviet contemporary Korolev believed

nuclear electric propulsion was the best option to go to Mars. Cancellation of the Apollo

program in the 1970s led to a sharp decline in nuclear propulsion and Mars exploration research.1

In 1988 President George H. Bush appointed NASA to create the Mars Office of Exploration as

part of the Space Exploration Initiative, resulting in a resurgency in reports on manned missions

to the moon and Mars.2 It was believed that the moon would be a stepping stone to Mars, as one

report states: "the Moon provides a unique database for life science and operational verification

in a reduced gravity environment, combined with the psychological realism of operations at a

harsh extraterrestrial location."3

Although the goals sought by the Space Exploration Initiative never came to fruition, a

renewed interest in moon and Mars exploration was expressed by George W. Bush under the

'Moon Mars and Beyond' initiative in 2002. NASA has indeed answered this call for a return to

the moon and then to Mars, doing so at a pace commensurate with its appropriated funding and

resources. The government agency, with the help of its industry partners, is currently in the

initial phases of its Constellation project, which hopes to return man to the moon by 2014 and

eventually replace the outdated and complex Space Shuttle with the Orion Crew Exploration









Vehicle. Despite all the preparation a return to the moon would provide for a future Mars

mission, one of the main differences would be that a vastly different propulsion system would be

needed to make the much longer journey to Mars. In fact it has been noted by mission planners

that "the choice of a transportation system is the key trade in performance for Mars

exploration."3 This is why it is paramount to invest now in Mars mission planning studies that

assess the various power and propulsion architectures that have the capabilities that will be

required to carry man to the surface of Mars and back.

Project Statement

Our study aimed to provide this necessary mission analysis and to do so from as broad a

scope as possible. Past studies have provided results for specific missions using only one

propulsion system, making it difficult to try and compare all possible propulsion systems for the

same mission. Our study attempted to provide a comparison of the major power and propulsion

system architectures considered for future manned interplanetary missions and to compare them

on an "even playing field." This involved elements of mission analysis, system modeling and

optimization, and parametric analysis that culminated in a large trade study. Although the end-

goal was to provide an answer to which power and propulsion system model best completes the

attempted interplanetary missions using the aide of a tradespace matrix, analysis of non-mission-

specific vehicle performance and interplanetary trajectories was also assessed. It was the hope of

this author that the approach taken was broad enough in context to allow application to a myriad

of future space mission profiles, yet specific enough to provide valid and useful results of vehicle

performance. This mission analysis study promises a unique perspective on the subject matter

given the combined systems engineering perspective and fundamentals of nuclear reactor design,

astrodynamics, and rocket propulsion.









Through investigation of previously completed Mars mission studies, a commonality

found was that the primary objective of each study was to safely send and return humans to

Mars.2 Although this objective was inherent as well to this study, a very important secondary

objective was to provide a comparison of different vehicle architectures on an 'apples-to-apples'

level, such that bias towards or against a particular system was completely withheld. Each of the

architecture models was pieced together following the same methodology, with the same design

constraints and without regard to particular mission profiles. Thus, the optimization of vehicle

performance for each mission profile produced results that could provide an honest comparison

of the different propulsion systems currently under consideration by both NASA and other

members of the aerospace community for future manned interplanetary missions. An extra

feature of this study was that even more challenging missions to Saturn and Jupiter were

attempted so that a more in-depth understanding of the capabilities of the technologies being

assessed could be gleaned.

Previous Work and Contributions

Many of the first formal studies comparing advanced propulsion methods for

interplanetary missions came about during the Space Exploration Initiative period, which

focused on missions to the moon and then to Mars. At that point in time, it was desired to

compare the conventional chemical propulsion system with that of a nuclear thermal or electric

propulsion system. One report that was done by Boeing for presentation to NASA was the

"Space Transfer Concepts and Analysis for Exploration Missions" report. This comprehensive

report attempted to compare the possible propulsion methods through various analyses,

diagrams, and charts. Initial Mass in Low Earth Orbit (IMLEO) charts were presented in many

different fashions, but one of the most useful was a chart showing the optimal IMLEO versus trip

time for all the opposition opportunities in the Earth-Mars synodical cycle.









This chart showed these opportunity bands, with the lower boundary of each propulsion

system option formed by the best year (2018), and the upper boundary by the most difficult year

(2025), given all opportunities between the years 2010 and 2025. The nuclear electric propulsion

(NEP) system was the front-runner for the shorter mission trips, yet it did not seem realistic to

look at total mission trips to Mars below 400 days, especially given the radically high power

requirement (120 MW) and IMLEO (1000 t) that would be needed for a fast NEP mission to

Mars. Looking at the more reasonable time span of 400-600 day missions, the advanced NTR

system was the front-runner, with the NEP system close behind, and lastly the rather impractical

chemical system. It should be noted, however, that if longer mission trip times were acceptable,

the NEP system would present the lowest IMLEO option given all the systems assessed.4

In more recent studies in line with the 2002 'Moon, Mars, and Beyond' initiative, the

mission analysis performed has commonly focused on one specific power and propulsion

system, going into a bit more detail on mission specifics, hardware, and mass and performance

estimates. Three such studies were reviewed that specifically focused on one main architecture

that is currently considered feasible for advanced interplanetary travel. These architectures

include a Nuclear Thermal Rocket (NTR) system, NEP system, and a Hybrid system that uses

one reactor to power both electric and thermal propulsion systems.

A study entitled "Nuclear Thermal Rocket Vehicle Design Options for Future NASA

Missions to the Moon and Mars" that was conducted by the NASA Lewis Research Center

analyzed a 2010 human mission to Mars. This mission consisted of a 344-day outbound transfer,

1153-day stay at Mars, and a 180-day return. Three NTR burns were performed with an Earth

swingby used for the last maneuver in order to reduce energy requirements. Two different

reactors were considered, one based on a NERVA reactor, and the second utilizing ternary









carbide fuels. The first had a specific impulse (ISP) of 900 (s), and the latter an ISP of 960 (s).

The reactors provided thermal power for the propulsion system, and electric power for any

spacecraft needs. This bimodal reactor design functioned by flowing hydrogen coolant through

the core to heat the propellant, and then a helium-xenon working fluid to remove heat for power

conversion purposes. The results for the NTR spacecraft included an IMLEO of 234 (t) for the

NERVA-type reactor, and 207 (t) for the carbide reactor, given four reactors each operating at a

thrust level of 15 (klbf).5

A team of engineers at NASA's Glenn Research Center conducted the NEP mission

analysis study entitled, "High Power MPD Nuclear Electric Propulsion (NEP) for Artificial

Gravity HOPE Missions to Callisto." The designed mission was a 4.5-year round-trip crewed

journey to Jupiter's moon Callisto, beginning in the year 2041. The vehicle was to be proceeded

by both a cargo and tanker vehicle, support a crew of six, provide artificial gravity for most of

the mission duration, and be refueled before leaving Callisto to return to Earth. The spacecraft

was to be powered by a high temperature gas-cooled fission reactor with a tungsten metal matrix

CERMET fuel element. Electrical power was generated using a high power, closed-cycle

Brayton heat engine. Hydrogen MPD thrusters provided propulsion with advanced performance

parameters including 8000 (s) ISP and 64.5% efficiency. A full mission analysis of this system

revealed that the 120-day stay, 2.1-year transfer to and from Callisto would require an IMLEO of

262 (t). This value accounts for an outbound propellant mass of 74 (t), but does not include the

return propellant load of 53 (t), which was carried to Jupiter aboard the tanker vehicle.6'7

The results of a mission analysis study performed to evaluate the feasibility for a mission

to Mars using a Hybrid propulsion system were found in "Mission to Mars Using Integrated

Propulsion Concepts: Considerations, Opportunities, and Strategies." Integrated Propulsion









Systems (IPS) is the term used to refer to a Hybrid system, one that uses both NTR and NEP

systems for propulsive means. The mission scenario in this study was a manned mission to Mars

in which the spacecraft was assembled in Low Earth Orbit (LEO) at about a 300 (km) altitude,

and then the reactor engine was turned on to begin the mission. In this particular study, the

nuclear and electric engines were both used to spiral out of the Earth's atmosphere until the

required Earth escape velocity was reached. The spacecraft then separated into two by means of

a tether for on-orbit artificial gravity, and restarted both propulsion systems for simultaneous

operation during the remaining voyage to Mars.8

The technology assumed in this study consisted of a "Rubbia Nuclear Rocket" and MPD

electric thrusters. In the Rubbia nuclear reactor, the heat exchange is essentially reversed from a

typical NERVA-type reactor, with fission fragments from subcritical fissions of an isotope of

Americium heating the coolant. This allows lower fuel operating temperatures, while enabling

higher ISP values to be reached. Four superconductive Magneto-Plasma-Dynamic (MPD)

thrusters were assumed for the electric propulsion system. It was assumed in the study that the

ISP of the Rubbia systems was 3500 (s) and that of the MPD system was 56000 (s). The mass

breakdown for this IPS spacecraft was found to be 378 (t) with an outbound propellant mass of

132 (t) and return propellant load of 92 (t).8'9

Study Overview

Study Methodology

Our study attempted to go farther than previous studies, which primarily focused on

specific power and propulsion systems applied to a singular mission, by assessing three types of

vehicles broken down into eleven different configurations over each of twelve possible mission

scenarios. Basic methodology used to perform such an expansive mission trade study includes

four major steps:









1. Defining the general mission concept and propagating this to twelve different reference

mission profiles

2. Creating ephemeris models

a. Generating optimized trajectories for each mission profile

b. Obtaining results of parametric study of ephemerides

3. Developing models for the space power and propulsion architectures

a. Running the vehicle models with mission trajectory requirements

b. Performing parametric study of vehicle design attributes

4. Comparing all vehicle configurations in tradespace matrix

a. Defining measures of effectiveness (MOEs) and scoring algorithm

b. Assessing all mission architecture blocks in tradespace matrix

The reference missions, or mission profiles, used in this study were built upon missions

studied in the past. Though it was desired to encompass a broad spectrum of possible planetary

destinations, stay times and transfer durations, this spectrum was derived from missions in

previously successful studies. Literature sources were also very useful in confidently making

assumptions and ground rules as to trajectory choices, vehicle staging, and planetary escape and

capture mechanisms. The choice of mission profiles, however, is inevitably up to the mission

planner, with no real wrong or right manner in which to make the choices. Major considerations

in designing the profiles were thus focused on designing missions that would be unbiased

towards a particular vehicle, and to obtain a set of missions that would adequately represent a

subset of realistic mission scenarios for the tradespace. It was essentially the design of the

tradespace matrix and designation of appropriate MOEs that will, in the end, determine the utility

of the final results.









The next step in the trade study methodology was to set up ephemeris tools to model the

trajectories of the chosen mission profiles. The term ephemeris refers to a table of the predicted

positions of astronomical bodies such as the planets or moon, and by extension, the predicted

positions of artificial satellites.10 The space vehicles modeled in this study can be seen as these

'artificial satellites,' thus determining their trajectories from Earth to the destination planets

depended on complex calculations that took into consideration the orbits of planets at different

dates in the future. Two ephemeris codes originally developed by NASA (IPREP and

CHEBYTOP) were used in order to model these trajectories with ease. By passing the codes

simple inputs based on the reference mission profiles, outputs were obtained that designated the

propulsive requirements for the system architectures. Optimization of the transfer time, stay

time, and departure date inputs was done so as to minimize these energy requirements.

Modeling the architectures of the space vehicles was the next phase of the study. The main

vehicle models created were for NTR, NEP, and Hybrid architectures. These models were then

further broken-down into multiple configurations that varied by fuel type for the NTR system

and thruster type for the NEP and Hybrid systems. The models were generated with Excel

spreadsheets that calculated requirements for power, thrust, propellant load, and vehicle

component sizing. The main consideration in these models was to accurately generate mass

estimates for each system component in order to calculate IMLEO based on the trajectory

requirements input into the vehicle models. Some of the components, such as reactor mass and

propellant varied by mission and vehicle type, but other components such as the Transhab had

constant mass estimates.l Using these spreadsheet vehicle models alone, parametric plots were

generated that characterized the performance of each power and propulsion system architecture.









The final phase of the trade study was actually assessing each of the vehicle architectures

for each of the mission scenarios. The tradespace matrix, which is seen in Figure 1-1 was

defined from the very beginning of the study. It was only after building the vehicle models and

assessing vehicle performance for the given mission scenarios, that vehicle attributes could be

defined. These vehicle attributes then become the parameters used to assess the MOEs that were

established for all of the missions. Finally, the MOEs were rolled up into a single score for each

mission architecture so that the architectures could be ranked in the tradespace matrix. This

ranking procedure was then used to come up with a final 'best answer' for three distinct

categories: each of the 12 mission profiles, each of the three planetary destinations, and for one

overall interplanetary mission.


MARS Jupiter Saturn
Conjunc. Oppos. Conjunc. Oppos. Conjunc. Oppos.
Fast Slow Fast Slow Fast Slow Fast Slow Fast Slow Fast Slow
__ /77_ / / _/ __ / _/ / / F77
77' 7' 77177'/77777 77'777777777777 77


III \ I mCpIIILU I u I I II IIulIal I CIxc Lau I
NTR (Composite Fuel; Thermal Reactor)
NTR (Carbide Fuel; Thermal Reactor)
NTR (CERMET Fuel; Fast Reactor)
BNTR (CERMET Fuel; Fast Reactor)


HYBRID (CERMET; Fast; MPD Thruster) y _
HYBRID (CERMET; Fast; Ion Thruster) 7 7 7
HYBRID (CERMET; Fast;Hall Thruster) 7 7 7 ~ 7 7
NEP( MPD Thruster) 7 7 7 7 77771 77
NEP (Ion Thruster) ~ __ _
NEP ( Hall Thruster) 7

Figure 1-1. Tradespace Matrix









CHAPTER 2
MISSION PLANNING

Astrodynamics

Essential to a successful mission analysis study is the understanding of astrodynamics

fundamentals and how a few specific parameters determine stringent mission requirements. The

definition of astrodynamics in the context of this project is the study of the motion of rockets,

missiles, and space vehicles, as determined from Newton's laws of motion and universal

gravitation. More specifically, it can be seen that astrodynamics deals with the trajectories

spacecrafts will assume on interplanetary transfers. The most basic parameters that stream down

from astrodynamics calculations will, for example, define the best time for a spacecraft to leave

Earth to travel to another destination and how much energy the vehicle will need in order to get

there. Due to the highly important yet complex nature of this field of study, a brief overview of

astrodynamics topics is discussed herein that will introduce basic terminology and concepts that

were integral in the early steps of the mission analysis process.

Orbital Motion

Understanding the transfer methods for interplanetary travel assumes an understanding of

orbiting objects. Orbital motion can be described by a family of curves called "conic sections"

which represent the only paths possible for the orbit of one body about another. The three types

of conic sections include open, closed, and a borderline case. Open conics are those in which the

orbiting body repeats its path, and they only consist of circles and ellipses. The orbits of planets

around the sun, and satellites around the Earth are all elliptical, with one real and one imaginary

focus. The circular orbit is a special case of the elliptical orbit in which the distance from the

orbiting body is constant, and the two foci overlap to create one central point of focus. Using the

energy equation for all conics found by Equation 2-1, where v refers to the orbital velocity, r to









the distance from the orbiting body, [t to the gravitation parameter (GM), and a to the semi-

major axis, the velocity of both an elliptical and circular orbit may be found.

V2 jU U
S2- (2-1)
2 r 2a

For a circular orbit with radius always equal to a, the velocity is found by Equation 2-2.


Ver -- 12 (2-2)


The borderline case between open and closed conic sections is for a parabolic orbit.

