Nuclear Systems for Mars Exploration

Nuclear Systems for Mars Exploration

(Parte 1 de 8)

2004 IEEE Aerospace Conference Proceedings

Nuclear Systems for Mars Explorationl.

Tibor S. Baht

Jet Propulsion Laboratory, California Institute of Technology 4800 Oak Grove Drive, MIS 301-17OU

Pasadena, CA 91 109-8099 818-354-1 105 tibor.balint@jpI.nasa.gov

Abstract-This paper identifies breakpoints for various power and propulsion technologies, with a special focus on fission-based sources in support of NASA’s Mars exploration program. Transportation, orbital, and surface missions are addressed through an assessment architecture developed for this study. This architecture is based on three key considerations: decomposition of generic Mars missions into phases, a lumped parameter approach, and a bounding case analysis. With these simplifications breakpoints are identified beyond which new technologies, such as nuclear fission power, are required to achieve mission objectives. It is found that in-space propulsion and power generation are sized by launch vehicle delivery limits and trajectory options. Similarly, power levels for surface-based reactors are affected by transportation system and EDL limits imposed by current technologies. Ailer summarizing the breakpoints for today’s state of the art, development targets are identified to enable space-based nuclear power and propulsion systems to perform at their full potential.

1. INTRODUCTION1
2. ASSESSMENT ARCHlTECTU RE2
3. TRANSPORTATION kWES2
4. EDL ISSUES7
5. POWER GENERATION ISSLIES9
6. CONCLUSIONS16
REFERENCES18
BIOGRAPHY20
ACKNOWLEDGMENTS17

1. INTRODUCTION

As the latest phase of our Mars exploration program draws to completion, NASA continues working on future mission concepts. Contributing to this effort, JF’L’s Advanced Mission Studies Office identified three possible exploration pathways, forecasting the next four decades. Assuming the success of our current missions, the first path in the roadmap covers a mainstream approach termed “cuent pathway”. It follows a conventional path with a number of planned missions for the next decade, followed by robotic and subsequent human explorations in the third and fourth decades. The second path, termed “reduced scope or go competed”, reflects an uncertainty of future exploration efforts, reacting to unfavorable technical, economical, or political influences. This rather pessimistic path sees the completion of the upcoming Mars Science Laboratory (MSL) program, but without predicting any follow-on missions. The third path responds to “momentous discoveries”, such as finding signs of past (or present) life on Mars, and consequently igniting an accelerated golden age of exploration in the form of advancing robotic exploration as early as the next decade, with subsequent human exploration and colonization of Mars during the following two decades (Figure 1). However, finding existing life on Mars could result in a delay to colonize

Mars, until bioethical issues are resolved.

The first two Mars exploration pathways predict a gradual increase in mission complexity, offset only by program timeline. To achieve the objectives of these missions, enabling technologies (e.g., propulsion and power systems) must evolve and in effect change our current technology paradigm. As shown in Figure 2, today’s space exploration can be characterized by mass dependence and consequent power limitation. Launch vehicle technologies limit the

Id decde

‘96 - ‘OS 2a decade Pd dewde I d5 decade Scensno 1 ‘@)-.I8 ’20-.30 1 ’j0-.40 c\

’ 0-7803-8155-61041$17.0OD 2004 BEE ’ IEEEAC papcr #I 118, Version 3, Updatcd Novcmbcr 19,2003

0-7803-81 5-6/04/$17.0 0204 IEEE 1 2958

- Figure 1- Mars exploration program evolution [I]

Authorized licensed use limited to: UNIV ESTADUAL PAULISTA JULIO DE MESQUITA FILHO. Downloaded on July 13, 2009 at 17:38 from IEEE Xplore. Restrictions apply.

5dJl sail

P* &Yn Fmuc I- --.. - " not _==. Figure 2- Technology paradigms maximal deliverable mass to Earth orbit and beyond. Space missions are designed around these hounds, thus hampering power availability for transportation, science, and housekeeping. Chemical propulsion, kel cells, and batteries belong here, characterized by restrictions to both power and duty cycle. In the future, a new paradigm can be envisioned, where advanced propulsion and power sources would provide power far beyond our current limits. While for the distant future we may consider exotic power sources

based on antimatter or nuclear fusion, for the near term nuclear fission power is the most likely candidate. Solar sails and tethers do not generate power by themselves, and may operate for an extended period of time; hence these solutions belong to a time-dependent category, not explored Mer in this study.

