4
Shielding Approaches and Capabilities

BASIC SHIELDING CONCEPTS

In principle, there are three ways to reduce a dose from external radiation:

  • Increasing the distance from the source,

  • Minimizing the time of exposure, and

  • Using shielding.

Increasing the distance from the source is often the simplest to implement for terrestrial radiation protection (e.g., the use of tongs to minimize exposure to hands, moving work spaces farther away from sources). In deep space, however, the spacecraft and crew are immersed in the radiation environment. Using distance from the source as a means of mitigating exposure is not possible.

Limiting the time that the crews are exposed to a radiation field was used in the early space program. During the Apollo era, human missions to the Moon used this method of minimizing exposure to the Van Allen radiation belts on the transits to and from the Moon. Transit times to Mars, however, may be driven by other considerations, such as available means of propulsion and spacecraft trajectory.

The most commonly used means of protecting terrestrial radiation workers is through the use of shielding (i.e., the placement of material between the human and the radiation source in order to reduce the intensity of the radiation field at the human’s location). In principle, shielding alone should be able to reduce exposure by attenuating the radiation and reducing the dose rates. For deep space missions, however, shielding alone cannot guarantee protection in all situations owing to the very high energies of the incident ions and the production of highly penetrating secondary particles, such as neutrons and light ions, coupled with mass constraints on the spacecraft and the large uncertainties in biological risk.

SHIELDING FOR PROJECT CONSTELLATION’S ORION

For operations within Earth’s geomagnetic field, little or no supplemental shielding is needed to ensure astronaut safety in a capsule or habitat. However, upon leaving this protective geomagnetic shield, the astronauts are subjected fully to the natural GCR environment and susceptible to serious radiation fluxes from solar particle events (SPEs).



The National Academies | 500 Fifth St. N.W. | Washington, D.C. 20001
Copyright © National Academy of Sciences. All rights reserved.
Terms of Use and Privacy Statement



Below are the first 10 and last 10 pages of uncorrected machine-read text (when available) of this chapter, followed by the top 30 algorithmically extracted key phrases from the chapter as a whole.
Intended to provide our own search engines and external engines with highly rich, chapter-representative searchable text on the opening pages of each chapter. Because it is UNCORRECTED material, please consider the following text as a useful but insufficient proxy for the authoritative book pages.

Do not use for reproduction, copying, pasting, or reading; exclusively for search engines.

OCR for page 62
4 Shielding Approaches and Capabilities BASIC SHIELDING CONCEPTS In principle, there are three ways to reduce a dose from external radiation: • Increasing the distance from the source, • Minimizing the time of exposure, and • Using shielding. Increasing the distance from the source is often the simplest to implement for terrestrial radiation protection (e.g., the use of tongs to minimize exposure to hands, moving work spaces farther away from sources). In deep space, however, the spacecraft and crew are immersed in the radiation environment. Using distance from the source as a means of mitigating exposure is not possible. Limiting the time that the crews are exposed to a radiation field was used in the early space program. During the Apollo era, human missions to the Moon used this method of minimizing exposure to the Van Allen radiation belts on the transits to and from the Moon. Transit times to Mars, however, may be driven by other considerations, such as available means of propulsion and spacecraft trajectory. The most commonly used means of protecting terrestrial radiation workers is through the use of shielding (i.e., the placement of material between the human and the radiation source in order to reduce the intensity of the radiation field at the human’s location). In principle, shielding alone should be able to reduce exposure by attenuating the radiation and reducing the dose rates. For deep space missions, however, shielding alone cannot guarantee protection in all situations owing to the very high energies of the incident ions and the production of highly penetrating secondary particles, such as neutrons and light ions, coupled with mass constraints on the spacecraft and the large uncertainties in biological risk. SHIELDING FOR PROJECT CONSTELLATION’S ORION For operations within Earth’s geomagnetic field, little or no supplemental shielding is needed to ensure as- tronaut safety in a capsule or habitat. However, upon leaving this protective geomagnetic shield, the astronauts are subjected fully to the natural GCR environment and susceptible to serious radiation fluxes from solar particle events (SPEs). 62

OCR for page 62
63 SHIELDING APPROACHES AND CAPABILITIES The return to the Moon plan for the Constellation program is quite similar to that of the Apollo program. However, Constellation will use a two-launch method. The Ares V Heavy Lift Launch Vehicle will launch the Earth Departure Stage (EDS) containing the propulsion system and the Lunar Lander. The four-member crew will be launched in the Orion Crew Exploration Vehicle to low Earth orbit (LEO) using the Ares 1 Crew Launch Vehicle, which will then dock with the EDS before heading to the Moon. Once the combined vehicle is in lunar orbit and systems checks have ensured a “Go” for landing, the Lunar Lander will separate from the Orion and initiate the lunar landing. The Orion Block-2 design requirements respond to two major mission scenarios: • Supporting crew and cargo transportation to the International Space Station (ISS) in LEO and returning them to Earth, and • Transporting crew to low lunar orbit (LLO) and returning them safely to Earth. As part of the lunar mis- sions, the Orion will rendezvous and dock with the Lunar Lander (already mated to the EDS) in LEO, and will provide piloting, guidance, and navigation to the combined cislunar spacecraft. On both of these missions, the Orion capsule would be the principal shielded volume; that is, it would be either the only volume, or it would be the most likely shielded module from which the crew could fly the mission and still benefit from a level of protection. The requirements provided by NASA to Lockheed Martin are as follows (T. Shelfer, Lockheed Martin, “Crew Exploration Vehicle RequirementsContractor’s Perspective,” presented to the committee on February 21, 2007): • Orion shall provide radiation protection, consistent with the principles of ALARA [As Low As Reasonably Achievable], to ensure that the tissue-averaged effective dose to any crew member does not exceed 15 cSv for the worst-case SPE, defined as the King parameterization of the August 1972 event. • The vehicle shall continuously measure and record the external fluence of particles of Z < 3, in the energy range 30 to 300 MeV per nucleon and particles of 3 ≤ Z ≤ 26, in the energy range 100 to 400 MeV per nucleon and integral fluence measurement at higher energies, as a function of energy and time, from a monitoring location that ensures an unobstructed free space full-angle field of view 65 degrees or greater. • The vehicle shall provide an omnidirectional, portable system that can continuously measure and record the dose equivalent from charged particles with linear energy transfer 0.2 to 1,000 keV per micrometer, as a function of time, at an average tissue depth of at least 2 mm. Preliminary analyses by NASA and Lockheed Martin indicate that the Orion capsule provides adequate shielding from its structure, avionics, life support, other hardware, consumables, and waste storage such that lower-energy SPEs would not be a threat. However, for the rarer, higher-energy events, doses could accumulate beyond the acceptable limit. For this reason, the Orion capsule itself must either incorporate sufficient shielding or else have the capability to reconfigure shielding and functional hardware to provide a radiation storm shelter for the astronauts. The duration of the most hazardous portion of an SPE or a close series of SPEs can be hours to a few days. Thus, the Orion capsule must be capable of providing the storm-shelter capability for a somewhat extended period of time, and astronauts must have access to food, water, and minimum hygiene facilities. Although it will be permissible to leave the radiation storm shelter for short periods (minutes to fractional hours) to meet personal needs or perform a task required for mission success, the astronauts should spend the duration of an SPE inside the shelter. As stated above, protection is accomplished by time, distance, and shielding. In this case, the astronauts can safely leave the shielded area if they return quickly. Lockheed Martin designers considered several solutions: hull shielding, deployable water shielding, shielding integrated into seats, and a deployable, high-density polyethylene (HDPE; defined by a density of greater than 0.94 g/cm3). At the time of this writing (summer 2007), the Orion project plans to provide 2.5-cm-thick slabs of HDPE for use by the astronauts to configure an in-space shelter inside the Orion capsule itself. The HDPE shield was the only one that could feasibly provide the necessary amount of shielding. The shielding would be 2.5 cm thick and would be stowed on the floor of the Environmental Control and Life Support System when not in use. At the time of this analysis, Orion was still in its first design cycle. More detailed radiation analyses and shielding configurations are planned in future iterations (T. Shelfer,