Parabolic orbits are rare in nature and an object traveling on one would continue traveling to

infinity until eventually coming to rest when all of its kinetic energy was exhausted. The orbital

speed required to do just this, overcoming the gravitational field of the orbiting body, is called

the 'escape speed'. From the energy equation for a conic section, and the statement that the

energy will be equal to zero at a distance of infinity, the escape velocity is given by Equation 2-

3.


Vesc -2 (2-3)


This equation reveals that the required velocity to escape a planet will be less the farther

the spacecraft is from the planet that it rotates. This is representative of why spacecraft designed

to leave Earth's atmosphere will first be launched to LEO to be assembled, and then escape the

Earth with a much lower velocity requirement.12

It is actually the hyperbolic orbit, categorized as an open conic section, which a spacecraft

would use to escape from Earth. This orbit differs from the closed orbits because the traveling

body does not retrace its path, and from the parabolic orbit because it will have some speed left









over after traveling an infinite distance. A hyperbolic trajectory will only be achieved with a

spacecraft that leaves its orbit at a velocity greater than the escape speed. This will result in

some residual velocity left over when the spacecraft reaches infinity. This residual speed is

referred to as the 'hyperbolic excess speed,' and is found by Equation 2-4, where v, is the

hyperbolic excess speed, Vbo is the burnout speed, and rbo is the orbiting radius at burnout. This

equation was again found via the energy equation for conic sections, where the energy is

constant between the end of the burnout and when the spacecraft reaches an infinite distance.12

2 2 2p/
V Vbo 4b (2-4)
rbo

In the context of interplanetary transfers from Earth, the reaching of infinity is assumed to

be the same as reaching the end of the Earth's 'sphere of influence.' Although a body never

completely escapes the gravitational field of the Earth, it can be assumed to be nearly zero at

some distance from the surface. The sphere of influence (SOI) is said to end when the

gravitational influence on the spacecraft is larger due to the sun than the Earth or other orbiting

planet. For the Earth, this distance has been approximated as 145 times the radius of the Earth.

With respect to the solar system this distance is negligible, but with respect to the Earth it is very

distant. It is in fact so distant, that the velocity at the edge of the sphere of influence is assumed

mathematically to be the velocity at infinity.12'13

Patched Conic Method

A method called the 'patched conic method' combines elements of elliptical, circular, and

hyperbolic orbits to describe the orbital motion a spacecraft assumes in order to complete an

interplanetary transfer. This method allows one to ignore the influence of the sun while the

spacecraft is within the Earth's SOI, to switch to a heliocentric (sun-centered) frame outside of









the SOI, and then to reverse this process upon arrival at the destination planet. The first step in

the trajectory design will be to determine the heliocentric transfer orbit.

An ideal minimum energy transfer to the destination planet would be a simple ellipse,

commonly termed a 'Hohmann transfer.' This heliocentric transfer assumes that the departure

and arrival planets are in circular orbits around the sun with velocity increments tangent to the

planetary orbits, and that the velocity changes occur instantaneously. These high thrust velocity

changes are commonly referred to as "delta-Vs," and constitute the relative velocities between

the respective circular planetary velocities and the perigee and apogee velocity which define the

transfer ellipse.14 For example, the Earth departure delta-V would be the difference between the

Earth's velocity relative to the sun, and the spacecraft's velocity relative to the sun as it exits the

SOI. The delta-V increments necessary to transfer from the departure planet (perigee) to the

arrival planet (apogee) for a Hohmann transfer are given by Equation 2-5 and Equation 2-6,

where Avp is the velocity increment at perigee of the departure planet, Ava is the velocity

increment at apogee of the arrival planet, rl is the periapsis distance of the Hohmann transfer,

and r2 is the apoapsis distance of the Hohmann transfer.15 Note that for an interplanetary transfer

the large center circle in the picture would represent the sun, and the planets would be

represented by the point masses at the outer edges of the blue and red arrows.


Av = ( 2 1) (2-5)
Vr r + r2



AVa= (12-) (2-6)
V 2 V 1 2

It is estimated that using a Hohmann transfer for an interplanetary voyage to Mars would

take approximately 260 days and would require an outgoing delta-V of 2.98 (km/s) for an









instantaneous thrust acceleration.13 Although the Hohmann transfer is ideal due to its minimum

energy solution, it is not always the most practical transfer method. For instance, once a

spacecraft reached Mars after such a transfer, it would have to linger for nearly 6 months before

it could return to Earth by means of another Hohmann transfer. Thus less optimal transfers are

usually taken despite the corresponding higher energy and delta-V requirements.

The delta-V of all heliocentric transfers will be determined based upon the departure date

and the travel time. The future date of departure allows for the determination of the relative

positions of the launch planet and the target planet at time of launch. The time of transfer will

then determine where the destination planet will be in its orbit when the spacecraft gets there.

The path that the spacecraft must follow to successfully intercept the destination planet will

determine the energy of the orbit and the delta-V at both departure and arrival. This delta-V will

always be the difference between the planetary orbit around the sun and the heliocentric

spacecraft speed given by Equation 2-7, where r, is the radius of the planet's orbit, and is the

specific mechanical energy of the transfer.


V,helio = 2 + t (2-7)
P

Optimal departure dates based upon the transfer time and stay time at the planet can be

determined based on ephemeris data that tracks a planets' synodic periods. This is the time it

takes to reappear at the same point in the sky as observed from Earth and relative to the Sun. For

instance, the synodic period for Mars is 2.135 years, thus the best launch opportunities that

would have minimum delta-V requirements would occur approximately every 780 days.12,13

Once the heliocentric delta-V is known, the patched-conic method continues with the

determination of the velocities relative to the planets at departure and arrival. The major









assumption that is made at this point is that the heliocentric delta-V is equal to the speed of the

spacecraft relative to the planet at the SOI. Using the previously introduced term in the

hyperbolic excess speed equation (2-4), the hyperbolic excess speed is obtained by Equation 2-8,

where A vhehocentnc is the heliocentric transfer delta-V, v,,heho is the velocity of the spacecraft

vehicle at escape or capture into the planet, and v ,heho is the velocity of the planet around the

sun.

Vo =A Vheliocentric Vv,helio Vp,helio (2-8)


Since the hyperbolic escape velocity V, can now be determined, equation (2-4) can be

rewritten as Equation 2-9 to determine the speed after injection burn, given the altitude at which

the bum takes place. This speed is essentially the perigee bum on the Earth escape hyperbola.


2 24
V r = Vbo (2-9)
rbo

Since the elliptical or circular orbital velocity before the bum can be determined given the

burn altitude, the actual delta-V experienced by the spacecraft will be that given by Equation 2-

10, where Vbo,planet was the speed after burn and vorb,panet was the orbiting speed before burn, both

relative to the planetary frame of reference.


AV planet bo,planet orb,planet (2-10)

It is important to note that although the delta-Vs calculated in the planetary and

heliocentric reference frames may be similar, they cannot be assumed equal as significant values

may result depending on the specific transfer being used. The delta-V in the planetary reference

is that which will be used to assess the amount of propellant needed onboard the spacecraft,









where the thrust maneuver is assumed to occur instantaneously at the bum altitude.13 A diagram

depicting the parameters introduced for the escape hyperbola relative to the Earth can be found

in Figure 2-1.

The main difference that will result between the hyperbolic escape from Earth and the

subsequent planetary capture, is that the transfer orbit was assumed tangent to the Earth's orbit at

departure, but the capture orbit will most likely cross the target orbit at some angle. This angle

will be taken into consideration along with the target planet's speed and the spacecraft

heliocentric speed to determine the speed at the target's SOI. The speed at the periapsis radius

from the target planet will then be calculated based on conservation of angular momentum.

Special attention must also be paid to the minimum distance from the orbital plane in which the

spacecraft can enter the planet's atmosphere, as entering below this distance will result in

collision with the target planet.12

Although the delta-V requirements are essentially derived from pre-determined physical

laws of nature involving the orbits of planets around the Sun, to a space mission planner they

define the 'cost' of the mission. Once the delta-V values are determined from calculations given

departure date and transfer time details for a specific mission, these values can then be used to

calculate how much propellant would be required given a specific type of propulsion system.

The 'rocket equation' is used to perform this calculation, enabling one to determine not only

propellant requirements, but essentially the entire spacecraft mass at the launch altitude. In the

space business, mass means money, thus the connection can be made between delta-V

requirements that stem from orbital mechanics and final monetary cost of any space mission.14

More specifics related to calculating propellant and system mass based on delta-V requirements









will be found in Chapter 3, which discusses the actual sizing of the spacecraft systems under

consideration in this trade study.

Mission Profiles

The difficulty in defining a mission profile is obtaining the seemingly best overall

approach given the many variables that play into mission design. Mission characterization

involves defining a mission concept which will describe how the mission will work in practice,

the mission operations which will detail how people will operate and control the mission, and the

mission architecture which links the mission concept to the major mission components.16 This

study was focused on developing mission concepts and mission architectures such that the

various power and propulsion systems being analyzed could be compared under different

mission scenarios. Mission operations was not a major focus in this study since the trade was

essentially made on the capabilities of the power and propulsion architectures to transport

humans, and was not concerned with human activities during the transfer periods.

The mission concept was essentially broken down into different mission profiles, each of

which characterized the specific parameters of departure date, destination planet, outbound and

return transfer times, and planet stay times (or class). The use of planetary swing-bys, and the

re-entry method at Earth are two other considerations that typically go into a mission concept but

they were not characterized in the mission profiles since all profiles used the same methods in

this regards.

The term 'mission architecture' was used in this study to refer to one of the three vehicle

architectures (NTR, NEP, and Hybrid) being coupled to a specific mission profile. The

motivation behind generating a large number of mission architectures is to obtain a better

understanding of the performance of each vehicle under different scenarios. Since it would be

inefficient to attempt to model every type of scenario for every vehicle architecture, the mission









profiles were configured in such a way that most of the subset of possible desirable missions

would fall into one of the mission categories. This was achieved by setting upper and lower

limits on quantifiable mission profile parameters, and then collectively optimizing these

parameters for each mission architecture.

The main driver behind assessing vehicle architectures given different mission profiles was

to see if one system was decisively superior to the rest for all profiles examined, or if instead

different systems were the optimal choice for specific mission profiles. These results would be

significant as they would highlight the importance of either focusing development efforts on one

vehicle architecture for all possible future manned missions, or focusing first on the mission

profile decision and then following with the appropriate vehicle design.

Considerations for Mission Planning

Mission trajectory

The trajectory chosen for each mission profile is dependent on parameters that determine

the launch window. These include the departure date, outbound and return transfer times and the

stay time on the planet. The latter parameter takes on a significant role in trajectory analysis,

often being divided into the two categories of "short-stay" opposition-class trajectories, or "long-

stay" conjunction class trajectories.

The opposition-class mission is generally characterized by short stay times on the order of

40-60 days, and round-trip Mars missions that range from 365-660 days. Most opposition-class

Mars missions take advantage of a Venus swingby on the return trajectory to Earth, though

discussion of this maneuver is briefly withheld. Other characteristics of this type of trajectory

include large propulsive energy requirements and the combination of both a short and a long

transit leg. Clear disadvantages to this method that have been noted include high variances in









energy requirements given departure date, and large spikes in escape and capture delta-V's due

to decreased transit time.2

The conjunction-class mission typically has stay-times between 400 and 600 days, and

total Mars mission trip times on the order of 900 days. These trips may be the more highly

desired of the two trajectory classes, since the stay time on the planet is much longer without a

significantly longer transfer time. In fact, relatively short transfer times on the order of 200 days

have been examined in previous studies for these mission trajectories. In addition, compared to

the high energy requirements of the opposition-class missions, the conjunction class missions

typically represent "minimum -energy solutions" for given launch opportunities.2

The use of a Venus swingby periapsis burn has been incorporated into methodologies for

opposition-class Mars missions, as mentioned previously. This swingby maneuver provides a

change in the spacecraft's heliocentric energy, which diminishes the delta-V requirement to enter

back into Earth's atmosphere. The maneuver typically requires a very small propulsive burst,

but will result in an overall reduction in propellant requirements. Past studies of this trajectory

type for Mars missions have indicated that adding a small propulsive maneuver during the

swingby increases "mission flexibility," and that a transfer at the periapsis of the Venus orbit is

"close to the optimum transfer point." Note that the periapsis burn can be in either the direct or

retrograde directions, such that the relative velocity may increase or decrease, respectively.17

Mission approach

Two approaches are also recognized for mission design: an all-up mission and a split

mission. All-up missions are those in which cargo and crew leave the orbit of the Earth at the

same time. In a split-mission design, cargo is flown to the destination planet first, and then is

followed by the crewed vehicle. One benefit of this design is that the cargo can be sent on a low-

energy trajectory to the destination planet and assure that supplies will be there to greet the crew









when they arrive. In addition, this method will reduce the payload of the crewed vehicle, which

already requires more energy due to the need to be sent on a faster trajectory to reduce time in

the radiation-filled vacuum of space. On-orbit assembly is a possibility for either approach, and

allows for different vehicle components to be launched into LEO on separate heavy lift launch

vehicles (HLLV). The assembly could utilize automated rendezvous and dock between

subcomponents, or the use of a space station or other docking facility.2

The capture methods at both the destination planet and following return to Earth are also a

consideration in the mission approach. Both propulsive capture and aerocapture methods have

been used in past studies for capture into the destination planet or moon. Propulsive capture

imposes a delta-V requirement on the propulsion system, thus necessitating a propellant

requirement for the propulsive burn. The aerocapture method instead uses the planet's

atmosphere to slow down the vehicle. This flight maneuver uses the friction from the dense

atmosphere of the destination planet or moon to slow down the spacecraft, thereby transferring

the energy from the high spacecraft speed into heat. This method thus saves on propellant load,

yet requires advanced heat shielding to protect the craft. An orbit elevator has also been

proposed as a method by which the spacecraft can move in order to reach various orbits once it is

already in its destination orbit, without the use of a propulsion system.6

Propulsive capture or direct Earth re-entry are the two methods suggested for return to

Earth. The propulsive capture, as its name suggests, uses a propulsive bum to capture into an

orbit around Earth. The other method, direct Earth re-entry, assumes that the spacecraft skims

the desired Earth orbit, without actually capturing into it through any propulsive means. The

Earth Crew Return Vehicle (ECRV) would then be released from the spacecraft allowing the

crew to descend back to Earth 'Apollo-style', with the remaining spacecraft hardware being









released into space.18 One of the considerations then in choosing between the methods is

whether or not the interplanetary spacecraft should be expendable.

Mission Profile Selection

A primary objective in the selection of specific mission profiles for this study was to

choose missions that represent, as much as possible, the entire subset of interplanetary missions

that may be desirable in the future. This was accomplished by incorporating different planetary

destinations, using the same transfer and capture methodologies for all missions, and by

allocating adequate ranges for the mission profile parameters. Through optimization of all

mission profile quantifiable parameters, singular data points were generated for each profile,

including an exact date of departure, number of outbound and return transfer days, and specific

number of stay-time days on the planet. Thus, although seemingly specific missions will be run

for the vehicle architectures, each mission will represent a larger subset of possible missions and

will reveal the best performance data for each transportation system.

The decisions that were made in the mission design process were influenced by two major

sources. The first very influential source was that of the set of literature found on past

interplanetary studies. Differences between methodologies arose, such as assuming delta-V

values for a mission versus using ephemeris codes to solve for them based on a specified launch

window. These literature sources also contained information on NASA DRMs that was

particularly useful in the mission planning process since they provided a comprehensive outline

of various mission concepts.