Thus, this paper focuses on two of the three paradigms, one dependent on mass and the other on power, with an emphasis on the role of nuclear fission power. First, the assessment architecture of this study is introduced, followed by descriptions and performance characteristics of o Transportation o Entry, descent, and landing (EDL) o Fission and decay-based power generation o Other conventional technology options

After identifying the limitations of today's technologies, key areas are stated where advancements could facilitate a transition from one paradigm to the other, demonstrated within the framework of the Mars Exploration Roadmap.

2. ASSESSMENT ARCHITECTURE

Power and propulsion system technologies cover a broad range of options developed to various technology readiness levels (TU). Space missions to date have utilized these technologies, which have been selected based on mission objectives. Therefore, to assess the breakpoints beyond which nuclear power sources represent the only viable alternative, their performance must be compared against more conventional technologies.

The assessment architecture used for this study consists of three components. The first reduces the number of parameters to only a few; the second limits the sensitivity analysis to the upper bounds of these parameters; and the third decomposes a generic Mars mission into distinct stages.

Lumped Parameter Approach

Space mission complexities pose hard challenges reflected through a multitude of dependent parmeters. To account for all is beyond the scope of this work. Instead, a lumped parameter approach is adopted, reducing these parameters to only mass, power, and time, from which all other parameters can be derived. Based on these key parameters main technology breakpoints are identified.

Bounding Case Approach

Technologies are sometimes scaleable and cover a wide application spectnun. The bounding case approach, adopted here, helps to minimize assessment effort by identifying these upper limits or technology breakpoints beyond which new technologies are needed to achieve mission objectives.

Mission Stages

A typical Mars mission consists of a partial or a full set of the following three stages:

o Transportation stage o In-orbit stage o On-surface stage

Each of these can be characterized by mass, power, and time.

With this methodology, generic mission concepts can be tested inexpensively on a conceptual level. Once one or more favorable answers are reached, further in-depth studies are needed to address utility and to identify the best candidate configuration for a given set of mission objectives. (Note that other considerations, such as safety and cost - though critical - are not discussed.)

3. TRANSPORTATION ISSUES

The orbit of Mars is more eccentric than that of Earth. It is at -1.5 AU (-1.4 to 1.6 AU) from the Sun, with an orbital inclination of I.85", relative to Earth. Due to orbital phasing, low-energy launch opportunities to Mars occur about every two years. It is typical to launch the spacecraft to parking Low Earth Orbit (LEO), orient it to an appropriate inclination, and then launch it to a transfer orbit between Earth and Mars with the last stage of the launch vehicle or using an onboard propulsion system. From that point on the spacecraft follows a trajectory, which is based

Authorized licensed use limited to: UNIV ESTADUAL PAULISTA JULIO DE MESQUITA FILHO. Downloaded on July 13, 2009 at 17:38 from IEEE Xplore. Restrictions apply.

on its initial impulse or its onboard propulsion system or both. Before placing any propulsion and power systems into the transportation framework, it is important to note the competing propulsion technologies and trajectory options. Cornhinations of these options define a Mars mission architecture trade space. Trajectories are influenced by launch date and propulsion system options. Similarly, trip time and payload mass requirements call for a suitable propulsion system. Hence, both trajectory and propulsion system options are discussed below.

Trajectories

The three main orbital transfer pathways are: o Ballistic (using high-energy impulse) o o Cyclers Low thrust (but high specific impulse, Isp) or

Each of these can be subdivided based on trip time and energy [2]. Due to the phasing between Earth and Mars and

the departure time, return trip missions can be optimized for a number of variables, for example shortest mission time, fastest transfer time, longest surface stay, or largest deliverable mass. Figure 3 summarizes these return trajectories, applicable to both manned and sample return missions. Detailed description of these trajectories is given in [3], and [4]. For most one-way scientific and cargo missions, the transfer time corresponds to the outbound leg of a given return trajectory. One-way manned missions do not have a mainstream acceptance; however, they can result in a 25 to 35% cost saving while building up a colony and resources on Mars [5][6]. It should be noted that trip times are dependent on assumptions for the propulsion system, final mass to be delivered to Mars, and the launch time. Therefore, the numerical values provided in this paper should be viewed only as rough estimates.