OCR for page 62
64 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION Lockheed Martin, “Crew Exploration Vehicle RequirementsContractor’s Perspective,” presented to the committee on February 21, 2007). Two instruments will be used to meet the monitoring requirements: the Radiation Assessment Detector (RAD) and the Tissue Equivalent Proportional Counter (TEPC). The RAD will be based on the instrument of the same name discussed in Chapter 2, used to monitor the martian radiation environment, in order to decrease develop- ment cost and leverage another currently planned project. The TEPC will be based on a design from Texas A&M University. Both of these instruments will take advantage of the ISS as a testing and evaluation environment (T. Shelfer, Lockheed Martin, “Crew Exploration Vehicle RequirementsContractor’s Perspective,” presented to the committee on February 21, 2007). LUNAR LANDER The Lunar Lander will be a highly specialized vehicle designed for operation in LLO, from LLO through terminal descent and landing, and then ascent from the lunar surface to rendezvous with Orion back in LLO. As described in the Exploration Systems Architecture Study (NASA, 2005) and NASA’s May 2006 Lunar Lander request for information (NASA, 2006), the Lunar Lander consists of three pressurized, habitable volumes: the Ascent Stage, the Descent Stage Habitat, and the Extravehicular Activity (EVA) Airlock. The airlock is discussed later in this chapter. All three portions of the Lunar Lander will travel from LLO to the Moon’s surface. The Descent Stage will carry the habitat for the crew to use during the surface segment of the mission, baselined at 16 m 3 (NASA, 2004). The crew would live primarily in this module during their stay on the surface, and they would retain access to the designated 10 m3 volume in the Ascent Stage. The Descent Stage and the Airlock would be left on the surface when the astronauts return to LLO in the Ascent Stage. The Ascent Stage represents the highest cost in total system mass, so it places the greatest premium on lightweight construction and will likely offer minimal shielding. However, the amount of time that crew members spend exclusively in the Ascent Stage will be quite limited, on the order of hours (G. Yoder, NASA, “Lunar Architecture Team Overview,” presented to the committee December 12, 2006.) The habitat is the most suitable module to afford radiation protection to the crew. Although the weight restric- tions are stringent, they may allow the provision of some radiation shielding in addition to the basic pressure-vessel structure, and thermal and micrometeoroid protection. The advantage of making the habitat serve as a radiation storm shelter is that it already contains the habitability accommodations to sustain the crew in relative comfort over the course of a solar storm. These habitability accommodations include life support, food systems, hygiene and waste management systems, sleep stations, and stowage for clothing and personal articles. The process used in the Orion capsule—developing a three-dimensional computer-aided design shielding model and a crew-radiation-exposure assessment based on some of the historical large SPEs—is also appropriate and reasonable to replicate for the various components of the Lunar Lander. SURFACE INFRASTRUCTURE In the past 2 years, NASA has published several versions of the lunar exploration timelines. Table 4-1 displays the timeline presented at the 2nd Space Exploration Conference (Lavoie, 2006), which was also presented at the committee’s first meeting by Geoffrey Yoder, NASA. The findings and recommendations of this report are still applicable to missions with slightly different timelines (such as those in NASA, 2004, 2005, 2006). Table 4-1 shows that in early years, missions will be short. The crew will need to bring along just about every- thing they need. However, the long-term goal is to gradually build up an outpost capable of supporting longer and more complex missions. The surface infrastructure, therefore, plays a significant role in controlling the total mission radiation dose.

OCR for page 62
65 SHIELDING APPROACHES AND CAPABILITIES TABLE 4-1 Potential Durations of Lunar Exploration Mission Mission Type Duration Outpost buildup, years 1-2 7 days Outpost buildup, year 3 14 days Outpost buildup, years 4-5 30 days Outpost occupancy 6 months Total/maximum mission duration >6 months SOURCE: Lavoie, 2006. Habitats Most space habitats, such as Skylab, Soyuz, Mir, and ISS modules, have been cylindrical, disk-shaped, or, in rare cases, spherical. Curved shapes provide the minimum thickness of the module walls for a given pressur- ized volume, responding to the need for minimizing mass (and therefore cost) of transportation into space. For the approximately isotropic incidence of space radiation, spherical shields provide the minimum mass for a given volume; however, a long cylinder is superior, presenting the minimum shielding at the ends, where it subtends the smallest solid angle; and that incoming radiation will have to traverse an increased amount of material, seen at an angle. These examples are for modules in free space, but the same will be true for habitat modules to be deployed to the Moon or to the surface of Mars. The difference is in shielding priorities. In free space, the only mass avail- able for shielding is the construction of the module itself, equipment, and supplies. Orion, on its longer missions, may be afforded a small amount of shielding-dedicated mass. However, on the surface of a planet, a habitat does not need to provide shielding in all directions; the planet itself grants 180° of protection, making a half-cylinder or dome more efficient as well as more practical. If the habitat is intended to provide the major component of radiation shielding during transit to the Moon, it may be useful to place the ascent and/or descent stages on the weakly shielded bottom of the habitat. Robots A menagerie of robotic concepts has been proposed to assist astronauts on lunar or planetary surfaces. A robotic assistant could carry a mobile, deployable solar-particle storm shelter with a mass of several hundred kilograms, accompanying an astronaut team or buddy pair on an EVA excursion. In the event of an SPE at a time when the crew is far from the base, crew members could unfold and deploy this shelter and wait out the storm for the few hours of greatest intensity. This strategy depends on the amount of life-support consumables carried by the crew (typically on the order of 8 to 12 hours, possibly increased by the robot’s carrying capacity) as well as their own comfort spending long hours immobile in the EVA space suit. Unpressurized Rover The unpressurized Lunar Rover Vehicle (LRV) used during the Apollo 16 mission proved a safe and reliable vehicle for the lunar crews. The rover extended the area of exploration outside the lunar module from meters to kilometers, allowed a greater mass of samples to be returned, and beamed images back to Earth. The LRV transported astronauts to places they never could have reached on foot. However, the astronauts were limited in their traverses by the available water and oxygen in their personal life support system (PLSS) backpacks. In the event the LRV broke down, the astronauts would need enough air and water to return safely to the Lunar Module. Unpressurized rovers will still be used in Constellation for short jaunts and trips around the outpost complex. They could still carry some portable radiation shielding in case of a breakdown, but their range is already limited, so it is likely that an astronaut could return to the outpost in time to avoid significant SPE exposure.