The second source that drove mission profile design was that of the ephemeris tools

themselves. Since a low-thrust and high-thrust ephemeris tool were both being used, and each

tool had some unique capabilities and disabilities, constraints were set from the very beginning

on possible mission platforms. For example, either the propulsive capture or direct Earth re-









entry methods could be designated for the high-thrust ephemeris tool, yet the low-thrust tool

assumes a capture into Earth orbit, thus only the propulsive capture option could be used if

commonality between mission profiles was to be maintained.

Mission trajectory

In order to increase the breadth of the mission profiles, and thereby increase the

applicability of this study to future research interests, three planetary destinations were

attempted, each with two different stay times and two transfer times. Many advanced propulsion

mission studies have been done using Mars as the planetary destination. In terms of present-day

space policy and thought, this is directly in-line with the 'Moon, Mars, and Beyond' initiative

proposed by George W. Bush in 2002. But there have also been various studies on mission

analysis for farther-out missions with destinations including Jupiter's moon Callisto, Pluto's

moon Charon, and objects in the Kuiper Belt.18'19

For this study, it was decided that Mars would be a logical choice for a planetary

destination, and that Jupiter and Saturn would also be investigated in order to test the vehicle

architectures against higher-energy requirements. The spacecraft orbit around Mars was a 250

(km) by 33793 (km) elliptical orbit, which is comparable to 1 solar day, and was taken from

Borowski's Mars mission study ground rules and assumptions list.18 The orbits around Jupiter

and Saturn will include some basic assumptions. The spacecraft itself will be assumed to be

orbiting Jupiter's moon Ganymede, and Saturn's moon Titan, as the spacecraft would not easily

find a safe orbit around the planet itself. However, the ephemeris tool input will include the

planet as the destination orbiting body and the orbiting distance will be that of the distance from

the planet to the orbiting moon. This amounts to the spacecraft seemingly taking on the orbit of

the moon around the planet, although it will actually be in a low orbit around the moon, with the

spacecraft and moon both being in orbit around the planet. Although some studies model the









three orbit insertion components (propulsive capture at radius of planet's moon, plane change

into moon's orbit, and propulsive capture into circular orbit around moon), the results of this

study are only able to account for the first component, and assume the other two components

negligible.19 The orbits that will be used in this study include a circular orbit at 998600 (km)

from Jupiter, and a 1162000 (km) orbit around Saturn, which again are synonymous with the

orbit of the selected moons around their respective planets.

Since both short and long stay times may be desired for future interplanetary missions

depending on the operational nature of the mission, both conjunction and opposition-class

missions were used for the mission profiles. A range was applied to the stay time so that an

optimization of the stay time could be found as part of the overall effort to find the minimum

delta-V requirements for each mission. It was decided that the opposition-class missions would

have a stay time range of 40-60 days, and conjunction-class missions would have a range of 400-

600 days, as these values are considered standard among mission analysis literature sources.5'12

Note once again that a singular data point will be found for the stay time after optimization,

despite the fact that the real CTV would actually be designed with sufficient propellant for a

somewhat longer/shorter stay to account for a safety margin.

Although high and low energy missions are already broken down as opposition and

conjunction-class in the mission profiles, it was decided that short and long transfer times should

also be included due to their own effect on energy requirements and safety concerns. Four of the

central issues for mission profile selection are said to be crew radiation exposure, crew time

spent in zero g, the component of mission risk that increases with mission duration, and added

cost of shortening trip time.20 Each of the first three of these components condones shortening

transfer times, though the last highlights the great expense, both in fuel and in dollars, of shorter









missions. It thus seems reasonable to want to analyze transfer times on both the low and high

ends of the spectrum.

Having two data-points for transfer time is also practical from an energy consideration

perspective. For instance, since a decrease in transfer time is generally associated with an

increase in delta-V requirement, the optimized point could likely be that of the longest transfer

time. If this transfer time is longer than that considered acceptable by future mission planners,

the dataset for this particular DRM may be deemed useless. It was thus found both fruitful and

necessary to include both short and long transfer categories in the breakdown of design reference

missions.

The transfer time ranges that were determined for this study's mission profiles were

decided by using past studies to determine typical transfer times and by trying to make

simplifying assumptions. The first assumption that this study made was to assume that the

outbound and return transfer times were equal. This was done to minimize time spent on

mission optimization using the ephemeris tools. Secondly, since the transfer times were to be

broken down into two groups, it was decided that the Mars transfer ranges would be 100-200

days for the short transfer range, and 200-300 days for the long transfer range, allowing for an

even split of 100 days for each. In a study that compared 21 different crew and cargo mission

studies, it was found that outbound transfer times ranged from 80 days to 335 days, with return

transfers being the same or somewhat quicker.2 It was thus felt adequate that the profiles

covered the 100-300 day transfer range, which equates to nearly 80% of the range covered by

numerous other studies.

In regards to the Jupiter mission, Ehricke has been cited as using a 640-day transfer for a

Europa mission, but many sources show a more practical 1000-day transfer to another of









Jupiter's moons, Callisto.9'19 Using the 1000-day transfer as a mid-point, the short transfer range

was decided to be 800-1000 days, while the long transfer range would be 1000-1200 days.

These categories are applied to both the Jupiter and Saturn mission as no sources were found for

crewed mission profiles to Saturn. In addition, research that was done previous to this study

showed similar trends for delta-V vs. trip time for Jupiter and Saturn destinations.

Mission approach

The current study serves to assess multiple vehicle architectures under a spectrum of

mission profiles for the crewed vehicle component of a split-mission architecture. It is assumed

that an unmanned cargo vehicle departs Earth orbit on a minimum energy, one-way trip to the

destination planet, carrying science payloads and any other payloads that would be necessary for

planetary exploration, but not necessary for the outbound and return trips on the CTV. Proposed

additional functions of the pre-deployed cargo vehicle noted in the literature include deploying

an unmanned semiautonomous rover to explore the landing site region prior to crew arrival, and

deploying navigation beacons to assist in landing the primary cargo payload.21 Beyond these

general assumptions about the cargo vehicle, no aspects of it are modeled or considered in this

mission analysis study.

The mission concept used herein assumed that the hardware components of the crewed

vehicle were assembled in LEO prior to departure. The components would first have to be

launched from Earth on HLLVs, and then assembled by means of an automated rendezvous and

dock between elements, or a more complex method involving the space station or other

construction facility.2 After assembly, the crew would be transferred from the International

Space Station (ISS) or a similar post to the CTV. Thus missions assessed in this study began

from an operational sense at the point were the assembled CTV was ready to leave LEO. In









staying consistent with orbital parameters found in Borowski's studies, the actual LEO altitude

the vehicle was assumed to be in was a circular Earth orbit of 407 (km).18

Certain aspects of the mission concept that were kept constant for all mission profiles

included major CTV components, and planetary and Earth capture methods. It was determined

that the CTV would only carry the mass needed for a full two-way trip to and from the

destination planet. Although this did require carrying both outbound and return propellant from

LEO, the propellant tanks could be dropped upon completion of major burn segments. The

propulsive capture method was used for all captures, both at the destination and upon return to

Earth to accommodate the capabilities of the ephemeris tools and still maintain the means for an

'apples-to-apples' comparison between architectures. The Earth orbit insertion maneuver

occurred at the perigee of a 500 (km) by 71165 (km) orbit, again in accordance with Borowski's

standard ground rules. The mission analysis did not consider anything past this point in the

mission, though it may be assumed that the crew would depart the ship at this point and perform

a r-eentry into Earth's atmosphere in the ECRV capsule.19

Mission profile summary

To provide a condensed account of the mission profile assumptions to be used for the

reference missions analyzed in this study, a list of the basic mission trajectory and mission

approach ground rules is given below.

* Delta-V requirements for each mission were determined through an optimization
methodology using high-thrust and low-thrust ephemeris tools

* Outbound and inbound transfer durations were assumed equalMars departure dates ranged
between 2030-2035; Jupiter and Saturn dates ranged from 2040-2045

* The following orbits were used for Mars, Jupiter, and Saturn:

* 250 (km) by 33793 (km) elliptical orbit at Mars
* 998600 (km) circular orbit at Jupiter (assume orbiting moon Ganymede)
* 1162000 (km) circular orbit at Saturn (assume orbiting Titan)









* A split-mission scenario was used for all missions; only CTV will be modeled

* Earth Orbit Rendezvous and Dock Vehicle Assembly at 407 (km) occurred after HLLVs
transport spacecraft components to orbit

* All CTV propellant for round-trip mission was carried onboard vehicle

* Propellant tanks were dropped after completion of major burn segments

* Propulsive capture method used to capture CTV into planetary orbit

* Earth Orbit Insertion of CTV occurred at perihelion of 500 (km) by 71165 (km) orbit

* TransHab was sized for a crew of 6, with consumables accounting for the total duration of
mission

The optimization of the trajectories using the ephemeris tools incorporated the transfer and

stay time ranges discussed in the Mission trajectory section. These ranges and their associated

design reference missions are given in Table 2-1. The first three categories of "Planet

Destination," "Mission Class," and "Transfer Type," all characterize the 12 basic mission

profiles. The outbound and return transfer duration and stay time are both given as a range in

days, and provide the bounds for the transfer type and mission class, respectively.

Ephemeris Tools

In recognition of the high-thrust NTR architecture and low-thrust Hybrid and NEP

architectures modeled in this study, the IPREP and CHEYBYTOP ephemeris tools were both

needed in order to model the interplanetary transfers for the missions previously discussed.

IPREP (Interplanetary PREProcessor) is a rapid grid-search optimizer for launch and

arrival windows, delta-V, and mass originally created by Martin Marietta Astronautics.22 It is a

commonly used tool for estimation of high-thrust trajectories, thus it was the logical choice to

model the trajectories for the NTR architecture in this study. Inputs to the program generally

include the order of the planets to be encountered, maneuvers to be performed at each encounter,

and the time-of-flight window for each mission segment. IPREP then calculates the delta-V









energy requirements using a patched-conic technique that assumes the planets to be point-masses

and finds the position and velocity of each planet from one of the ephemerides. A transfer orbit

for each leg of the trajectory is then found by solving Lambert's problem.23

CHEBYTOP (Chebyshev Trajectory Optimization Program) is a program built in the late

1960's by the Boeing Company, which provides a two-body, sun-centered, low-thrust trajectory

optimization and analysis. Its capabilities extend to preliminary mission feasibility studies such

as the one undertaken. Its strengths include requiring a small number of simple inputs and

having fast run times for quick interplanetary mission modeling. CHEBYTOP uses the

CHEBYCHEV Optimization Method, which is a series of approximations to the control problem

that breaks it down into a group of classical calculus optimizations.24'25 The algorithms used are

only relevant to spacecraft with very low thrust-to-weight (T/W) ratios, which makes it an ideal

modeler for NEP systems. This code has been used in numerous studies done in the past on low-

thrust trajectory systems such as solar electric propulsion (SEP) and NEP for Mars missions and

for other planetary destinations. NASA Marshall Space Flight Center (MSFC), for example,

used the code to generate thrust-to-weight (T/W) versus delta-V curves for an NEP mission to

Pluto.25 Besides its use for the NEP architecture in this study, it was also used for the Hybrid

architecture due to its capability to model high-thrust NTR burns at escape and capture, with an

NEP mid-trajectory low-thrust bum.

The transfer methodology used by each of the ephemeris models developed with these

codes was slightly different depending on which propulsion system was being modeled. The

IPREP code used to model the high-thrust NTR propulsion systems assumes one bum at Earth

Escape, a second bum at Planetary Capture, a third bum at Planetary Escape, and finally a fourth









burn at Earth Capture. These burns can be thought of as instantaneous burns for modeling

purposes, as is commonly done in traditional delta-V calculations for high-thrust orbit transfers.

The NEP system that was modeled with CHEBYTOP used a spiral trajectory out of

Earth's orbit and a spiral capture into orbit at the destination planet. This is due to the fact that a

low-thrust engine cannot achieve the necessary escape speeds in a very short amount of time as

the NTR systems do. Note that the transfer time designated by the user in the IPREP input file

did not account for the number of days needed to spiral out of or into a planetary orbit, but only

accounted for the days on the interplanetary trajectories.

The Hybrid system, which was also modeled with CHEBYTOP, assumed an NTR bum to

escape Earth's orbit, an NEP burn for the majority of the transfer duration, and an NTR bum to

capture into the destination planet's orbit. This same method was then used in reverse to escape

from the planet and return to Earth. Even though the Hybrid and NEP systems used the same

ephemeris tool, they were indeed modeled in separate ways. Due to features inherent to both the

CHEBYTOP and IPREP ephemeris codes and the manner in which the scripts were set up for the

ephemeris models generated in this study, the delta-V requirements given by the output files took

into account the different trajectory methods used by each vehicle.









Table 2-1. Design Reference Mission Categories
Planet Destination Mission Class Transfer Type Out/Ret Tran. (days) Stay Time (days)
Fast 100-200 40-60
Oppos. Slow 200-300 40-60
Fast 100-200 400-600
Mars Conjunc. Slow 200-300 400-600
Fast 800-1000 40-60
Oppos. Slow 1000-1200 40-60
Fast 800-1000 400-600
Jupiter Conjunc. Slow 1000-1200 400-600
Fast 800-1000 40-60
Oppos. Slow 1000-1200 40-60
Fast 800-1000 400-600
Saturn Conjunc. Slow 1000-1200 400-600


-^ *


,SOI


V

Vr


Figure 2-1. Escape Hyperbola









CHAPTER 3
MODEL DEVELOPMENT

Space Reactor Background

Reactor Configuration

All reactors rely on the principle of thermal energy production from the fission process of a

fissionable atom such as 235U. This energy production results from the conversion of the kinetic

energy of fission fragments and neutrons to heat after slowing down from collisions and

interactions with other atoms. This thermal energy is then transferred to a coolant that flows

through the reactor core. Space reactors will typically use a lightweight gas such as hydrogen as

both the coolant that extracts heat from the reactor, and in the case of a nuclear thermal rocket,

the propellant that immediately thereafter is shot out of the rocket nozzle to create momentum.26

Propellants with low molecular weights are most effective for thermal propulsion as they

produce the highest specific impulse.27 Space reactors that are specifically used for propulsion

are known for having high specific impulse and high thrust levels, providing a clear advantage

over alternatives such as chemical propulsion.

The actual configuration of a space NTR system is similar to that of a chemical system,

except for the reactor heat source. The hardware consists of a reactor, propellant tank, radiation

shielding, a feed system, and a nozzle. The major reactor components consist of the radial

reflector, reactor pressure vessel, moderator, fuel-element assembly, and control drums. The

reflector surrounds the outside of the core and functions to reflect neutrons produced in the chain

reaction back into the core, helping to maintain a controlled chain reaction. The pressure vessel

is needed in order to maintain reactor pressure and must be made of an aluminum or composite

material that will withstand the high radiation, heat flux, and pressures from the reactor. A

moderator material is used in a thermal reactor to slow the neutrons produced from a nuclear-









fission reaction to energies in which they are more apt to undergo another fission. The fuel-

element assembly contains the actual heat-producing uranium fuel, along with the flow channels

for the coolant. Control rods are also found in the core and serve to absorb neutrons to decrease

neutron population and maintain the ability to control the reaction rate or even shut down the

reactor.26

Reactor Fission Spectrum

A major choice in overall reactor configuration concerns the type of fission spectrum it

will operate under. It may operate using fast neutrons produced by fission reactions (a fast

reactor), or neutrons slowed to thermal energies and thus more likely to produce subsequent

fission reactions (a thermal reactor). A fast reactor functions based on a chain of reactions

propagated by high-energy fission neutrons. The low probability of fast neutrons producing

fission reactions results in a large fuel requirement for this type of reactor. However, highly

concentrated fuel will allow a compact reactor, typically of smaller size and mass compared to a

thermal reactor. The probability of neutron capture and consequent fission is much higher for

thermal energy neutrons that have been slowed by interaction with moderator materials than for

fast neutrons. Although thermal reactors are typically larger than fast reactors, they require

much less reactor fuel than fast reactors and also a less complicated fuel element and core

design.28

The reason that both fast and thermal reactors have been considered for space reactors is

that they are inherently better equipped for different types of missions. The fast reactors are

preferred for long-life, low power operations since the high fissile loading allows high total

energy operation. Missions that require large bursts of power but small total lifetime energy may

benefit more from a thermal reactor. The relatively long neutron lifetime and large delayed

neutron fraction found in thermal reactors would help maintain precise control of burst power.27









Nuclear Thermal Power and Propulsion

History of Nuclear Thermal Rocket (NTR) Systems

This history of the NTR engine began in 1953 when the ROVER program began at Los

Alamos Scientific Laboratory in order to develop a reactor for the operation of a nuclear rocket.