High-thrust frajectories refer to ballistic transfers. While a Hohmann transfer does not take planet phasing into account and in this pure form the trajectory cannot be used, it is considered an ideal transfer, minimizing the total energy.

Such a trip to Mars requires a round trip AV of 1.2 km/s. This can only be achieved by dividing the round trip into an outbound leg, a stay period, and a return leg. The resulting round trip time is 2.6 years (-971 days), including a stay time of 1.24 years (453 days) [7]. Type 1 and 2 transfers can be faster than Hohmann transfers and can he envisioned as Hohmann transfers to a dummy orbit beyond Mars, terminating and/or initiating at one of the two Mars orbit crossings. Type 1 (Tl) round trips transfer to the first Mars orbit crossing. For this case the stay time increases to 1.65 years, hut the round trip time is only reduced by 0.2%. Type 2 (T2) round trips pass the target orbit and transfer at the second opportunity on the way back. It reduces roundtrip time by about 8%, but significantly reduces stay time (to 0.43 years) [7]. Free-return flyby is the simplest and least energy intensive trajectory, based on the ballistic

Hohmann transfer and a single spacecratt, which minimizes propellant requirements. During the outbound trip the spacecraft passes Mars, achieving about 2 hours of optimal viewing. There is insufficient time for a piloted landing. The return requires 1.5 heliocentric revolutions due to planet phasing. The total flight time is about 3 years. While it fulfills technological requirements, the long flight time and short stay time makes this option undesirable [3][4].

Flyby-rendezvous or short-stq are similar to the ballistic transfer type free-retum flyby, but they utilize two spacecraft. The first spacecraft arrives and lands 30 days before the second spacecraft's flyby. It takes off in time to rendezvous with the second spacecraft and does the same return as the free-rehun flyby mission. It still results in a proportionally too long transfer time compared to a short 30-day stay. For this case and also for the free-return flyby, the Earth-Mars flight time is -230 days, while the Mars- Earth return flight-time is -840 days due to planet phasing.

[3][4] Conjunction class represents a long-stay mission architecture, with a total round trip time of -950 days, which includes a stay time of up to 560 days [SI. Such a mission may require a Saturn class rocket and In-Situ Resource Utilization (ISRU). Fast 150-day one-way trip times would need Nuclear Thermal Propulsion (NTP)

(sometimes referred to as Nuclear Thermal Rocket (NTR)). Opposition class represents a short-stay architecture, with short outbound and long return transit times (or reverse). The advantage is a short 1.6 to 1.9 year total mission time without or with Venus swing-by, respectively. The Venus swing-by is more favorable from an energy point of view, but it subjects the spacecraft to greater thermal and radiation loads. Minimum energy opportunity for this class occurs every 26 months. The stay time on the surface is, however, only 30 days; hence 95% of the total mission time is spent in transit. Niehoff et al. [3] provides a summary table of available launch dates between 2002 and 2015 for conjunction, opposition, and sprint class trajectories. The above "brute force" methods are expensive, requiring the

Authorized licensed use limited to: UNIV ESTADUAL PAULISTA JULIO DE MESQUITA FILHO. Downloaded on July 13, 2009 at 17:38 from IEEE Xplore. Restrictions apply.

largest available launch vehicles (e.g., Delta 1V Heavy or Titan IV).

Low-thrust trajectories require longer trip times than those for high-thrust trajectories, and the spacecraft spends a significant time crossing the Van Allen radiation belt. Due to a higher specific impulse propulsion system, the delivered mass may be higher. Adding AY by a Venus swing-by can further reduce propellant requirement and consequently further increase payload mass, but it adds to the trip time. Venus is about 30% closer to the Sun than Earth). Low thrust transfer options are used by electric propulsion systems. A typical mission includes a 50-day outward spiral from LEO. AAer reaching escape velocity the trip to Mars takes about 510 days. Another 40-day spiral-in follows from Mars orbit capture to reach Low Mars Orhit (LMO). The stay time is between 100 and 200 days. The retum trip consists of a 25-day spiral-out from LMO, a 230-day transfer from Mars to Earth, and a 16-day spiral-in from Earth capture [3]. For manned missions it is suggested that, to avoid the Van Allen radiation belt, embark the crew on a high-thrust rocket and join the low- thrust spacecrafi after it spiraled beyond Geosynchronous

(Parte 1 de 8)

Comentários