OCR for page 62
66 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION Pressurized Rover Several initiatives during the 1980s and 1990s considered larger, pressurized lunar rovers that could carry greater quantities of consumables for the astronauts of post-Apollo missions. A pressurized vehicle for lunar exploration was seen as the next logical step. Several universities, NASA centers, and NASA contractors produced a number of pressurized lunar rover designs. A notional pressurized lunar rover (PLR) considered by Anderson et al. (2006) has a mass of about 5,150 kg and a traverse range of several hundred kilometers. Like the space suits that it supports, the pressurized rover would essentially be a complete spacecraft in which the crew could live while exploring the lunar surface. The rover could incorporate fuel cells to provide mobile power and water for the life-support system on long journeys. These cells produce water and electrical power by combining lunar oxygen and hydrogen brought from Earth. With this enhanced mobility, the crew can service remote facilities, such as lunar telescopes, and conduct long-range geological traverses while making geological surveys and sample collections. The lunar south pole is at the edge of the largest crater in the solar system, approximately 2,500 km across. Exploring areas this large will require traveling for weeks on end. However, extending the EVA time also increases the probability that it will coincide with an SPE. Because it is mobile, the pressurized rover—like the other rovers—could not depend on in situ shielding at the outset of an SPE unless it found a convenient, naturally occurring “garage” in a lunar lava tube or cave. Therefore, this pressurized rover would need to carry its own radiation shielding to protect a crew throughout an SPE. A three-dimensional computer-aided design radiation shielding model could be developed for the pressurized rover, and parametric SPE radiation analyses be performed to ascertain the intrinsic shielding capability of the pressurized rover. If needed, it could also carry deployable radiation shielding materials, such as the polyethylene shields carried by Orion. In Situ Shielding Bringing all the radiation shielding all the way from Earth to the lunar or martian surface poses a substantial mass penalty. The most popular ideas for mitigating this mass penalty involve using surface resources to provide shielding. These resources include regolith, lava tubes, caves, and other landforms. Regolith Shielding In situ shielding made from regolith is an appealing possibility because regolith is found everywhere on the lunar surface and certainly will be available at the south pole where NASA’s Outpost-First Strategy plans to estab- lish the first lunar base. The poles tend to resemble the highlands more than they resemble the maria, so it is less likely that explorers can find a lava tube or cave at the eventual location of choice for the outpost site. Regolith consists of dust, rocks, and pulverized rock, all typically composed of aluminosilicates and iron oxide minerals. Approximately one-half of the mass in such materials is oxygen, and because of this the average atomic number of the material is as low as or lower than aluminum (Z = 13). Although not as optimal as high- hydrogen-content materials in terms of shielding effectiveness per unit of mass, regoliths are nonetheless as least as effective as aluminum (Zeitlin et al., 2002). The simplest approach to shielding using regolith is to place the module on or as close as possible to the ground and to construct a simple embankment by piling material against the module to shield in the azimuthal directions. Using regolith to shield in the zenith direction can be aided by providing a simple scaffolding or flat- roof structure on which to load additional material. The ends of the modules, where airlocks may be located, can be effectively shielded by creating a blocking berm at a standoff distance. As long as the line of sight to space is blocked, there is no requirement to providing a contiguous shield. The regolith can be loose, or converted into convenient forms such as sandbags or pressed blocks. Habitat designs that consist of a module on long legs (e.g., the Apollo lunar module) or a landing truss could be much more difficult to shield using regolith because of the challenges of raising tons of material to the necessary heights and providing sufficient support. Simonsen (1997) describes a method for determining an appropriate regolith thickness. First, one selects a radiation environment scenario—a combination of GCR and SPE exposure representing “worst-case conditions.”

OCR for page 62
67 SHIELDING APPROACHES AND CAPABILITIES (See Chapter 2 for further discussion on this topic.) Next, the radiation transport codes HZETRN and BRYNTRN are used to compute the effective dose behind a slab of regolith as a function of thickness. This curve is used to identify a candidate thickness or thicknesses that would keep effective dose below the permissible exposure limit (PEL). Then, the HZETRN and BRYNTRN analyses are repeated, this time using the selected thickness of shield- ing over the actual geometry of the shielded habitat. If the resulting effective doses are below the PEL, then the shielding thickness is acceptable; otherwise, the shielding thickness must be increased. The following example will give a rough idea of the amount of mass involved. In a basic but realistic analysis, Simonsen (1997) estimates that 90 metric tons of regolith is adequate to shield a cylindrical habitat 12.2 m in length and 4.6 m in diameter with 50 g/cm2 thickness of regolith. Analyzing the required thickness for a habitat of Mars is basically similar, although the radiation design envi- ronment will be slightly different, the PELs will be lower (since the overall mission is longer), and the regolith has a different composition. There is also one additional step: modeling the propagation of the radiation through Mars’s atmosphere, which offers a substantial degree of dose reduction. (More detail on this method can be found in Simonsen [1997]). Simonsen points out that for short stays on the martian surface, it may make sense not to use any regolith shielding at all, because the construction would add length and risk to the mission, negating the additional protection of the shielding. The process detailed above will generate a design that meets standards, but not one that satisfies the ALARA principle. Since regolith is abundant, the goal might be “the thicker, the better.” As thicknesses are increased to exceed a few meters, a practical limit is reached, however, because of structural load factors even in the low lunar gravity, the diminishing marginal shielding associated with large thickness, and the efforts and safety risks for construction. Implementation becomes a matter of the economies of scale. Substantial masses of equipment will be required to excavate and move surface material; to sort, crush, or condense it for packaging; to package it by pressing it into bricks or packing it into bags; to move these packaged units to the construction site; and finally to build the protective structure over the habitat. Because astronaut time is expensive and the crew will have many higher priorities than “filling and stacking sandbags,” this construction will most likely require automated and teleoperated machinery to extract and process regolith, and robotic sys- tems to move and emplace it. The equipment might resemble skid-steer loaders or compact excavators but would be designed for operation in a low-gravity space environment. It will weigh at least several metric tonsa very significant payload for the Lunar Landerand also require significant amounts of power from the solar-power generators, which are also large payloads. Thus, the economy of scale becomes a system tradeoff of how much of this equipment it is feasible to deliver to the lunar surface and when to deliver it. For example, if there is not sufficient power generation and transmission capability in place yet, it does not make sense to deliver the regolith- handling equipment that will exceed the available capacity. For sortie missions, it is likely that regolith shielding will not be practical. However, as missions lengthen, the utility of the shielding increases. Furthermore, constructing an outpost means that the cost of shielding construction is amortized over the duration of the program. This will be true of either sculpted regolith or of shielding that has been brought from Earth but left behind when the Lunar Lander departs. Geomorphology-Dependent Shielding Lava tubes, caves, tunnel borings, and other shielding solutions based on existing lunar formations offer the potential to reduce but probably not eliminate the need for heavy construction equipment. Lava tubes are believed to exist in the maria, where the rills appear to be collapsed lava tubes. Caves may exist in crater walls or tumbled boulder fields. In the most basic scenario, the outpost would be located next to a steep hill or cliff (perhaps the side of a crater), cutting off some portion of the line of sight. In the ideal scenario, explorers would find a lava tube of sufficient size and structural stability to contain the outpost and its proximity operations. Theoretically, missions could include equipment to dig down into the regolith instead of piling it on top of a habitat. The resulting “shield thickness” would be very large, but such construction is more energy-intensive and complex. This concept is not being currently considered within the scope of Project Constellation.