The major reactor series that went through design, build, and test phases during this program

included KIWI, Phoebus, Peewee-1, and Nuclear Furnace-1.28

The KIWI reactor series holds the accolade of being the first NTR reactor built and tested.

It allowed for advances to be made in the areas of instrumentation and control, fuel element

design and fabrication, structural design, and testing techniques. The Phoebus reactors that

followed had design specifications intended to meet the need of interplanetary propulsion

systems, with special focus on manned missions to Mars. The major results from research on

this reactor series included control of rocket parameters over a wide range of operating

conditions, along with finding that large nozzles for NTR applications was feasible. The Peewee

reactor series came next chronologically, and was meant to investigate performance

characteristics of a smaller reactor. This was followed by the last stage of ROVER, the Nuclear

Furnace series, which was designed to test advanced fuel elements containing composite fuel.28

The end of the research-focused ROVER program led to the start of the Nuclear Engine for

Rocket Vehicle Applications (NERVA) program, which focused more heavily on concept

development. During this 11-year program, the NRX reactor series was developed,

incorporating the non-nuclear system components (propulsive components) into the reactor

designs developed during ROVER. The NRX-XE' engine was the main focus as vertical

downward firing tests of the engine in a simulated space vacuum were conducted. This allowed

for investigation of the engine start-up and shutdown characteristics, along with the resulting

engine performance parameters. Over the course of this program, more than 20 NTR reactors









were built and tested at the Nuclear Rocket Development Station at Nevada's Nuclear Test Site.

Although the NERVA program was canceled in January of 1973 due to a change in national

priorities, the most poignant outcome of the work done during ROVER/NERVA was the

confidence that an NTR engine could be developed to meet the objectives of structural integrity,

restart capability, predictability, control, and reliability.28

Fundamental Research

Two of the primary areas of current research for NTR systems are reactor fuels and reactor

operational capabilities. Research of space reactor fuels is of high importance due to the fact that

the reactor is heating a coolant for propulsive purposes. Increasing the operating temperature of

a space reactor fuel thus has implications for overall propulsive performance. Reactor operations

are also being analyzed in terms of providing dual-mode functions instead of a single mode, the

implications of which may have profound effects on overall spacecraft capabilities.

Improvements in the area of fuels research have been consistent since the ROVER years, while

dual-mode bimodal reactor configurations have only been studied in recent years. It is important

to note that for both areas of research, however, no full-scale hardware testing has been

undertaken for nearly 35 years.

The four main types of reactor fuels that have undergone serious consideration for space

application include graphite, composite, and carbide fuels for thermal reactors, and fast reactor

CERMET fuel. Graphite fuel is the oldest and most mature, as it was studied and used during

the ROVER/NERVA programs. The original ROVER engine had rods 54 inches in length, with

a mixture of uranium, zirconium, and carbide in a graphite matrix.29

Both the composite and carbide fuels were developed based upon experience gained with

graphite fuels, yet they only underwent minimal testing near the end of the ROVER/NERVA era.

The composite fuel composition differed from the graphite fuel in that it was a 'composite' of the









UC2 and ZrC used to make up the fuel pellet and coat in the early generation coated-particle

matrix graphite form. It was found that the ZrC coating increased both the lifetime and the

integrity of the fuel as it protected the fuel from the hot hydrogen propellant. Carbide fuels, on

the other hand, eliminated the protective carbide coating required for matrix fuels.

Improvements were made with carbide fuels since it was found that the composite fuel coating

severely limited the endurance and temperature performance of the fuels.

Since the termination of the ROVER/NERVA programs, many advances in fuels research

have been made. Some stem from improvements in other related areas such as materials

research, and include an increase in hydrogen turbopump efficiency, improvements in titanium

pressure vessel manufacturing techniques, and improvements in nozzle cooling. The term

applied to updated NERVA NTR engines that use these latest fuel technologies available is

NERVA Derivative Reactors (NDR). NDRs typically have carbon-based matrix fuel elements

with graphite moderator and ZrH moderator sleeves in the support structure. Typical chamber

temperatures for NDR graphite, composite, and carbide fuels are 2500, 2700, and 3100 (K)

respectively, while ISP values are 885, 921, and 1020 (s) respectively.28

Bimodal reactor designs, which have been under development only more recently, have

consistently been designed based on the use of CERMET fuel. This is primarily a result of both

CERMET fuels and bimodal designs being based upon fast fission reactors. CERMET fuel is

termed according to its ceramic metallic formulation, and is primarily configured of U02 fuel

encased in tungsten and tungsten-rhenium alloys.30 Properties of CERMET fuel include high

strength, thermal conductivity, temperature capability and burnup, in addition to giving reactors

a long operating life and the ability to restart. It is important to note that use of CERMET fuel

and fission reactors is not limited to bimodal reactor configurations.









Bimodal reactors, named appropriately for their two modes of operation, have been kept

behind their unimodal counterparts due to the significant costs associated with their testing and

production. Their discussion, however, reaches back to the ROVER/NERVA program, during

which the potential benefits of such reactors were recognized. The basic reactor design used for

that program was assessed for possible modifications that would allow electric power generation.

In more recent years, the Air Force Phillips Laboratory has conducted studies on bimodal reactor

designs for possible military applications. The DOE Office of Nuclear Energy has since

collaborated with the Phillips Laboratory in order to develop bimodal bus designs along with

initial performance requirements.30

Bimodal reactors have increased complexity in both design and function due to their dual-

mode operation. During the power and propulsion modes, the reactor must operate under very

different sets of conditions and must perform extremely different functions. In addition to these

capabilities, a space reactor is expected to have both high reliability and a long lifetime. The

engine must function in cooperation with more hardware than a unimodal engine, given that a

power conversion unit and heat radiator are required to produce electric power and get rid of

waste heat. While the amount of electric power generated is the important performance

parameter in power mode, the thrust-force and specific impulse are the driving parameters in the

propulsion mode.

In a typical bimodal reactor design, the core will consist of heat pipes and CERMET fuel

with numerous propellant channels. In the propulsion mode, liquid hydrogen may be run

through reactor components for cooling purposes, and will then run through the CERMET fuel

element one time before expanding out the rocket nozzle. In the power mode, the heat pipes will

serve as the energy transport medium from the reactor fuel to the power conversion system. A









small gap between the fuel elements and heat pipes serves to allow for thermal radiation between

them, creating the primary energy transfer mechanism for the power mode. A working fluid

such as sodium or xenon then runs through the heat pipe, transferring the energy to the power

conversion unit.

A reactor design by the Phillips Laboratory uses 93% enriched CERMET fuel with finned

heat pipes. The CERMET fuel elements are nine-sided blocks with 52 axial propellant channels.

They are 59.5% UO2 by volume and 40.5% tungsten. The heat pipes provide both energy

transport and structural support, which relieves the need for tie-tubes. This reactor design was

said to have a 10 (kW) electric power output capability with a 10-year lifetime, and produced

220 (N) of thrust with a specific impulse of 825 (s) for the propulsive mode.30

Development of the NTR Model

The NTR architecture, which utilizes a nuclear reactor for thermal propulsion, was one of

the three main architectures studied in this project. This architecture was broken down into five

systems that all achieve NTR propulsion through different selections of fuel and reactor types.

The five NTR systems that were assessed in the tradespace matrix are found in Table 3-3, which

categorizes each system according to reactor mode and energy spectrum along with fuel type.

Although the NTR systems were broken down according to characteristics of the space

reactor, it is important to understand how the reactor functions within the entire spacecraft. The

term 'architecture' itself was used in this study to describe not only the reactor, but also each of

the subcomponents that makes up the entire CTV that will carry humans to an interplanetary

destination. Thus it should be noted that the NTR architectures were all based upon the same

basic principles and had almost identical vehicle schematics. The exception to this generality

within the NTR architecture group was the one system operating with a bimodal reactor. While

unimodal reactors' main functionality was to provide heat to the hydrogen coolant that is









expelled through the rocket nozzle, the bimodal reactor had a secondary coolant loop that

provided electric power to the spacecraft. Within the group of unimodal reactor systems, the

only real differences were in the physical makeup and operation of the reactor itself.

The four unimodal reactor systems can all be described using the schematic found in

Figure 3-1. The divisions of the entire CTV were broken-down into power, propulsion, and

spacecraft categories. This schematic was intended to provide an overview of how the

components within these groups interacted, along with a very rough idea of where they were

physically located respectively within the space vehicle structure. The solid black lines depict

flow of hydrogen propellant from the tanks all the way through its exit from the rocket nozzles.

The dotted black lines depict transfer of electricity from the fuel cell power source to the

components that require electricity for functionality. The power and propulsion system boxes in

the schematic are specific to the NTR unimodal systems, and will thus be discussed in the

following section. The general spacecraft components, indicated in yellow, will be discussed

later in the text, as their modeling was not dependent on the type of propulsion and power system

architecture.

The one NTR system that used a bimodal reactor had a significantly altered architecture

schematic as shown in Figure 3-2. As can be seen in the schematic, fuel cells utilized in the

unimodal systems are no longer used for onboard spacecraft power needs. In addition, power

conversion units, power management and distribution (PMAD) systems, and radiators were all

added to accommodate the reactor power generated by the bimodal reactor, and a secondary

propellant was added for power conversion purposes. The solid black lines in the figure again

indicate propellant flow, while the dashed lines indicate flow in the power distribution process.









Power

One of the ways in which an NTR reactor will differ from a purely NEP power reactor is

that the required thermal power will be defined based on the desired temperature and mass flow

of the propellant instead of the mission energy requirements. The derivation of this power can


be found through both Equation 3-1 and Equation 3-2, where mprop NTR is the rocket mass flow

in (kg/s), T is the thrust in (N), ve NR is the exit velocity of the propellant in (m/s), P,_prop is the

thermal power in (MWt) for the propulsive mode, and tH2 is the temperature in (K). It is

interesting to note that the thrust value, which is needed to determine the power requirement, is a

parameter that is chosen by the designer. In this study it was chosen to be 15 (klbf), which is on

the lower end of typically considered thrust ranges. Thus it can be reasoned that the power is

determined by both the designer and the properties of the fuel itself.


mprop_NTR T= (3-1)
/e NTR


P prop = prop NTR* (0.018061 *tH2 -5.715417) (3-2)

After the power required of each type of reactor fuel core for a thrust of 15 (klbf) was

calculated, the actual mass of the NTR power and propulsion system had to be determined. The

basic components consisted of the reactor, pressure vessel, internal and external radiation shield,

and propulsive hardware. The propulsive hardware was further broken down into the three main

components of the nozzle, turbopump assembly, and nonnuclear support hardware such as lines,

valves, actuators, and instrumentation thrust structures. The data-points for these three

propulsive hardware components were taken from an SAIC report for the 'SAIC ELES-NTR'









design. The mass estimates for the three components were 421 (kg), 104 (kg), and 1264 (kg)

respectively, which gave a total of 1789 (kg) to be added to the reactor and shielding masses.31

The mass of the reactor and pressure vessel were estimated from SAIC plots, which gave

mass as a function of power, pressure, and temperature. Linear interpolation tools were

developed in order to be able to confidently generate mass estimates based on the characteristic

parameters of each fuel. The fuel parameter values that were used for the estimation of the

reactor mass included chamber temperature, ISP, and calculated power, all of which are seen in

Table 3-2. Although the interpolators required the three inputs of temperature, pressure, and

power, the ISP was also needed for the calculation because the mass flow parameter that defined

the power requirement was dependent on ISP. Also, the pressure used for all cases was assumed

constant at 1000 (psia).

Once the reactor masses had been calculated, the internal and external shield masses were

calculated based on these values. It was decided to estimate the shielding masses by using ratios

of shielding to reactor mass found from data-points in the literature. It was thus determined from

data that characterized a 75 (klbf) Westinghouse/NERVA design, that the internal shielding was

to be 26.96% of total reactor mass, and external shielding 80.27% of reactor mass.28 Calculating

the total NTR power and propulsion system mass was then accomplished using Equation 3-3,

where f,,, h,I~d was set to 29.6% and fExt h,,eid was 80.3%.


NTR = reactor + m propul. + reactor (fInt shield + Ext held) (3-3)

The reactor and shielding masses for each type of fuel, along with the total NTR system

mass for each fuel is given in Table 3-3. Note that although only one reactor was used to carry

out power and propulsion requirements, an extra reactor system equal in weight to the primary

reactor was carried on board for redundancy.









A unimodal NTR architecture uses all of the heat generated from the reactor for heating of

the rocket propellant. Any electric power needed by the spacecraft had to come from an

alternate source. Due to proven performance on the Space Shuttle, fuel cells were chosen for

electric power generation for the NTR unimodal system. Fuel cells are self-contained generators

that operate continuously without the need for sunlight. They work by converting the chemical

energy of an oxidation reaction to electricity. An added benefit of their use is that they actually

produce water as a byproduct, which can the be used for drinking.16 Two key parameters of fuel

cells used in the calculation of their mass were the specific power of 275 (W/kg), and the lifetime

estimate of 100 days. The equation used to calculate the mass of fuel cells based on electric

power requirement and the mission duration is given by Equation 3-4, where P, is the electric

power in (kWe).

masspe, cl = (1/.275) P, (mission days /100) (3-4)

The bimodal reactor provided both the hydrogen propellant heating and electric power

requirement for the spacecraft. Thus, although the actual mass of the bimodal reactor was on

average higher than that of a unimodal reactor, no extra mass was assumed for power generation.

There was, however, mass associated with the power conversion that is required to convert the

reactor's thermal energy to electrical power. The three main components required to fulfill this

function include the PMAD system, the Brayton Conversion Unit, and the radiator. The

equations defining these component masses can be found by Equation 3-5, Equation 3-6, and

Equation 3-7, respectively, where Ps,spacecraft is the thermal power requirement needed to satisfy

spacecraft electrical power needs in (kW), r/ is thermal to electric conversion efficiency, and

SpMassrad is the specific mass of the radiator (kg/kW).









massp, =.635*Pe


massCBC = 0.875 P (3-6)


massrad = ((,spacecraft )(1 s )) SpMaSSad (3-7)

These equations were derived from both literature data-points and textbook material.16,32

Further details on these components and their derivations is given in the NEP propulsion section,

which focuses more predominantly on the power conversion process.

Propulsion

The NTR propulsion equations were developed based on the principles of the "rocket

equation," which take into account both delta-V requirements and the mass of the spacecraft.

The ephemeris tool that was used to model the NTR architecture assumed that there was an

escape bum at Earth and the destination planet, and also a capture bum at the respective planets.