OCR for page 62
68 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION Extravehicular Activity Space Suits An astronaut performing an EVA is primarily shielded by the planetary body on which he or she is standing. Clowdsley et al. (2005) calculated a point estimate that the maximum daily effective GCR dose for an astronaut in an EVA suit exposed on the lunar surface is 0.085 cSv (about half of the calculated amount for free space). The 1977 solar minimum environment was used as a worst-case GCR environment. This estimate neglected the minimal radiation protection capabilities of the EVA suit itself. Further protection can be added through operational plan- ning, the use of portable shelters, and possibly pressurized rovers. Space suits are not intended to provide significant protection. Since mass is tightly constrained during space suit design, any mass taken up solely by shielding could otherwise be used to store more oxygen, enable a thicker thermal shield, or make the suit less cumbersome. Like many parts of Project Constellation, space suits are subject to risk leveling: finding that an extra kilogram is avail- able, a designer will figure out how that kilogram could be used to provide the greatest reduction in risk. However, space suits do also provide some minimal degree of radiation protection. Therefore, it is useful to quantify their shielding properties and to incorporate the ALARA principle into their design (Wilson et al., 2006). Current space suit concepts encompass three main types: the soft suit, the hard suit, and the hybrid suit. Each of these three broad categories covers a range of advantages and disadvantages for general and specific applica- tions and criteria. • Soft suit: a space suit in which the only rigid component is typically the helmet and perhaps the soles of the boots. The potential of the soft suit as a tight-fitting pressure envelope is to minimize mass while maximizing flexibility. The Mercury, Gemini, and Apollo suits were all soft suits. • Hard suit: a space suit in which all the components of the pressure enclosure are made from a rigid torso, joints, spacers, and sizing rings. The pressure-sealed joints turn on circular ball bearing races. Although NASA has explored a few concepts, they have proven too heavy, uncomfortable, and expensive to be used on missions. A hard suit offers the potential for the fabrication of its parts from carbon composite, a better material for radiation shielding than aluminum or fiberglass is. • Hybrid suit: a space suit that combines hard and soft components, typically a hard upper torso and the rest soft goods. The space shuttle and ISS extravehicular mobility units (EMUs) are hybrid suits. Other features are generally standard across suit design, updated as each model incorporates the technology available at the time. Helmet visors are coated to protect the astronauts’ eyes from certain wavelengths of harmful light; they also reduce the dose of ionizing radiation to the eyes, therefore also lowering the risk of cataracts. The personal life support system (PLSS) backpack contains oxygen, water, and batteries. It provides a fair amount of radiation shielding over a limited area through its mass. Gloves are usually fairly thin, given the need for dexterity and manipulation. The layer of material closest to the skin is the liquid cooling and ventilation garment (LCVG). It supports a network of small, water-filled tubes used for convective cooling. The LCVG is contained by a bladder and restraint to maintain the suit’s pressure. The outermost layer is the thermal micrometeoroid garment (TMG), which actually consists of several layers of insulation and ripstop fabric. In 2003, NASA published a collection of radiation studies that had been performed on the EMU and the Orlan-M, the space suit used by the Russians on the ISS (Cucinotta et al., 2003). These studies included beam- line experiments using electrons with an energy (6 MeV) typical of those in Earth’s Van Allen radiation belts and protons with energies representative of SPEs and Van Allen belts (up to 232 MeV). Validation of the BRYNTRN space radiation transport code, which was developed at NASA Langley Research Center and typically used for estimating doses from SPE and Van Allen protons was also carried out. Some relevant findings from this group of studies included the following:

OCR for page 62
69 SHIELDING APPROACHES AND CAPABILITIES • When doses were measured behind swatches of space suit material for a beam of 60 MeV protons, secondary radiation caused the dose to actually increase by 10 to 40 percent for the suits and helmets owing to the increased slowing down of the protons, resulting in greater energy deposition rates (Benton et al., 2003). • Beam-line experiments were also performed on a phantom torso (a human mock-up, designed to provide the same shielding as a human body; it can be outfitted with dosimeters at various organ locations). The proton beam energy was set to 232 MeV in order to penetrate the torso. The torso was instrumented with TLD-600 and TLD-700 dosimeters, positioned to measure doses and dose equivalents in the eye, brain, lung, stomach, and thigh. The eye and brain were shielded by the helmets; the lung, stomach, and thigh were shielded by the suit fabric. Overall, the helmets were found to reduce the doses by 22 to 27 percent for the eye and by 13 to 21 percent for the brain. Dose-equivalent reductions ranged from 14 to 25 percent for the eye and less than ~8 percent for the brain. For the lung, stomach, and thigh, dose reductions ranged from <1 percent (thigh) to 13 percent (stomach). Dose-equivalent reductions ranged from 0 percent (thigh) to 14 percent (lung) (Benton et al., 2003). • The BRYNTRN radiation transport code was found to model proton exposures accurately when compared with measured doses taken in a beam line (Zapp et al., 2003). Typical differences between calculated and measured doses were less than 6 percent for the various organs. • Space suits have components that do provide some measure of radiation protection. Certain minor changes in the EMU design, for example, adding a thin layer of material to the dorsal side of the glove, could yield gains in radiation protection (Moyers et al., 2003). Note that the beam experiments involved protons and electrons at energies similar to those found in LEO. Constellation astronauts will have to pass through these regions, but the bulk of their time will be spent in deep space or on a lunar or planetary surface, where the contributions of electrons are negligible but GCR and SPEs play a much more prominent role. Despite the differences, these measurements, methods, and computational models can serve as a starting point for evaluating new space suit designs. Wilson et al. (2006) suggest some other possible design changes inspired by the analysis of ISS suits: replacing the water tubes of the LCVG with a solid water jacket, using a thin layer of polyethylene as part of the suit layup, and redesigning the PLSS so that the mass of its components are more evenly distributed around the body. The Johnson Space Center EVA Systems Project Office issued an EVA Systems Architecture and Reference Suit System Approach briefing (Dutton and Johnson, 2006) that has been approved by the Constellation program. Like the EMU, the Constellation suit will be fairly modular; it will have two configurations: an emergency suit (for launch, entry, and abort and for contingency EVA) and a sortie EVA suit. These two suits will share many components, although they will have different soft upper torsos, TMGs, and visors. A third configuration may eventually be designed for outpost EVA; it will use many of the same components. The Constellation TMG will evolve out of the TMGs used in the Apollo program, substituting modern, advanced materials such as hydrogen-rich radiation shielding. The TMG will have to be redesigned for martian missions owing to atmospheric requirements. During the design of the EMU, materials with radiation shielding properties were used, but certification only required that the materials could stand up to the environment (e.g., 50 EVAs), not that they would guarantee a certain amount of protection. In addition, the visors of the helmet will be subject to a variety of requirements on transmittance and reflectance of certain wavelengths of light. Although dose limits for the Constellation suits have not yet been set, designers are expecting them to be similar (S. Cupples, NASA, “EVA Suit Radiation Attenuation,” presentation to the committee on February 21, 2007). The setting of requirements, and the efforts of the space suit designers to incorporate radiation shielding early in the process are both very positive signs. Airlock The Exploration Systems Architecture Study (NASA, 2005) and request for information (NASA, 2006) both include an EVA airlock to enable the safest and most efficient means of transit between the pressurized cabin of the lander or habitat and the vacuum of space. The EVA airlock may offer an alternative module for a radiation storm shelter, particularly if NASA implements a requirement for a hyperbaric capability, such as on the ISS (Barratt,

OCR for page 62
70 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION 1996). In this case, the airlock could be operated at 6 atmospheres, allowing the airlock to serve as a recompres- sion chamber in case of decompression sickness. An airlock that holds 6 atmospheres (and will be pressure-tested to 12) will be significantly more robust and thicker in its primary structure than one that operates at 1 atmosphere and tests to 2. This thicker structure and material would afford a de facto shielded module. However, outfitting the EVA airlock to serve as a solar storm shelter poses some significant penalties. The most obvious is that a primary purpose of the airlock is to provide a buffer for the living quarters to exclude, mitigate, and control the intrusion of lunar dust, which the Apollo astronauts found highly irritating for breathing and on the skin, and which also may prove toxic. Secondly, the airlock would become cramped during an SPE lasting several days. A crew taking shelter in the airlock would need to bring their sleeping arrangements from the habitat with them into the airlock, as well as food, water, first aid, laptops, and other necessities. The crew would need to leave the airlock to use the hygiene facilities in the habitat. Because of these drawbacks, it may be preferable to provide some shielding in the habitat so that the airlock would only need to be used during the highest flux peak of the event. Portable Shielding The purpose of a portable radiation shield would be to provide temporary shielding in a contingency situationfor example, the sudden onset of an SPE. The portable radiation shield is envisioned to be a rapidly deployed and easily transported “blanket-design” made of a lightweight material that exhibits good radiation- mitigating properties, such as high-density polyethylene. The “blanket” could be transported on a lunar rover or a multipurpose trailer. In the event of an SPE, it would be advisable for the EVA crews to seek out terrain that would provide additional shielding, such as a high rock or cliff formation. It is also envisioned that mission controllers would direct the crew when to seek shelter and to use the portable shielding, and also when an “all-clear” period was reached and it was permissible to continue the EVA or to return to a base location. Another design could be for individual use, such as a hooded, poncho-type shield arrangement that the crew member could slip into quickly. The thickness of these portable shields would be dictated by how much radiation mitigation would be warranted. Concept of Operations Operational protocols will be critical to radiation safety during any EVA beyond the protective influence of Earth’s geomagnetic shield. These protocols will involve gathering two kinds of data: solar activity indicators and actual radia- tion dose as experienced by the astronauts conducting the EVA. The second can only be accomplished by including one or more active dosimeters on their persons or on nearby equipment, such as a rover or robotic assistant. When solar activity is observed that could result in an SPE, the EVA will likely have specific contingency plans already in place for seeking shelter or for using available shielding, much as there are contingency procedures for anomalous launch or other hazardous operational situations. During EVA, access to shielding on a timescale of less than 1 hour, and preferably only minutes, is desired in order to avoid possible excessive exposure. For the most expedient mitigation, astronauts might use a two-tiered approach. Dose rate will build up over a period of time, so that immediately taking on even a small amount of strategic shielding (e.g., covering the torso) prior to reaching a more substantial shelter area would be effective in reducing the total exposure received. Nuclear Power Space missions usually rely on solar panels and/or fuel cells to provide electricity. However, as anticipated by NASA, the future long-term missions to the Moon and Mars would require a power source on the order of 1 MW that can reliably supply electricity and provide heat for a few decades. Nuclear reactors are the most efficient method of generating this power (Figure 4-1). Presented here is a discussion of the engineering challenges asso- ciated with the safe operation of fission reactors in space. Not only must the reactors have adequate shielding to protect astronauts during nominal operation, but they must have enhanced safety and reliability as well.

OCR for page 62
71 SHIELDING APPROACHES AND CAPABILITIES FIGURE 4-1 The relative applicability of various space-based sources of electrical power. SOURCE: NRC, 2006. R01155, Figure 4-1 Fixed image, not changeable Under the present plan for the lunar architecture (Lavoie, 2006), the Shackleton Crater Rim is selected as the outpost site location. The average monthly illumination at the outpost location is in excess of 70 percent, which would allow the outpost to be powered entirely by solar-power generation. The Solar Standard Power Unit will generate 10 kW of daytime power and 2 kW of eclipse power. In combination with two Make-up Power Units, the system will deliver 6 kW of continuous day/night power. The Make-up Power Unit primarily includes a self- contained proton exchange membrane regenerative fuel cell subsystem with gaseous H 2/O2 storage. The Prometheus Power and Propulsion program, originally tasked to investigate nuclear propulsion for long- duration missions to the outer planets, was restructured to support the long-duration stays on lunar and martian surfaces called for by the Vision for Space Exploration. This program’s primary focus is nuclear fission surface power (FSP) systems. The FSP is the most credible solution for power generation at levels of tens of kilowatts to support lunar surface operation independent of day/night cycle or lunar surface location. The FSP is similarly enabling for the poorly illuminated surface of Mars. The FSP module contains main components that are intended to generate and at the same time to control the thermal power and reactor-induced radiation for the lunar or martian habitat. These components include the following: • Reactor core and reflectors, • Primary heat transfer, • Radiation shield, • Reactor instrumentation and control, and • Power conversion and radiator. According to NASA’s investigations and missions planning, the power requirements for human-tended outposts are projected to range from 25 to 100 kW in the early stages, with the requirements approaching 1 MWe following