These four high-thrust burns typically last on the order of a few hours and are represented in the

equations by four separate delta-V values. The four separate equations that were used to

calculate the propellant mass at each of these bums are shown in Equation 3-8 through Equation

3-11, where the exit velocity is given by Equation 3-12, and the mass totals at the beginning of

each of the four burn segments are given by Equation 3-12 through Equation 3-16.

e Avl/v NTR
m prop NTR mTOT1 *(1 -ev ) (3-8)


mprop2NTR= TOT2 *(1-e Av2 ) (3-9)


mprop3NTR mTOT- 3 ( Av e 3 NTR ) (3-10)


mprop4NTR = mTOT4 (- e Av4 NTR) (3-11)


Ve NTR = Sp g (3-12)


(3-5)










(3-13)


MTOT1 power prop inert spacecraft propl,2,3,4


mTOT2 = power prop inert + spacecraft + mprop2,3,4 mprop tankl (3-14)


mTOT3 =m power prop inert + spacecraft + mprop3,4 prop tan kl,2 (3-15)


MTOT4 = power prop inert +m spacecraft + prop4 Mprop tankl,2,3 (3-16)


The Av terms in the propellant mass equations signify the delta-V requirements

determined by the ephemeris tools, and the power prop inert, spacecraft mprop, and


mprop tank values are masses in (kg) calculated by the NTR model.

Another important parameter in these equations that is a function of fuel parameters is the

ext velocity ve NTR It is seen from the exit velocity equation above that this parameter is a

function of ISP, increasing for fuels with higher operating temperatures, and thus decreasing the

required propellant load. Note also that the delta-V values from the ephemeris model output

were used with the assumption of no gravity losses. Therefore, any losses that were not

accounted for in the original code, were not accounted for elsewhere in the propulsion models.

The weight of the spacecraft at LEO was also an important parameter in the propulsion

model. This can be seen by the dependence of propellant load on spacecraft mass as well. All of

the spacecraft components, total propellant, and four propellant tanks were included in the first

mass estimate (nron1). At the end of each stage, however, the fuel that was spent, along with the

mass of the tank, had to be deducted from the mass estimated before the bum. The mass

represented in the equations above by the notation spacecraft comprises that of general

components including the communication and navigation systems, along with the TransHab and









ECRV. The mpower/prop inert notation refers to any components required for power or propulsion

including the reactor itself, fuel cells, and power conversion components.

Nuclear Electric Power & Propulsion

History of Electric Propulsion (EP) Systems

The history of electric propulsion (EP) dates all the way back to 1903 when the visionaries

in the scientific community began exploring its potential for spaceflight. In that year

Tsiolkovsky released his article "Investigation of Universal Space by Means of Reactive

Devices," which contained the Tsiolkovsky rocket equation, notably the most fundamental

mathematical expression in the field of space propulsion. It was only eight years later that he

published an article that mentioned the idea of electric propulsion, saying "it is possible that in

time we may use electricity to produce a large velocity for the particles ejected from a rocket

device." The combined knowledge of cathode ray tube development at that time, along with his

appreciation for the importance of rocket exhaust velocity to space propulsion led to

Tsiolkovsky's anticipation of the future of electric propulsion.33

Around the same period, Robert Goddard was also investigating electric propulsion ideas

as a natural result of his physics work on electricity and his passion for propulsion. His initial

musings on the electrostatic acceleration of electrons eventually led to thoughts on the reaction

of ions in an electrostatic accelerator, and the neutralization of the charged exhaust with a stream

of oppositely charged particles. By 1917 Goddard had developed the world's first documented

electrostatic ion accelerator, with the intention that it could be used for propulsion.33

It was Hermann Julius Oberth, however, who first declared to the technical community that

EP was a serious and worthy consideration for future astronautics. His 1929 text Ways to

Spaceflight devoted an entire chapter entitled "The Electric Spaceship" to the capabilities and

future role of EP in propulsion and attitude control. This book is also what brought EP into the









spotlight for science fiction writers. Unfortunately, after the 1930s this is where EP stayed as

scientific advancements halted and efforts eventually switched to chemical rockets that were

needed for the second World War.33

The following decades after the war did see a renewal in interest in EP research. In 1949

British physicists L.R. Shepherd and A.V. Cleaver declared that ion rocketry was impractical,

and in 1951 American astrophysicist Lyman Spitzer found that ion propulsion was perfectly

feasible. Ernst Stuhlinger wrote his classic text Ion Propulsion on the same subject in 1964. The

1960s in particular saw a growth in coordinated research and development programs addressing

EP technology, mostly due to the pervasive upswing of U.S. space ambitions. Following

experimental flight tests in the 1970s came the first commercial applications of EP in the 1980s

as attitude control thrusters on commercial spacecraft. The early 1990s saw electrothermal arcjet

use on many communication satellites, and by 1998 electrostatic ion thrusters had been used on a

planetary mission for NASA.33

Current activities in electric propulsion are being conducted by NASA and throughout

academia and include basic research all the way to flight demonstration. A cooperative

agreement between NASA and MIT has allowed for research in the area of Hall thruster

modeling to provide a theoretical understanding of the physical processes occurring in this

electrostatic thruster type. In the flight arena, NASA currently has one active electrically

propelled satellite called Earth Observing-1 (EO-1) that operates with pulsed plasma thrusters.

In addition, NASA is currently developing the DAWN asteroid science mission that will use

three NSTAR ion engines, and is providing a set of colloid thrusters in support of the ESA

Smart2 spacecraft as part of the New Millennium Program. NASA's Project Prometheus, which

was canceled due to budget constraints in 2004, aimed to combine space nuclear power with









electric propulsion technologies to allow for sophisticated active/passive remote sensing, greater

launch window flexibility, and increased science data rates. The technology areas that were

under development for the project's Jupiter Icy Moons Orbiter (JIMO) mission included high

power and high ISP gridded ion thrusters, increased thruster lifetime, high power uptake power

processing units (PPUs), and radiation-hardened components. Continued research in these areas

is essential if advanced robotic and human missions to interplanetary destinations is ever to be

accomplished.34

Fundamental Research and EP Description

Within the field of electric propulsion, three main subdivisions have arisen: electrothermal

electrostatic, and electromagnetic propulsion. In electrothermal propulsion, the propellant is

heated by an electrical process and expanded through a suitable nozzle. Three subclasses of this

division include resistojets, arcjets, and inductively and radiatively heated devices. In

electrostatic propulsion, the propellant is accelerated by direct application of electrostatic forces

to ionized particles. The most common thruster type in this class is the ion thruster, though some

development work has also been done on field emission electric propulsion (FEEP), which

produces minute amounts of thrust. In the third category of electromagnetic propulsion, the

propellant is accelerated under the combined action of electric and magnetic fields. Subclasses

in this electric propulsion category include magnetoplasmadynamic thrusters (MPDT), Hall-

current accelerators, and pulsed plasma devices.35

When considering potential EP thrusters for interplanetary, high-energy requirement

missions, it is currently only feasible to consider the electrostatic and electromagnetic classes of

propulsion. The electrothermal method falls behind the other two classes due to fundamental

thermal limitations for exhaust speeds and lifetimes that primarily result from the heating and

expansion processes required of electrothermal accelerators.35 Acceleration of the propellant by









external forces, however, will allow for higher efficiencies and specific impulse values. The

specific thruster types within the electrostatic and electromagnetic categories that were

considered in this study include ion, MPD, and Hall thrusters.

Ion thrusters

One of the electrostatic devices capable of satisfying the acceleration levels described

above is the ion thruster. It functions by accelerating a beam of atomic ions with a suitable

electric field and then neutralizing it with a flux of free electrons. Positive atomic ions are

liberated from the propellant source and accelerated by the electrostatic field created with

strategically placed magnets. These atoms are then combined with the electron source outside of

the grid to produce a net zero charge stream with speed determined by the net potential drop over

the distance between atom release and the neutralization plate.35

Although the ion engine is one of the most complex within the group of EP engines, they

also have undergone the most research and development over the years. They are appealing for

space travel due to past demonstration of long lifetime, high efficiency, and high specific

impulses. For the purposes of this study, the lifetime was assumed to be 20000 hours, the ISP

was estimated at 6000 (s), the thruster efficiency at 90%, and the thrust per thruster at 1 (N) with

xenon as the propellant choice.26'35 Disadvantages with this thruster include the indicated low

thrust density, high system complexity, and high PPU specific masses. This technology has been

demonstrated on over a dozen U.S. and Soviet flight tests dating back to 1962. The first

operational use of ion engines was in 1994 for the Japanese ETS-6 and COMETS satellites, and

was followed by use on a commercial satellite bus in 1997 and the interplanetary Deep Space

NASA mission in 1998.35 Future research on these systems will most likely focus on developing

higher power capabilities, decreasing the mass of high power PPUs, and developing extremely

small propellant feed systems.36









MPD thrusters

Despite having considered only one electrostatic thruster in this study, both the MPD and

Hall thruster engines were considered from the electromagnetic group. This follows

appropriately from the fact that electromagnetic acceleration presents many possibilities for

implementation given that the applied fields and internal currents may be steady, pulsed, or

alternating over a range of frequencies. There are also a variety of propellant types, electrode

configurations, and means for injection and ionization of atoms that are available when using

these electromagnetic thrusters.35

An MPD thruster is configured with a central coaxial geometry cathode, an annular anode,

and an interelectrode insulator. In this thruster type, gaseous propellant flows into the upstream

part of the channel whereby the atoms are ionized with a uniform electric arc, compressed into a

hot plasma just beyond the cathode tip, and expanded out the thruster as plasma exhaust. These

thrusters are capable of high specific impulses, moderate efficiencies, and high thrust. The

parameters assumed for this study include a lifetime of 2400 (h), an ISP of 5000 (s), an

efficiency of 65%, and a thrust of 100 (N).9'26'35'37 It is to be noted however, that megawatt

power levels and the subsequent propellant flow rates for MPD thrusters have not yet been tested

and are both technological and economic problems.35 A major objective for MPDT research is

the achievement of thrust efficiencies greater than 50% with non-condensable propellants while

operating the thrusters below the plasma instability threshold.35

Hall thrusters

Electromagnetic Hall thrusters rely on use of the same-named "Hall effect," which refers to

the potential difference on opposite sides of a thin sheet of conduction material through which

electric current created by a magnetic field applied perpendicular to the Hall element is flowing.

In this type of thruster, propellant is accelerated by an electric field in a plasma discharge with a









radial magnetic field. Channel and field geometries lock the plasma electrons into nearly

collisionless cross-stream drifts, leaving the positive ions to accelerate downstream under the

applied electric field.35 This thruster device is sometimes thought of as a cross between an

electrostatic and electromagnetic accelerator.

Although Hall thrusters have a relatively high specific impulse and thruster efficiency, they

have a low thrust output per thruster, and past testing has shown life-limiting erosion and

unstable and oscillatory plasma discharges to be of concern.34 The assumed performance

parameters used in this study include a lifetime of 8000 hours, a specific impulse of 400 (s), a

thruster efficiency of 62%, and a thrust of 0.4 (N).38 Research interests for Hall thrusters include

using krypton as propellant, increasing the lifetime to greater than 8000 hours, and eliminating

the erosion in the plasma discharge chamber.34

Despite knowledge of current performance parameters including ISP, efficiency, and thrust

for each of the EP thruster types under consideration, the type of missions that the thrusters are

being applied to requires some assumptions on the future state of the technologies. Advanced

interplanetary missions will have very high delta-V and burn time requirements that may be

feasible for these EP systems on paper, but have not actually been tested experimentally. As

alluded to earlier, testing of these thrusters at high powers and for long-life durations will, in

itself, be a demanding task. However, as the missions being studied have departure dates on the

order of 2030 for Mars and 2040 for Saturn and Jupiter, a "leap of faith" will have to be taken by

assuming that the technologies will be available to satisfy the mission requirements. The

assumptions made in this study also included the thrusters being able to operate at very high

power with PPUs that can process power coming from megawatt-level nuclear reactors, and

thrusters that can burn for considerably long times. It is suspect that these capabilities will









improve over the next 20-40 years and the performance parameters will most likely increase

along with them.

Development of the Nuclear Electric Propulsion (NEP) Model

The NEP architecture was broken-down into three systems based on the type of NEP

thruster being used. This was different than the breakdown for NTR systems, which varied by

reactor fuel type. Since the propulsive performance of NEP systems was primarily dependent on

thruster type and this performance can vary greatly between different thruster options, this was

the primary focus. The reactor needed for an NEP system was unimodal in the sense that it only

needed to generate power (in contrast to the usual connotation of 'unimodal' as only producing

propulsive heat). It was thus not extremely important what type of reactor was used since the

masses of all reactor types were only slightly different and the reactor mass was merely a small

fraction of the total mass of the CTV.

The thruster types chosen for analysis in this trade study included MPD thrusters, ion

thrusters, and Hall thrusters. Each of these thruster types would provide different ISP,

efficiency, and overall performance capabilities to an NEP spacecraft, thus it was paramount to

properly explore the breadth of NEP systems by assessing each of these thruster options.

The schematic shown in Figure 3-3 depicts the spacecraft components that together make-

up the NEP vehicle architecture. It is apparent in this diagram that the reactors served only as

power generation for the spacecraft subsystems and electric propulsion mechanism. While the

blue components are part of the power generation subsystem and the yellow components

comprise the set of common vehicle components, the pink components now represent the electric

propulsion subsystem. In an NEP system, electric thrusters provide the sole means of propulsion

and require both liquid propellant and electric power for functioning. The arrows in the









schematic are divided between the solid black arrows that depict a coolant or propellant flow,

and the dashed black lines that indicate flow of power to each system.

Power

The considerations for determining the reactor mass for the NEP power reactor were

much different than those for sizing the different NTR reactors. Since space reactor fuels

provided the trade within the NTR architecture, the different mass sizes of each of the NTR

reactors was pivotal to the results since mass was based upon fuel choice. In addition, the NTR

reactors provided the propulsive energy for the spacecraft, making fuel parameters such as

temperature and ISP important variables in an overall comparison of NTR systems.

In the case of the NEP power reactor, the only parameter that went into the NEP

architecture model based on the reactor was that of the reactor mass. In addition, only one

reactor was needed for all of the NEP systems, since the NEP trade was essentially on the type of

electric thruster being used. The propulsive properties of the EP thruster were what

differentiated these systems, and thus the mass of the reactor simply became another term to

lump into the total mass carried along with the spacecraft.

It was thus decided to label the NEP power reactor as a fast reactor, without specification

to its specific fuel form. This generalism was used so that various data-points from literature

sources could be used to develop an equation to be used in calculating reactor mass based on the

power requirements calculated by the NEP model. Three NEP power reactor descriptions were

taken from previous NEP mission analysis studies, and their main descriptive parameters were

compiled in Table 3-4.32,37,39

Although each of the three reactors from these studies had slightly different fuel forms,

they were all fast reactors, as expected for an NEP system. As discussed previously, fast reactors

are ideal for NEP systems since they can provide long-duration steady power. The mass and









power data-points corresponding to these three reactors included the total mass of the power

reactor, pressure vessel, and internal and external shielding. By plotting each of the points on the

curve seen in Figure 3-4, a trendline was used to calculate Equation 3-17, where the mass in (kg)

is dependent on the electric power requirements Pe in (MWe) of the CTV.

Massreactor EP = 53.939* Pe2 + 976.12 Pe + 2814.9 (3-17)

The reactor masses were thus calculated separately for each mission since the calculated

power requirements varied for each mission. It should also be noted that the ISP was assumed

for the NEP reactor to be 925 (s), since this is the same ISP as the CERMET fast reactor used in

the NTR model.