OCR for page 62
72 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION the lunar base development. The power level and life for the current reference design of each FPS are 40 kW and 8 years, respectively. The FPS radiation shield component provides the attenuation of the reactor-induced radiation to the rest of the FPS module (J. Nainiger and L. Mason, NASA, “Plans for Affordable Surface Fission Power,” presentation to the committee on December 12, 2006). The radiation protection measures and systems designs for fission power on Earth are well known. However, when applied to remote planetary destinations, direct transfer of the existing technology becomes difficult, if not impossible. The heavy shielding used on Earth is too massive to transport to the Moon. Unlike terrestrial radiation workers, astronauts will have to live within a certain range of the reactors. As noted in Chapter 2, solving these problems will require engineering solutions rather than more research. NASA is currently considering two fission surface power system (nuclear reactor) shielding approaches: • Emplaced configuration: The light-material-shielded reactor will be placed in an excavated cavity so that external shielding is provided by lunar or martian regolith (Figure 4-2A). Unless a convenient crater can be located, this will require astronauts to be involved in excavation, construction, and the installation of the reactor. The overall effect of using regolith as a shield is a reduction of the neutron and gamma-ray doses from the reactor by five and three orders of magnitude, respectively. (See also the discussion in the subsection “In Situ Shielding” earlier in this chapter.) • Landed configuration: A reactor equipped with an above-grade, 4π shield made of Earth-delivered materials will be placed on the regolith surface. The landed configuration requires very large mass and volume of high-Z and low-Z materials to reduce the neutron and gamma-ray dose to the levels equivalent to those provided by underground emplacement of the reactor module (Figure 4-2B). The main sources of radiation are neutron and gamma rays from the fissions in the reactor, prompt gamma emission from neutron interactions with the regolith, and gamma rays emitted by the decay of activated regolith and other external shielding materials. The activation of the soil near the reactor will reach near-equilibrium during the 10- to 20-year lifetime of the surface station. The activity following reactor shutdown will be dominated by Na24, with a half-life of roughly 15 hours; soil activation will decrease within days. For the reference below-grade excavated configuration, the total radiation dose is less than 5 cSv/yr at a radial distance of 100 m from the reactor axis. For the above-grade configuration, the 4π shield reduces the dose at the habitat area that is 1 km away from the reactor axis to less than 5 cSv/yr, and for nonhabitat areas to less than 50 cSv/yr (J. Nainiger and L. Mason, NASA, “Plans for Affordable Surface Fission Power,” presentation to the committee on December 12, 2006). The radiation rate from the reactor core is dependent on its power. The power generation expected for this mission is easily attainable with current technology; therefore, other aspects could be optimized: the weight and type of the shielding materials and the distance from the habitats to the reactor. Optimization methodologies and computer simulation tools needed for this sort of analysis are already available. On Earth, where weight is a much lower concern, there has been limited attention to the development of very lightweight shielding. On a mission to the Moon, NASA will be limited to available lightweight shielding materials or materials from the lunar surface. Engineering challenges in the development of the fission surface power system design studies and technology demonstrations include the following: • Fission surface power system and balance-of-plant design: Conceptual studies that would lead toward technology demonstration would require analysis of the reactor design, technology selection, and material shield design development and demonstration. • Micrograity: Analysis of microgravity effects would need to be carried out, with selected experiments to verify the operational safety of the ground fission power systemfor example, effects on the flow of coolant through pipes. • Autonomous operation: Lunar surface power systems cannot be constantly attended to and thus they will require more complex, robust, self-contained control systems.

OCR for page 62
73 SHIELDING APPROACHES AND CAPABILITIES A π B FIGURE 4-2 (A) Reactor power module in excavated cavity (emplaced configuration). (B) Reactor power module on landed configuration with external 4π shield. NOTE: This is a conceptual figure for illustrative purposes only. It does not imply a finalized design, nor the committee’s endorsement of the concept pictured. SOURCE: J. Nainiger and L. Mason, NASA, “Plans R01155, Figure 4-2 for Affordable Surface Fission Power,” presentation to the committee on December 12, 2006. Each half is a fixed image, not changeable To print in color • Shielding:  hat is the optimized lightweight shielding design for the expected range of the fission surface power W systems?  hat is a safe distance from the fission surface power systems to ensure acceptable annual exposure, W including the variations of power and shielding from the base-design when added to the natural radiation exposure on the lunar surface?

OCR for page 62
74 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION  What course will be taken in case of accidents in the presence of the crew? What would be the expected exposures to the crew? Would crew intervention be required, and what additional dose would that confer? Would these considerations necessitate additional shielding? — For the case of accidents not in the presence of the crew, once the crew returns, what action is required and what dose would that involve? — What would be done with the reactor core at the end of life? What shielding requirements would that pose and what dose levels may be expected? Further analyses are needed to understand the best policy on planetary and space pollution with radioactive materials. • Launching criteria, requirements, and technology: NASA has never launched a nuclear reactor into space. The risks associated with launching a reactor, its transport and extreme environmental constraints on the Moon and Mars can be addressed through research in the following areas.  Launch payload: optimization of weight and size regarding the shielding.  Hazardous material safety consideration: safety protocols.  Operational lifetime: continuous power generation for life support and shielding material durability.  Environmental constraints and effect on shielding: safe operation with intact shielding.  Shielding in case of accidents: “breakup” during launch or reentry; water-immersion accident in which the reactor falls into water and becomes moderated; compression in the case of an impact; individual component failures. NUCLEAR PROPULSION NASA has not yet decided on an approach to providing propulsion to Mars, so it is unknown if nuclear pro- pulsion will be employed. Therefore, this section provides only an overview of some of the pluses and minuses associated with the topic. Nuclear propulsion includes a broad range of propulsion methods that use nuclear reaction as a primary source of power. Nuclear propulsion is usually divided into two categories, depending on the source of propulsion. 1. Radioisotope decay has been a main source of electricity and propulsion for numerous missions in the past four decades. Russia has flown more than 30 reactors and numerous radioisotope systems. The United States flew the SNAP-10A in 1965 and has flown dozens of radioisotope systems (most recently on the New Horizons mission to Pluto). Energy is released from the radioactive decay of the selected radioisotope (in most cases plutonium-238) providing a continuous, reliable, and proven source of power. The energy released in radioactive decay of the isotope in the device called a radioisotope thermoelectric generator (RTG) is converted into electrical power that is used to accelerate and eject propellant at a speed high enough to boost the spacecraft. The major limitation of such a system is its low power. RTG is usually applied for long-term (i.e., low-acceleration) missions or missions to the outer planets, where solar power is inadequate. This approach currently represents the safest space nuclear propulsion system with the lowest requirements for the shielding, as long as the container holding the radioisotope does not leak. The environmental impact study for the Cassini-Huygens probe launched in 1997 (NASA, 1997) for Saturn estimated the probability of contamination accidents at various stages in the mission. For example, the probability of an accident to cause radioactive release from one or more of its three RTGs during the first 3.5 minutes following launch was estimated at 1 in 1,400; the probability of a release later in the ascent into orbit was 1 in 476; after that, the likelihood of an accidental release fell off sharply to less than 1 in 1 million. 2. Thermal nuclear propulsion in submarines is an industry with developed, proven technology and more than 50 years of accumulated knowledge and experience. However, there has not yet been an analogous system that propelled spacecraft into space. There have been ideas to develop nuclear power in space since the earliest days of the space program. A nuclear thermal rocket and chemical rockets operate by the same basic principlesnamely, the expansion of hot gas (propellant) through a rocket nozzle to provide thrust. The propellant flows through cool- ant channels of the solid-fuel reactor core where it is heated to very high temperatures (>3,000 K proposed for pseudo-ternary carbides). To achieve high performance, the fuel is required to operate at very high temperatures. Hydrogen has been used as a propellant during all rocket reactor tests and is preferred because it has the lowest

OCR for page 62
75 SHIELDING APPROACHES AND CAPABILITIES molecular weight. However, hot hydrogen can react with the fuel, resulting in corrosion and mass loss. Furthermore, mission cost constraints require a compact, lightweight reactor necessitating high power densities (high neutron flux) with associated radiation damage and increased susceptibility to fracture. Because of its high performance potential, nuclear thermal propulsion could be used for human explora- tion missions and cargo transport to the Moon or Mars, for outer-planet robotic explorations, and for Earth-orbit transfers of satellites. The main benefit of nuclear propulsion is that it can provide a greater specific impulse, or the amount of thrust provided per unit mass of propellant. Producing the same thrust with less required fuel creates two complementary possibilities. First, one could have a spacecraft of the same weight but with more shielding. Second, one could have a lighter spacecraft, with a higher velocity that could reduce the transit time and radia- tion dose. In reducing mission length, nuclear propulsion will reduce the risk to astronauts from cosmic radiation in addition to the other health, psychological, and operational benefits associated with shorter mission durations. Although this would be offset in part by the radiation of the reactor itself, Nealy et al. (1991) found that cosmic radiation was more dominant. It is estimated that nuclear propulsion could reduce the length of a short human- exploration mission to Mars from a year and a half (using chemical propulsion) to under a year (NASA, 1989). Nuclear propulsion could also reduce the total time of a longer-duration mission by 50 to 100 days for the low- Earth-orbit mission (Bennett and Miller, 1991). A comparison of two similar missions is shown in Figure 4-3; in addition to a shorter mission length, the nuclear mission includes a longer stay on Mars, which raises the value of the mission as well as reducing the total radiation exposure. Figure 4-4 shows the tradeoffs between mission masses and travel times for a long-duration mission launching between 2008 and 2011. ALTERNATIVE METHODS Active Shielding The use of active shielding that involves the generation of an electromagnetic field to deflect or capture the incoming radiation particles and thereby protect the crew has been proposed. However, there are many practical and technical difficulties with implementing such a method of defense. First, the active shield may require large amounts of power. Second, such powerful electromagnetic fields can have disruptive effects upon the onboard microelectronics. Third, possible exposure of the crew to such powerful electromagnetic fields may carry addi- tional unknown health effects. There are also technological challenges that need to be addressed including the safe dissipation of the stored energy in the event of a spontaneous quench of the magnet, refrigeration of the coils to 2.1 K on a long space mission, and the development of new superconducting cables. Electromagnetic shields often seem attractive when compared with bulk shielding. However, a true comparison between the efficacy of the two must include the following: 1. Complete and realistic representations of radiation environments expected in deep space (SPEs and GCR), which consist of particle types ranging from protons through iron nuclei with energies up to several GeV per nucleon. It is not sufficient to choose one particle type and energy for the analyses and then base the conclusions of the study on that limited environment. 2. It is not possible to deflect all of the incident particles, since many will have energies above the cutoff energy of the electromagnetic field. Just as it is not possible to stop all of the particles in a bulk shield; some par- ticles will be transmitted through both types of shielding. Estimates of the shield effectiveness must include detailed calculations of the biological risks from those transmitted particles, both primary and secondary. Thus, dose and dose-equivalent comparisons between active and passive shields should include detailed transport analyses that consider all relevant secondary particle production mechanisms. Estimates of bulk shield mass requirements based on simple range-energy relationships are not adequate for choosing between passive bulk shielding and proposed alternative shield configurations.

OCR for page 62
76 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION FIGURE 4-3 Mars mission duration for chemical and nuclear propulsion. SOURCE: Stafford, 1991. Finally, there is a difficulty that is ubiquitous to the lunar surface: lunar dust. The lunar dust contains nanophase R01155, Figure 4-3 iron inclusions that give it electromagnetic properties. Creating a powerful electromagnetic field could loft dust over the outpost. The Apollo crews foundhalfdusta fixed image, the skin and respiratory system during their Each the is to be irritating to not changeable short 2- to 3-day stays on the lunar surface. Since the dust is highly abrasive to machinery and probably hazardous to the crew’s health, levitating the dust could create additional threats to the crew and mission.

OCR for page 62
77 SHIELDING APPROACHES AND CAPABILITIES Total mission mass in low Earth orbit (metric tons) One-way transit times to and from Mars (days) FIGURE 4-4 Comparison of transit times for a long-duration Mars mission using either a chemical or a nuclear propulsion system. Shorter-duration missions require more propellant and thus a higher mission mass. SOURCE: Stafford, 1991. R01155, Figure 4-4 Fixed image, not changeable, Radioprotectants except x- and y- axis labels The only radioprotectant that has been approved for use in humans at this time is amifostine (ethyol), which has a dose-reduction factor (ratio of dose of radiation to cause an effect in the presence of the drug to the dose of radiation to cause the same effect in the absence of the drug) of 1.8 to 2.7 when the drug is administered prior to an exposure. The dose-reduction factor varies with the amount of time that has passed since exposure. This drug was originally developed by the Walter Reed Institute (and called WR2721) as a thiol-containing compound with free-radical scavenging capabilities (Hall and Giaccia, 2006). Amifostine is used clinically to prevent late tissue toxicities (such as salivary gland damage) in patients who are being treated for head and neck cancers. The drug has been available to NASA for several decades, predominantly because of its ability to inhibit the induction of mutations following radiation exposure, thus suggesting the possibility of reducing stochastic effects (cancer) following radiation exposure. There is some confusion in the literature on the ability of amifostine to protect against damage from high-LET (linear-energy-transfer) types of radiation such as those encountered in space (Hall and Giaccia, 2006). On the one hand, the free-radical scavenging activity of amifostine should allow it to protect against indirect damage to DNA (induced by free radicals, predominantly following low-LET expo- sures) rather than direct damage to DNA, which predominates following high-LET exposures. On the other hand, there are several studies demonstrating that amifostine reduces cancer induction in mice when administered prior to exposure either to low-LET gamma rays or to high-LET neutrons (Grdina et al., 2002a,b), although the mecha- nism is still to be elucidated (Grdina et al., 2000). Nevertheless, its use as a chemoprevention agent is limited by toxicity, which includes hypotension. Various other radioprotectors have been examined in the literature, including phosphanol, Mn-SOD mimetic drugs, and others (Greenberger and Epperly, 2007). None has been shown to have the same level or as broad a level of protection as amifostine. During the past 2 or 3 years, in an effort to protect against possible radiation attacks coming from terrorists and others, the National Institutes of Health has invested in several large center grants to develop a better understanding of radiation effects, including the development of new protectors. NASA participates, along with other agencies, in the continuing evaluation of these developments, and it is well placed to take advantage of any breakthroughs in radioprotectant developments as they occur.