In order to better understand the flow of power and the associated losses from the reactor

to the thruster, a pictorial representation of this flow, along with parameter descriptions, is found

in Figure 3-5. Because the Brayton power conversion unit was indiscriminate of the final

destination of its electric power, it provided the function of converting all thermal power

generated by the reactor to electric power for spacecraft needs. The PMAD system, which

consisted of cabling, fault protection, and switching gear, then directed the power to the

appropriate spacecraft loads.16 Losses were found in the Brayton Conversion Unit, PPU, and

electric thrusters, as shown and described in the power and efficiency diagram. Some of the

major equations used to describe the power and efficiency calculations visualized in the flow

diagram are given by the thermal power Equation 3-18, and the electric power Equation 3-19,

where the power for propulsive needs is given by Equation 3-20, and the overall system

efficiency is given by Equation 3-21


s= Ps,prop +s,spacecraft (318)


Pe = P' *7s = Pe,prop + e,spacecraf (3-19)









Pe, = Pp + P (3-20)
e, prop pp th


rIT = s 7pp 7th (3-21)

Propulsion

The size of the propulsion system was dependent on two main factors: the energy

requirements for the mission, and the remaining spacecraft vehicle mass. The energy

requirements for the NEP spacecraft consisted of a delta-V value for the outbound and inbound

transfers and the bum time of the thrusters during these transfers. It is important to note that this

was the only architecture that required spiral trajectories out of the Earth and destination planets'

atmospheres. This was necessary due to the need for the spacecraft to pick up speed in order to

escape from the atmosphere of each respective planet. Although the propellant was sized

according to the total requirements of both the transfer and spiral durations, the crew would only

be on the CTV for the transfer duration as it was assumed that they were transferred to the

spacecraft after it picked up the necessary speed for escape. This was important so that the

requirements for transfer time given in the mission profiles would not have to include the rather

long spiral times required by the NEP system. This drawback to the NEP system was not

neglected as it was accounted for in the final tradespace analysis.

In addition to energy requirements, the sizing of the system propellant was also dependent

on the weight of all the components it had to propel through space. This weight included the

communication and navigation components, TransHab and ECRV, truss structure, power reactor,

Brayton Conversion Unit, radiators, and the propulsive thrusters and PPUs. More simply put,

enough propellant had to beonboard to carry the rest of the spacecraft weight to its destination. It

must not be forgotten that the weight of the unused propellant itself also played into the weight

equation since it too was carried onboard the spacecraft until it was used.









The two main mass estimates needed to size the propellant mass for outbound and return

journeys were the IMLEO (m,,,o) and the initial mass at departure from the destination planet

(moT2). The propellant loads required to carry these masses were divided into two separate

equations (Equation 3-22 and Equation 3-23) since the outbound and return burns are two

distinct entities.

mpropl = mOT1 *(-e-Aviv) (3-22)


mprop2 TOT2 (1_ 2v ) (3-23)

This allowed for the second propellant load equation to be calculated assuming a smaller

mass due to the first propellant load being burned off, and from dropping its associated

propellant tank. The initial mass quantities at initiation of the two main burns are given in

Equation 3-24 and Equation 3-25, where it can be seen that the propellant and total mass

equations are interdependent.

mTOTI = NEP inert + Mpropl,2 (3-24)

mTOT2 = mNEP inert -m roptankl +mprop2 (3-25)


The term mNEp inert used in the mass equations represents the total spacecraft mass minus

the propellant, an equation for which is given by Equation 3-26, where the terms represent the

mass of the power reactor, propulsion hardware, communication and navigation components, and

TransHab, ECRV and structural truss masses, respectively.

mNEP inert power + mpp+th + mcomm,nav Transhab,ECRV,truss 1(3-26)

As the mass sizing of the power reactor and its associated pressure vessel, shield, and feed

system have been previously discussed, and sizing of the spacecraft communication and









navigation systems is to follow, only the sizing of the propulsion system will be discussed

herein.

The mass estimates for the PPU and thruster hardware were taken from both text and

literature sources that had specific mass (kg/kWe) values for ion, MPD, and Hall thrusters. The

MPD thruster and PPU combined specific mass was the lowest of the three, at 0.6 (kg/kWe).

This parameter was mostly due to the very high power levels processed in this simple and robust

thruster device, and has made MPD thrusters highly desirable for interplanetary missions.26'35

The specific mass estimate for the Hall thruster system was based off of both an advanced

thruster and advanced PPU design. The High Voltage Hall Accelerator (HIVHAC)

Development Program predicted a possible future thruster specific mass of 1.3 (kg/kWe), while

PPU specific masses ranged from a progressive 5 (kg/kWe) to a very advanced 2

(kg/kWe).37'38'40 Given the distant departure times of the reference mission profiles under

consideration in this study, the advanced PPU specific mass was assumed, for a total PPU and

thruster specific mass of 3.3 (kg/kWe). Although ion thrusters are known to require very heavy

PPUs (-10 (kg/kWe)), a similar technology advancement was assumed for the ion system and a

value of 4 (kg/kWe) was used for the PPU specific mass.35 Combined with a literature data-

point of 1.2 (kg/kWe) for the ion thruster itself, a total specific mass of 5.2 (kg/kWe) was

assumed for the thruster and PPU combination.37 The equations for the combined PPU and

thruster mass for the MPD system is thus given by Equation 3-27, the Hall system by Equation

3-28, and the ion system by Equation 3-29, where the term Pe,prop is the electric power supplied

to the EP thrusters.

masspp+ = 0.6 Pe,pro (1 + burn duration /lifetime) (3-27)

masspp+ = 3.3 ,po (1 + burn _duration / lifetime) (3-28)









masspth = 5.2 Pe,pro (1+ burn _duration/lifetime) (3-29)

Although these specific mass estimates alone provided the sizing for the thrusters based

upon satisfying the mission energy requirements, they do not account for the functional lifetimes

of the thrusters that may limit the duration of their use. Thus contingency weight had to be

added for thruster hardware for missions that required burns greater than the lifetime of the

thruster selected. The burn durations for the NEP system were defined as the spiral out/in times,

along with burn times between planets. This duration is not the same as the crewed mission

duration, which would include the transfer times and stay time on the destination planet. Given

this assumption, thrust correction maneuvers needed during crew stay at the destination planet

were assumed supplied via alternate methods such as reaction control systems.

The mass estimates for the xenon and liquid hydrogen propellant tanks differed based on

the propellant, not the thruster, and were approximated with textbook equations.16 The tank

estimates for the ion and Hall thrusters thus used the same equation since both required xenon

propellant, while the MPD thruster used a tank equation based upon hydrogen propellant.26 The

differences in the densities of the propellants led to a different required volume per unit mass,

and thus different propellant tank volumes and masses. Textbook sources gave Equation 3-30

for MPD tank mass, while the equation for the xenon propelled ion and Hall thrusters was given

by Equation 3-31.

masstak H2 = .287* massprop (3-30)


masstank e = 52 + 0.075 masspp +.154 mass 2/3 (3-31)

For both equations, the term massprop represented the total amount of propellant used for

the mission.16,32









The power conversion, power management, and heat dissipation system masses were not

dependent on the thruster used, thus common equations could be used for all thruster models.

Due to the lack of mass approximations for the PMAD component in textbook material, three

data-points from the literature were used to approximate the PMAD mass equation. Microsoft

Excel was used to derive a logarithmic approximation to the data-points' trend line, a plot of

which can be seen in Figure 3-6. Issues arose for power requirements less than about 1000 kWe

due to the nature of the logarithmic function, thus a regressed linear function was used for lower

power requirements. The formula for PMAD mass is given in Equation 3-32, which shows the

calculation based on the electric power requirement P, in (kWe).

massPAD = 7772.7 Ln(P) + 1067.5(forP, > 1000);.635 P (forP, < 1000) (3-32)

In the calculation for the Brayton Unit, a specific mass of 0.875 was used and was taken

from a literature data-point for a similar NEP crewed vehicle.32 The Brayton Unit, or CBC, mass

is given by Equation 3-33.

massCB = 0.875 (3-33)

The mass of the radiators was determined by a straightforward equation depicting the

radiator mass needed to rid the system of excess heat due to losses. The value for the specific

mass of the radiators was estimated to be 1.5, which is slightly more conservative than a text

estimate of 0.1-0.4 kg/kW, and in close approximation to the Brayton to radiator mass ratio from

the previously mentioned literature source.16,32 The equation used for the radiator mass estimate

is seen in Equation 3-34, where SpMassmrd is the specific mass value of 1.5 for the radiator.


massrad = P,prop p *( *- ,s)+(P Ps,prop)( s)) SpMassad (3-34)









An additional equation that was used to estimate the mass for the truss structure for the

entire spacecraft was Equation 3-35, where again, mass,,er, refers to the entire spacecraft mass

not including propellant.

masstrus = 10% massznert (3-35)

In addition to sizing the NEP system components, the propulsion equations also derived

the required thermal and electric power for the spacecraft, along with the thrust required by the

EP thrusters. One important point to note is that the thermal power generated by the reactor was

converted to electric power and then split between the propulsion system that operates the

thrusters, and the components requiring constant power. These components included

communication and navigation systems along with the TransHab module. This effect is shown

through the thermal power equation, Equation 3-36.


mprop (ISp g)
Ps = /2 'rq + (Ps,COMM,NA V,Transhab) (3-36)


/ prop

The total electric power requirements for these two separate entities combined is given by

Equation 3-37.

Pe =Ps *'s (3-37)

The thrust variable T is then based only on the thermal power devoted to the propulsion

system, Ps,prop, as seen by Equation 3-38.

2* 77T *P (3-3
T T 7Isp* g (3-38)









Knowing the value of the thrust is only significant for calculating the number of thrusters

that would be needed for a particular EP system given by the simple relation in Equation 3-39,

where the term in the denominator indicates the characteristic thrust per EP thruster.

T
# thrusters = (3-39)
T thruster

It should be clear from inspection of the analysis performed to develop the NEP propulsion

equations, that thruster-specific parameters played a major part in determining spacecraft

performance. The EP thruster efficiency, specific impulse, and thrust per thruster had a large

impact on power and thrust requirements, overall system efficiencies, and wet and dry mass

estimates for the spacecraft. These thruster characteristic parameters were even used as inputs in

the ephemeris models, which determined the energy requirements for the system. The specific

mass, fuel, and lifetime parameters for each thruster were the main drivers for the mass estimates

of the actual thruster and PPU hardware, along with the propellant. The parameters used for the

ion, MPD, and Hall thrusters in this study are briefly summarized again in Table 3-5.

Hybrid Power and Propulsion

Background

Up until this point, Hybrid vehicle architectures have only been minimally analyzed

through paper studies. One such study indicates that an integrated nuclear thermal and nuclear

electric propulsion system would grant the ability to obtain from two systems much more than

one can obtain from a single technology.6 A NASA presentation on 'Bimodal Nuclear Thermal

Rocket Propulsion for Future Human Mars Exploration Missions' gives pictures, schematics, and

explanation of how a bimodal nuclear thermal reactor could apply electrical power production to

a secondary electric propulsion system. The basic methodology is that the bimodal nuclear

thermal rocket (BNTR) would produce short bursts of high thrust for escape mechanisms,









followed by a long power generation phase wherein the electric thrusters would be operational.41

It is the hope of mission planners that a Hybrid spacecraft would combine the benefits of both

NTR and NEP to increase overall performance for challenging interplanetary missions.

Development of the Hybrid Model

The Hybrid vehicle architecture schematic was devised as a cross between the NTR

bimodal system and the NEP system. As can be seen in Figure 3-7, the schematic for the Hybrid

system maintained the power and propulsion subsystems seen previously in the NTR bimodal

schematic, along with the electric propulsion subsystem seen in the NEP schematic. Dashed

lines in this schematic refer to power generation flow, while the solid black arrows refer to

propellant or coolant flow.

The unique vehicle design in the Hybrid schematic allowed for the reactor to provide

power for the spacecraft needs and the necessary thermal energy and electric power for two

modes of propulsion. The location of the two propulsion systems relative to the TransHab was

indeed intentional, as the reactor system would most likely reside at the opposite end of the

TransHab, while the EP system would be on the same end. The two thrust mechanisms would

also therefore produce thrust in opposite directions relative to the spacecraft itself, requiring

realignment of the vehicle before changing propulsive mode.

Power

The same bimodall' NTR reactor that was used for thermal power generation for the NTR

bimodal architecture was also used for the three Hybrid architectures. A bimodal engine was

required for this architecture in order to provide thermal propulsion for the NTR engines and

electric power for the EP thrusters and other spacecraft functions. The bimodal reactor utilized a

fast fission spectrum and CERMET reactor fuel. An assumption of two reactors onboard the

spacecraft was used just as for the other architectures. In order to estimate the total reactor mass,









a dependence on the electric power requirement was used just as it was for the NEP system.

This estimate was found by using the same equation for reactor mass as the NEP system, but

adding an additional weight to account for NTR reactor fuel and propulsive hardware. The

bimodal engine at 5272.3 (kg) was designed to have an electric power output of 25 (kWe). An

NEP reactor operating at 25 (kWe) would have a calculated mass of 2839.3 (kg). This difference

between the two of 2432.9 (kg) was thus assumed to be the "additional" NTR weight added to

the calculated weight of the reactor based only on the NEP reactor mass. The NEP reactor mass

equation for one reactor is shown below in Equation 3-40, and was updated to account for the

additional mass needed for the Hybrid reactor.

MaSSreactor Hybrid = (53.939 Pe2 + 976.12 Pe + 2814.9) + 2432.9 (3-40)

Propulsion

The equations that defined the EP thruster sizing, along with the parameters for the EP

thrusters, were the same as those used for the hybrid architecture. The EP aspect of the Hybrid

vehicle was identical to the NEP architecture with all three types of EP thrusters assessed.

Although the vehicle had an NTR engine, all configurations utilized the bimodal CERMET

reactor, thus the trade was solely on EP thruster type, not on NTR fuel type as it was for the NTR

architecture.

The equations that changed compared to the NEP system were those for the propellant

sizing. The Hybrid propulsive equations represented operationally both NTR and NEP delta-V

requirements, in which one NTR delta-V bum occurred for planetary escape and was then

followed by a long-duration NEP delta-V burn that allowed for capture. The first three of the

four tanks associated with the four delta-V maneuvers was dropped after its respective bum

segment. For instance, the mass of the propellant required for the second bum used the IMLEO









mass minus the propellant burned in the first NTR segment along with the tank that housed that

propellant. These tanks were non-recoverable and inevitably became space 'junk.' The

equations that defined the propulsive mass requirements for the Hybrid architecture's four burns

are given by Equation 3-41 through Equation 3-44, where the mass equations at initiation of each

burn are given by Equation 3-45 through Equation 3-48.


mproplNTR = mTT *(1 e Rv,_ ) (3-41)


mprop2NEP = TOT2 (1 Av2 P ) (3-42)


mprop3NTR = m O3 *(1 -eAv3 R ) (3-43)


mprop4NEP mTOT4 *(l- 4e v ) (3-44)


mTOT = mNTR inert +mNEP inert + mpropl,2,3,4 (3-45)


mTOT2 NTR inert + NEP inert + prop2,3,4 prop tankl (3-46)


mTOT3 =mNTR inert + mNEP inert + prop3,4 mprop tankl,2 (3-47)


mTOT4 mNTR inert + mNEP inert + mprop4 mprop tankl,2,3 (3-48)

Note that the propellant mass subscripts indicating NEP or NTR burns are not included in

the second set of total mass equations, but are instead merely indicated by the number, or order,

in which they occur.

Common Vehicle Subsystems

TransHab

Every vehicle system that was analyzed for the advanced interplanetary missions in this

study had a crew living module called the 'TransHab'. This term, commonly found in mission









analysis literature, was coined for the actual structure that was still in the design phase when

research efforts ended in 2001. The TransHab was the first space inflatable module designed by

NASA, conceived as a technology capable of supporting a crew of six on an extended space

journey. The space habitat had an endoskeletal design, with a light and reconfigurable central

structure and deployable pressure shell. The shell was elaborate in that it was made of several

layers, could contain water or hydrogen for radiation shielding, and had an outer layer a foot

thick. Recommendations drawn from the research completed during this NASA program

included having a three-level internal layout with isolated crew quarters at the center, and

exercise and hygiene stations situated on different levels from the public functions in the kitchen,

dining and conference areas. Unfortunately budget cuts along with reported values of $100

million to build the structure led to the project's demise.41 After the four-year research period

between 1997 and 2001, Bigelow Aerospace purchased the rights to the patents developed by

NASA and is currently continuing development of a similar structure for future commercial

space applications.1

The ECRV was an additional habitation environment designed to be attached externally to

the TransHab for a majority of the mission duration. In a hypothetical mission, when the CTV is

being assembled in LEO prior to flight, the crew will be transferred from the launch vehicle to

the CTV via the ECRV. The TransHab will then be inflated and flooring and partitions deployed

so the crew may move from the ECRV into the TransHab for the remainder of the trip. The

ECRV will also likely be the capsule the crew will use to aerobrake into the Earth's atmosphere

after the interplanetary voyage. This would follow the propulsive capture of the entire CTV back

into LEO and disengagement of the ECRV from the main space vehicle.









A vital consideration for the future design and sizing of the TransHab will be radiation

shielding. It has been noted by several literature sources that the hazards of space radiation

create the primary restraint on crewed interplanetary space missions. This is a problem that has

not yet been resolved by nature due to the many complexities space radiation presents. Future

crewmembers will be exposed to many types of radiation (ionizing and non-ionizing), which

fluctuate considerably over time and location. This implies that different crew procedures and

levels of protection will be needed over the course of a mission. More research will have to be

done on the shielding properties used for protection, as human-rated structures must be

approached with a very radical design perspective.42 For example, some studies have proposed

surrounding the crew module with hydrogen propellant tanks for radiation protection. Although

the hydrogen-containing materials would be good radiation absorbers, the aluminum tank

structure would actually produce dangerous secondary radiation particles. There is also

uncertainty in the biological response to high-charge radiation. Currently, no exposure

limitations have been set for deep space missions, and recommendations simply rely on exposure

limits for operations in LEO.43

Some work has been done in the past to try and model how much and what type of

radiation could be expected on interplanetary missions. The two main forms of ionizing

radiation that are estimated include the constant, but low intensity galactic cosmic radiation

(GCR), and the low frequency, but high intensity solar particle events (SPE). One study reported

that with 1280 (kg) of polyethylene TransHab shielding for a Mars journey, 13.38 (cSv) of

radiation dose equivalent was detected for SPE and 17.65 (cSv) for GCR for each one-way 180 -

day transfer. The surface habitation on Mars resulted in 28.05 (cSv) for GCR and 2.23 (cSv) for

SPE.44 The lower rate of exposure on Mars was a result of the natural radiation shielding from









the Martian atmosphere. Although this study found a total radiation dose equivalent rate under

the NASA LEO limit of 50 (cSv/year), the rate of exposure during the transfer period would

indeed have been over the limit if it had been received for an entire year. A long duration

transfer would have led to 63.5 (cSv/year) exposure, which is a significant violation of the

current limit.

Although radiation shielding would be a significant factor in the design and development

of the TransHab, it was deemed too difficult to account for shielding mass for all of the reference

missions run in this study. It was assumed therefore that the 1900 (kg) that was assumed for the

TransHab shielding weight, in addition to other materials such as wastewater and propellant

tanks, would provide the necessary shielding for the trip. The breakdown of the TransHab-

related masses thus became quite simple and was modeled from previous mission analysis

sources. The module itself had a mass of 21000 (kg), the 6 crewmembers and their belongings

equated to 600 (kg), and the ECRV was set at 5100 (kg). The consumables were dependent on

the length of the mission (including stay time at planet), and were sized at 2.45 (kg) per person

per day for food and water.18 The breakdown of masses for the TransHab components can be

found in Table 3-6.

The power requirement for the TransHab was also an important parameter to consider

when generating the total operating power requirements of the spacecraft. The actual value of

constant power that must be provided for the crew module so that everyday living and working

needs can be satisfied was a difficult value to estimate, especially given the scarcity of

information in the literature. In the lecture notes of G.L. Kulcinsky, however, a value of 100

(kWe) was given for a "manned outpost" on the surface of Mars.45 It was believed that this









would be a reasonable estimate for the TranHab since it would require similar power needs to a

planetary ground station.

Spacecraft Functional Components

The remaining spacecraft components that are common to each of the space architectures

were those that provided for various communication and navigation functions to get the

spacecraft to its destination and back. These system components were broken down into the

following groups: command and data handling (C&DH), tracking, telemetry and command

(TT&C), attitude determination and control system (ADCS), the reaction control system (RCS),

and truss structure used to link all spacecraft hardware. The C&DH subsystem distributed

commands and accumulated, stored, and formatted data from the spacecraft. The TT&C

subsystem linked spacecraft with ground or other spacecraft by receiving commands and sending

status telemetry. The ADCS provided stability and orientation in the desired directions during a

mission despite external disturbances. The propulsive RCS subcomponent was linked to the

ADCS by performing the thrust actions that were needed by the spacecraft as determined by the

ADCS.

Although the general functions and equipment needed by these subcomponents was

important to understand from the systems engineering perspective, the two most important

parameters in the spacecraft model development were the mass and power estimates for the

components. It was again difficult to estimate the size and power of these subcomponents given

the limited number of papers found in the literature that went to similar depth-levels for

spacecraft sizing, yet one paper provided values for a reasonable estimation. For the purposes of

this study, the ADCS was estimated at 70 (kg), the C&DH at 35 (kg), and the TT&C at 205 (kg).

The total power required was taken from the same reference, which designated the

"communication loads" as requiring a power of 5 (kWe).37 The RCS mass was taken from two









data-points, one from an NTR study, and the other from an NEP study. The NTR study showed

an RCS mass that was 5.7% times that of the propulsion propellant mass, the NEP study gave a

value of 6.9%, and one textbook suggested a range from 3-10%.16,18,37 It was thus decided to set

the RCS mass to 6% of the propellant mass required by the main propulsion systems for a given

architecture. Since the RCS would only require a minimal amount of power at various small

time increments over the course of the duration of the flight, it was decided to neglect its power

requirement when sizing the overall power of the nuclear reactors. The structure of each

spacecraft was based on the total mass of the system, scaling with system mass. A common

estimate which was used to size the truss structure mass was 10% of the overall spacecraft

mass.16 A summary of the mass breakdown between common subcomponents is given in Table

3-7.

Analysis Tools

ModelCenter Software

The main analysis tool used in this study was Phoenix Integration, Inc.'s ModelCenter.

This tool created a visual environment for process integration and employed complex design

exploration techniques, making it ideal for performing trade studies and exploring a design

space.46 One major benefit to this tool was that it reduced both the time to run individual

programs and the possibility for human errors that lead to costly mistakes. In addition,

ModelCenter offered powerful and timesaving design tools that included parametric plots,

optimization tools and linking of models. Each of these assets together provided the ability to

conduct a large trade study within a reasonable time period.

The Analysis Server, a product by the same company, was used in conjunction with

ModelCenter in order to be able to convert codes being used into reusable components that were

directly accessible within ModelCenter. The Analysis Server was a flexible, Java-based software









server that allowed applications to be wrapped for integrated engineering processes. The

'wrapped' components, which are termed FileWrappers, are built from the FileWrapper utility,

which was designed to automate the execution of analyses based on file I/O programs. The

FileWrapper components thus had the ability to automatically edit the appropriate input files, run

the code's executable, and parse the output file whenever it is executed. A template file, which

is a copy of the input file in standard format, had to be created for use with a FileWrapper

component, as the FileWrapper loads the template file and substitutes any user-inputs into it.46

Incorporation of Models into ModelCenter Framework

ModelCenter was used in this study by wrapping individual Excel worksheets and

ephemeris codes into ModelCenter system models that allowed for user inputs, automatic

passing of data from one worksheet to another, optimization and converging of parameters, and

finally parametric analyses.

Ephemeris models

The IPREP and CHEBYTOP ephemeris codes were used to generate the NTR and the

Hybrid and NEP ephemeris modeling tools, respectively. This was accomplished through the

use of ModelCenter in conjunction with Analysis Server. The IPREP and CHEBYTOP

executables were first input into the models by generating FileWrappers and input templates that

the Analysis Server used to connect with the ModelCenter files. An example of the FileWrapper

and templates created for the three propulsion system-specific ephemeris models is given by the

CHEBYTOP Hybrid model files found in Appendix C: ModelCenter Template and Filewrapper.

Creating these FileWrappers allowed the execution of IPREP and CHEBYTOP to be done using

components within their respective ephemeris models, with input and output variables appearing

in easy-to-use lists within the ModelCenter GUI.









The IPREP component was able to run an entire two-way mission in one sweep, but the

CHEBYTOP component had to be run twice within the NEP and Hybrid models to simulate the

outbound and return transfers, and was thus input as two separate components in each of the

models. The data for the apogee and perigee variables for each destination was held within an

Excel worksheet in each of the three models so that selection of planetary destination would

allow the correct orbit parameters to be passed to the IPREP or CHEBYTOP components. A

tool to aid in converting a date input by the user to a Julian date required by the executable was

also incorporated as an Excel component.

For the two models that used the CHEBYTOP components, an additional Excel component

was added that uploaded the thruster-related parameter values upon user selection of a thruster

type. These parameters included the assumed values for thruster efficiency, ISP, and thrust force

per thruster for the selection of Hall, ion, or MPD thruster type. Once all of these components

were incorporated into the model, input variables that included departure date, stay time and

transfer time were incorporated into the user interface and then linked to the appropriate

variables in the model components. Key output parameters that were desired from the models

included the delta-Vs, electric thruster bur times, and NEP spiral times. A screenshot of the

ModelCenter NEP ephemeris model's graphical user interface (GUI) is given in Figure 3-8.

ModelCenter's Darwin optimization tool was also incorporated into each of the three

ephemeris models. This genetic algorithm-based tool with fully featured GUI was integrated as

a trade study plug-in within ModelCenter. The genetic algorithm method is based on

probabilistic rules with evolution achieved through application of probabilistic operators that

mimic biological genetics.46 The goal that was achieved by using this optimizer within the

ephemeris model was to minimize the overall mission delta-V for a given mission profile. Since









each mission profile was designed with ranges on the parameters for departure date, stay time,

and transfer time, these variables became the design space for the optimization. The lower and

upper bounds for these parameters became inputs that could be manually set by the user

depending on the mission profile under review. Execution of this tool enabled the energy

requirements to be minimized based on manipulating each of the four variables in the design

space to find the very best answer given the specified mission profile requirements. Attempting

to perform such an optimization manually would have undoubtedly required many hours of

running the ephemeris model trying to make improvements one variable at a time. A screenshot

of the Darwin Optimization tool GUI is shown in Figure 3-9.

Vehicle architecture models

Incorporating the vehicle architecture models into ModelCenter was much more

straightforward than the ephemeris models since only Excel models that were fully self-

generated were used. This involved using the Excel component tool to insert the various Excel

models needed for each architecture, and then using the linking tool to link parameters within the

model. The NEP and Hybrid architectures both had two main Excel worksheets incorporated

into their ModelCenter models. The first was the main propulsion and power model that sized

the reactor, while the second was a thruster selector tool. This latter worksheet allowed the user

to select the desired thruster type from the ModelCenter GUI, automatically updating the

appropriate propulsive parameters characteristic to that thruster in the sizing model.

The NTR model was slightly different in that its secondary Excel worksheet was a reactor

fuel selector, which would input the fuel parameters for the fuel selected into the main NTR

propulsion and power model. A screenshot of the NEP model showing the input parameter and

output parameter lists, along with a block component input and output diagram is shown in

Figure 3-10.









Mission architecture models

The term 'mission architecture' was used in this study to refer to a vehicle architecture

performing a mission profile. From the modeling perspective this would equate to combining

the architecture's ephemeris model and its power and propulsion model. The mission

architectures were then developed into their respective ModelCenter models, and were created

by combining both the ephemeris and main architecture model. This allowed a user to then run

an entire mission, deriving the IMLEO and other output parameters by simply inputting the

mission profile scenario and fuel or thruster type.

Some difficulties were encountered, however, when trying to incorporate certain levels of

automation into the models, which would have been timesaving mechanisms for the user. For

instance, the ephemeris model's Darwin optimizer tool that manipulated departure date, transfer

and stay time in order to minimize delta-V energy requirements would not function properly

within the larger mission architecture ModelCenter model. This was due to the fact that the term

it was trying to optimize was being passed to the power and propulsion model, and was no longer

an output parameter of the ephemeris model as required. This finding affected the planned

methodology for the data analysis process. It was decided to maintain use of the ephemeris

ModelCenter models for optimization purposes, record all optimized departure dates, transfer

and stay times in a database, and manually enter these parameters into the mission architecture

models to run the full-scale models. When the mission architecture models were run, the

optimized delta-V energy requirements were automatically generated in the ephemeris model

and passed to the power and propulsion model. Figure 3-11 provides a visual of a mission

architecture model GUI that shows these user inputs, the two models incorporated into the larger

model, and the desired output parameters.









Validation of models

Following completion of the mission architecture models it was decided to attempt to

validate the models by comparing the IMLEO results from alternate sources to the results of the

mission architecture models developed in this study. One literature source for each of the three

main architectures was used for comparison. Attempts were made in each case to try and model

a very similar mission to that presented in each study, though lack of specific mission analysis

data such as the energy requirements used for sizing was recognized as a weakness in the

validation methodology.

The NTR architecture model was validated using the results from a very detailed Pratt &

Whitney Rocketdyne NTR model. This model was built upon numerous smaller models, high-

level analysis and very thorough attention to reactor design and development details.47 All of the

fuels analyzed in the NTR model studied in this paper were also analyzed by the P&W model,

thus it was decided to compare the IMLEO results for both models at the 15 (klbf) thrust setting.

One of the main differences between the models, which may have some effect on the validation

results, is that the P&W model assumes 1, 2 or 3 engine configurations wherein each engine

provides the indicated thrust. The NTR architecture model developed in this study assumed only

one engine was providing the thrust, yet an extra engine's weight was included in the model for

redundancy purposes.

The mission that was analyzed for the NTR validation was a round-trip Mars mission with

these specific delta-V energy requirements:

* TMI (Trans-Mars Injection)=3.60 (km/s)

* MOC (Mars Orbit Capture)=1.44 (km/s)

* TEI (Trans-Earth Injection)=1.09 (km/s)

* EOC (Earth Orbit Capture)=0 (km/s)









By running the NTR architecture model with the above energy requirements for each of the

fuel selections (graphite, composite, carbide, and CERMET), a comparison of IMLEO estimates

was made with the P&W model. The results from both models, along with the percent error

between the two sets of results, can be found in Table 3-8. The percent error data-points ranged

from 0.4% for the graphite model up to 8.63% for the CERMET model, with the lowest error

cases pertaining to the more researched graphite and composite fuels. Considering the differing

assumptions that most likely went into development of each of the two models, along with a vast

difference in development time for the models, these comparison results provided strong support

for validation of the NTR model.

The NEP literature source used in validating the NEP architecture model analyzed a 2033

Mars crewed round-trip mission that used MPD thrusters and a fast power reactor. The exact

delta-V values used in the study were not given, thus a mission was set up in the NEP ephemeris

model to generate approximate values for the energy requirements. A known difference between

the paper-study analysis and the approximations used herein was that an Earth flyby was used in

the paper-study versus the direct Earth re-entry assumed in the ephemeris model. To account for

this difference, it was decided to match the propellant mass of 276 (kg) given in the literature

study in order to determine the correct delta-V for the return trip. Using an outbound delta-V of

19.3 (km/s) and return delta-V of 19 (km/s), a transfer time of 275 days, and MPD thruster

selection, the resulting IMLEO from the NEP architecture model was found to be 598.8 (MT).

The literature source quoted an IMLEO of 560.7 (MT).32 The NEP model thus gave a 6.8%

higher estimate of the IMLEO than did the source, a rather reasonable value given the

assumptions and approximations that had to be made to compare the two estimates.









A study on hybrid vehicles was also used in an attempt to validate the Hybrid architecture

model. The mission was again a round-trip crewed Mars mission with a short stay. Similar to

the previous validation methodology, a longer stay-time was assumed. This assumption was

made since energy requirements were not explicitly cited in the paper-study, and the ephemeris

model gave far less demanding energy requirements for a longer stay-time on the surface of

Mars. The outbound and inbound NTR delta-V requirements were found to be 3.17 (km/s) and

0.22 (km/s), respectively, while the NEP requirements were 14 and 11.8 (km/s). The literature

source cited an IMLEO of 298 (MT) for the hybrid vehicle with MPD thrusters, and the Hybrid

architecture gave a result of 275 (MT), a difference of only 7.7%.6

The validation results presented for each of the three architecture models satisfied the

objective of comparing the results seen in a literature study for each vehicle type, to those

generated by the models developed in this study. The outcome was highly satisfactory in that the

models created herein did a relatively good job of predicting the same results as professional

mission analysis results done by experts in each particular power and propulsion field. These

results were important since they provided credibility to the predictive nature of each of the

individual architecture models. This was crucial for conducting a meaningful trade study of all

three architectures, since they were compared directly against each other to determine the best

performing vehicle architecture.









Table 3-1. NTR Architecture Systems
NTR System Reactor Mode Energy Spectrum Fuel Type
1 Unimodal Thermal Graphite
2 Unimodal Thermal Composite
3 Unimodal Thermal Carbide
4 Unimodal Fast CERMET
5 Bimodal Fast CERMET


Table 3-2. Fuel Operating Parameters
Chamber Temp. (K)


ISP (s)


Power (MWt)


Graphite 2500 875 306.6
Composite 2700 925 316.6
Carbide 3100 1000 342.0
Cermet 2700 925 316.6


Table 3-3. NTR Reactor Mass Sizing


* prop. hardware added to total


Reactor (kg) Intern. Shield (kg)


Extern. Shield (kg) Total (kg)


Graphite 3665.0 988.4 2,942.0 9384.4
Composite 3062.3 825.9 2,458.2 8135.3
Carbide 3404.0 918.0 2,732.5 8843.5
Cermet-uni 1059.1 285.6 850.2 3983.9
Cermet-bi 1680.8 453.3 1,349.2 5272.3

Table 3-4. NEP Power Reactor Descriptions
Power (MWe) Mass (kg) Type Spectrum
5 9044 Liquid-metal Fast
1 3845 Particle bed reactorFast
10 17970 Li-cooled pin type Fast

Table 3-5. EP Thruster Characteristic Parameters
Type Efficiency ISP (s) Thrust (N) Spec Mass (kg/kWe) Fuel Lifetime (days)
Ion 0.90 6000 1.0 5.2 Xenon 850
MPD 0.65 5000 100.0 0.6 Hydrogen 350
Hall 0.60 2500 0.4 3.3 Xenon 100


Table 3-6. TransHab Mass Breakdown
Component Mass (kg)
Module 21000
ECRV 5100
Crew/Suits 600
Consumables 14.7/day









Table 3-7. Common Subcomponent Mass Breakdown


ADCS 70
C&DH 35
TT&C 205
RCS 6%*mass_prop
Structure 10%* total mass


Table 3-8. NTR Model Validation Mass Estimates
P&W Mass (kg) Models Mass (kg) % Error
Graphite 239000 238042 0.40%
Composite 200000 211738 5.87%
Carbide 180000 195854 8.81%
Cermet 170000 184666 8.63%


Mass (kg)


Component





























Figure 3-1. NTR Unimodal Vehicle Architecture Schematic




Poweirand H2 Propellant
Propulsion Tanks
Turbopump &
Pressure Vessel Feed System He/Xe Spaecraft
He/Xe



ReactorRadiators Ppn
'' ~ R~~t~l Prpellant Nanganron ~enjngiontrol Trans ab IECRV
Power Power Management Tee.."ni tin

SReactor ... Ra _Aran tavl ynt

Radiation Shield



Figure 3-2. NTR Bimodal Vehicle Architecture Schematic













Power

Pressure Vessel

Reactor



Pressure Vessel

Reactor
'.._ ; ;


He/XeCoolant

Radiators 1

Power Power Management
Convers'ion" and Distribution

Radiators


Radiation Shield


Figure 3-3. NEP Vehicle Architecture Schematic


20000

18000

16000

S14000
-
S12000
E
S 10000
U,
8000
-

8 6000
-

4000

2000

0


Spacecraft
Data m i!, i

Navigation Sensing/Control

Telecommunications

Attitude Control System

Electric
OPro ulsion


0 1 2 3 4 5 6 7 8 9 10 11
Power (MWe)


Figure 3-4. NEP Reactor Mass vs. Power


















1Ps


bP.


Proputaon Svyem



lpp. Ppp
P e PrC ................j............
S. p... .............-.....-.--.. .. PI

lo Mgsn & -g $PpCcraft
Pe Otr tPe spacecraft 105 kWe/CBC

* Note Pe=Pjrop Spacecraft power is separate from calculations


P,: Raw power from power source required by NEP system
rls: Efficiency of converting raw power to electric power
Pe: Electric power output of CBC required by NEP spacecraft
P : Power required for power conditioning
rpp: Efficiency of power conditioning for the thruster
Pth: Thruster Power
r6: Efficiency of converting electric power to thrust power


Figure 3-5. Power and Efficiency Flow Diagram


30000
y = 7772.7Ln(x) + 1067.5
25000


a 20000


| 15000
a

S10000


5000


0 4


0 5 10 15
Power (MWe)


20 25 30


Figure 3-6. PMAD Mass Sizing Equation












































Figure 3-7. Hybrid Vehicle Architecture Schematic







-a Departure_[
-a Departure_
-a Departure_
a Transfer tim
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c-.. Transfer tim
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c-. NEP_Thrust
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[ NEP_Thrust
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S Julian Conv
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El- Time_optimi;
Stay_low _b
Stay_upp_b
Transfer lo
Transfer_up
a Departure_
Departure_
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-a NEP Thrust




Figure 3-8. NEP Ephemeris Tool GUI


Spacecraft
Data Processing

Navigation Sensing/Control

Telecommunications

Attitude Control System
. .....[ ... .o .. '.

Electric
a Propulsion


Power aPd
Po r ad H2 Propellant
Propulsion Turbopump & Tanks
Turbopump --
--__ Pressure Vessel Feed System

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a.Conversioi and Distribution
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,ay 17.32
Month 9
ear 2040
e out 998
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er Eff 0 82
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ie_out 125852
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ie_ret 112318
piral_Time 386 52
mission delV 121 164
Mission_Days 2051
ission_Days 2381 7
zer
found 40
found 60
v bound 800
p_bound 1000
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-..a DestinationPlanet Mars
Hybrid_Ephemeris
+- H Hybrid_Standalone

c- IMLEO 250792
-..... Mass Inert 125127
a.o-. Mass_Propellant 112565
c-o, Powerelec 3507.49
-0... No_Thrusters 96.6348
-..... NTR DV1 3.17481
c-o. NEPDV2 14.03
NT-0 NTRDV3 0.22138
-o... NEPDV4 11.81
c-o. Crewed_Mission_Days 950
c-... NEP_Burn_Days 375.79


Figure 3-11. ModelCenter Mission


Architecture GUI









CHAPTER 4
TRADE STUDY RESULTS

Due to the breadth and depth of this mission analysis trade study, the data analysis phase of

the investigation was divided between three main areas: ephemeris analysis, vehicle architecture

analysis, and tradespace assessment. The ephemeris analysis utilized the ephemeris tools in

order to model the energy requirements for the three major vehicle systems based upon departure

dates and mission duration. The vehicle architecture analysis determined performance

parameters of the power and propulsion systems for these missions. The tradespace assessment

then used these results to assess the entire tradespace matrix based on pre-determined MOEs.

The ability to create the parametric and carpet plots that were needed for the data analysis

and tradespace analysis to follow were done using the time-saving functionality inherent to

ModelCenter. ModelCenter's data analysis toolbar was used to set various parameters for

plotting, producing multiple charts for different types of interpretation. Figure 4-1 shows the

actual ModelCenter Toolbar with intuitive buttons for each data analysis option, along with four

examples of plots that can be generated, including carpet, contour, surface, and statistical

dependence plots.

Ephemeris Analysis

One of the key steps in performing mission analysis is using ephemeris tools to determine

system energy requirements for a desired mission. Although the energy requirements for the

specific reference mission profiles were needed to assess the tradespace matrix, it was desired to

first derive some more general ephemeris results from the models created in this study for the

NTR, NEP, and Hybrid vehicles.

The single most important parameter to be obtained from the ephemeris models was the

delta-V value, which is essentially a requirement on how much energy will be needed by the









propulsion system. This requirement will determine how much propellant must be carried on

board in order to get the useful hardware to its destination in space. The analysis to follow will

allow one to see how this delta-V parameter is dependent on other variables such as stay time,

transfer time, and departure date, the effects of which are determined by celestial mechanics.

The following plots attest to the capabilities of both the ephemeris models developed for

this study using legacy NASA codes, and the ModelCenter software tool used to run the

ephemeris models. Their presentation is separated between departure and duration categories,

which refer to whether the delta-V parameter was plotted against the departure year and month,

or versus transfer and stay time.

NTR

Departure

The LEO departure date is one of the major factors in determining delta-V values with

ephemeris models. This is why the departure year, month, and day were all used in the process

of trying to find the optimal time to leave the Earth. The idea of the synodic periods for each

planet can be visualized in Figure 4-2 as a wave-like pattern is seen for the delta-V values over a

range of 15 years. The peaks and valleys of each curve represent the cyclical nature of the

location of the Earth relative to each destination planet over time. It can be seen that the number

of cycles is directly proportional to the distance from Earth, giving more minimum-energy

opportunities to Mars, and less to Saturn. Note that the data seen in this plot, and similar ones to

follow, assumes a transfer time of 200 days for Mars, and 1000 days for both Jupiter and Saturn,

with a 500-day stay at all planets.

Contour plots were also created to assess how the delta-V energy requirements were

affected by both departure month and year. Figure 4-3 shows the dependence of the energy

requirement on both of these parameters, where the color legend in the upper right-hand corer









relates delta-V values in units of (km/s) to the colored lines seen on the plot. This plot indicates

that any year and month between the two green curves in the upper left-hand area would be the

optimal time of departure. The range of values for delta-V based on all results is between 39.1

and 109.8 (km/s), though changing the stay and transfer time would certainly provide flexibility

in further optimizing these results.

The Saturn and Jupiter contour plots could not be created for the NTR model using the

IPREP ephemeris code. Consistent with prior knowledge of this code, the simulations seemed to

break down at these more distant planets and years, specifically with Saturn, and at years after

2042. This also meant that the optimizer tool could not be used to optimize the Jupiter and

Saturn reference missions, thus they had to be estimated using a manual optimization process.

Duration

In addition to departure date, the transfer time and stay time play a major part in

determining energy requirements. Delta-V versus transfer time curves were plotted for all

destination planets across a range of 200 to 1200 days, fully encompassing the ranges allowed

for the reference missions. It can be seen in Figure 4-4 that the requirements for the Mars

missions actually get more difficult after 300 transfer days, and the Jupiter and Saturn

requirements get less difficult after approximately 400 days. This observation can again be

related to the distance of each planet from the Earth and/or the sun. Note that the departure year

used for the Mars analysis was 2031, while that for Jupiter and Saturn was 2041.

The delta-V values were also plotted against stay time to see how this parameter affects

requirements. It can be seen from Figure 4-5 that each planet's curve has two valleys and one

peak. This seems to support the idea of using conjunction and opposition-class missions since

the valleys correspond relatively to the 40 to 60-day stay opposition-class missions, and the

longer 400 to 600-day stay of the conjunction-class missions.









Next, the energy requirements for each planet were assessed separately using a surface plot

to see the affects of both transfer time and stay time on delta-V. Figure 4-6 shows the results

given a transfer time range of 100-300 days, and a 400-600 day stay time range. It appears the

NTR system would have a very difficult time completing this mission at low transfer times that

also have a short stay time, but will have a much simpler mission at all other transfer and stay

times.

Jupiter's surface plot shows a delta-V assessment given transfer times of 1000 to 1200

days and stay times between 400 and 600 days. This plot, seen in Figure 4-7, shows much more

variation in energy requirements across the transfer and stay time ranges along with a more even

influence of either independent parameter on delta-V.

Saturn's contour plot in Figure 4-8 reveals much smoother curves than that of Jupiter's

when plotted over the same transfer and stay time ranges. However, similar patterns are seen

between the areas of maximum and minimum delta-Vs. Jupiter's optimum point for the specific

departure date used is around 900 days for the transfer and 500 days for the stay time. Saturn's

optimal mission profile is around 1150 transfer days and 500 stay days for the same departure

date as Jupiter.

NEP

Departure

The NEP mission profile analysis plots were obtained just as they were for the NTR

ephemeris model, with expected variations to be seen due to the different type of propulsion

system. The delta-V versus departure year plot seen in Figure 4-9 shows both similarities and

differences compared to the comparable NTR plot. A similar synodical pattern is seen in the

Mars curve, which has exactly six peaks, just as in the NTR curve. The Jupiter and Saturn

curves, however, appear somewhat different, as only one peak and valley can be seen for each.









This may be due to the fact that the NEP system has less propulsive power and will be

accelerating much slower to its destination, thus increasing the periods of the curves of these

more distant planets.

The contour plot of delta-V vs. departure month and year for Mars is shown in Figure 4-

10. It can be seen that the optimal departure is in the early months of 2031 and a bit later in the

year 2033, representing the Earth-Mars 2.13 year synodic period. The delta-V values range from

40.2 to 99.7 (km/s) at these departure dates.

The departure contour plot for Jupiter can be seen in Figure 4-11. This plot shows a much

larger relative area of optimal energy requirement as compared to the Mars plot. This is in

agreement to the earlier finding that there are more cycles over time in the Jupiter and Saturn

curves versus Mars for the same range in years. This plot serves to portray the 1.09 year Earth-

Jupiter synodic period, as similar delta-V minimums and maximums are repeated just over every

12 months. However, it is apparent from the high delta-V results ranging from 73.6 to 133.4

(km/s), that further optimization of mission parameters would be needed to make this flight

practical.

The Saturn contour plot in Figure 4-12 shows the trend seen in the Jupiter plot even more

clearly as much less variation is seen for delta-V requirements over the range in dates plotted.

Most of the plot reveals an area of optimal departure dates, while only the lower left-hand corner

shows departures dates of high delta-V. Again, the synodic cycle of 1.03 years can be seen as

the delta-V values repeat just about every year. The delta-V values for the Saturn mission are the

highest, ranging from 126.4 to 169.6 (km/s).

Duration

The dependence of delta-V on transfer time for the NEP missions can be seen in Figure 4-

13. In comparison to the NTR plot, this plot shows fewer peaks in delta-V across the spectrum