OCR for page 62
78 MANAGING SPACE RADIATION RISK IN THE NEW ERA OF SPACE EXPLORATION SUMMARY Finding 4-1. State of radiation protection plans for lunar missions. The use of surface habitat and spacecraft structure and components, provisions for emergency radiation shelters, implementation of active and passive dosimetry, the scheduling of EVA operations, and proper consideration of the ALARA principle are strategies that are currently being considered for the Constellation program. These strategies, if implemented, are adequate for meeting the radiation protection requirements for short-term lunar missions. Recommendation 4-1. Strategic design of Orion. As the design of Orion continues to evolve, designers should continue to consider and implement radiation protection strategies. Finding 4-2. State of radiation protection plans for Mars missions. For longer-duration lunar and Mars missions the currently large uncertainties in radiological risk predictions could be reduced by future research. Without such research, it may be necessary to baseline large shielding masses and reduced-length missions, and/or delay human exploration missions until uncertainties in risk prediction and radiobiological methods of risk management have advanced to the point that they can be conducted within the limits of acceptable risk. REFERENCES Anderson, M., A. Hanford, R. Howard, and L. Toups. 2006. Lunar surface scenarios: Habitation and life support systems for a pressurized rover. P. 84 in Earth & Space 2006: Proceedings of the Tenth Biennial ASCE Aerospace Diision International Conference on Engineering, Construction, and Operations in Challenging Enironments, eds. R.B. Malla, W.K. Binienda, and A.K. Maji. American Society of Civil Engineers, Reston, Va. Barratt, M.R. 1996. Space Station Hyperbaric Medicine Ad Hoc Committee Meeting. NASA-CP-10140. Available at http://ntrs. nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19960014811_1996009073.pdf. Bennett, G., and T. Miller. 1991. Planning for the space exploration initiative: The nuclear propulsion option. Pp. 1-4 in Proceed- ings of the Eighth Symposium on Space Nuclear Power Systems, eds. M.S. El-Genk, and M.D. Hoover. AIP Conference Proceedings 217. American Institute of Physics, New York. Benton, E.R., E.V. Benton, and A.L. Frank. 2003. Characterization of the radiation shielding properties of U. S. and Russian extra- vehicular activity suits. Pp. 71-96 in Radiation Protection Studies of International Space Station Extraehicular Actiity Space Suits, eds. F. Cucinotta, M. Shavers, P. Saganti, and J.Miller. NASA/TP-2003-212051. NASA, Washington, D.C. Clowdsley, M.S., J.E. Nealy, J.W. Wilson, B.M. Anderson, M.S. Anderson, and S.A. Krizan. 2005. Radiation protection for lunar mission scenarios. Space 2005. Paper No. AIAA 2005-6652. American Institute of Aeronautics and Astronautics, Reston, Va. Cucinotta, F.A., M.R. Shavers, P.B. Saganti, and J. Miller. 2003. Radiation Protection Studies of International Space Station Extraehicular Actiity Space Suits. NASA/TP-2003-212051. NASA, Washington, D.C. Dutton, J., and B. Johnson. 2006. Constellation Space Suit System. CSSS Overview. Operational Assumptions/Concepts. CSSS Industry Day presentation, NASA Johnson Space Center, Houston, Texas, August 25. Available at http://procurement.jsc. nasa.gov/csss/Suit%20Systems.ppt. Grdina D.J., Y. Kataoka, and J.S. Murley. 2000. Amifostine: Mechanisms of action underlying cytoprotection and chemo- prevention. Drug Metabolism and Drug Interactions 16(4):237-279. Grdina D.J., Y. Kataoka, J.S. Murley, N. Hunter, R.R. Weichselbaum, and L. Milas. 2002a. Inhibition of spontaneous metastases formation by amifostine. International Journal of Cancer 97(2):135-141. Grdina, D.J., J.S. Murley, Y. Kataoka, and W. Epperly. 2002b. Relationships between cytoprotection and mutation prevention by WR-1065. Military Medicine 167(2 Suppl.):51-53. Greenberger, J.S., and M.W. Epperly. 2007. Review. Antioxidant gene therapeutic approaches to normal tissue radioprotection and tumor radiosensitization. In Vio 21(2):141-146. Hall, E., and A. Giaccia. 2006. Radiobiology for the Radiologist. 6th Edition. Lippincott Wilkins and Williams, Philadelphia, Pa. Lavoie, T. 2006. Lunar Architecture Team Overview. Presented at NASA-AIAA 2nd Space Exploration Conference, Houston, Tex., December 4-6. Slides and Webcast available at http://www.nasa.gov/mission_pages/exploration/main/ 2nd_exploration_conf_prt.htm.

OCR for page 62
79 SHIELDING APPROACHES AND CAPABILITIES Moyers, M.F., G.D. Nelson, and P.B. Saganti. 2003. Energy measurements for extravehicular activity space suits. Pp. 19-53 in Radiation Protection Studies of International Space Station Extraehicular Actiity Space Suits, eds. F. Cucinotta, M. Shavers, P. Saganti, and J. Miller. NASA/TP-2003-212051. NASA, Washington, D.C. NASA (National Aeronautics and Space Administration). 1989. 90-Day Study on Human Exploration of the Moon and Mars. Washington, D.C. November. NASA. 1997. Environmental impacts. Chapter 4 in Cassini Final Supplemental Enironmental Impact Statement. Available at http://saturn.jpl.nasa.gov/spacecraft/safety-eis.cfm. NASA. 2004. Crew Exploration Vehicle (CEV) Request for Proposals. Available at http://prod.nais.nasa.gov/cgi-bin/eps/sol. cgi?acqid=113638. NASA. 2005. NASA’s Exploration Systems Architecture Study. NASA-TM-2005-214062. NASA, Washington, D.C. NASA. 2006. Lunar Lander Concepts Requirements Description. Available at http://procurement.jsc.nasa.gov/NNJ06LSAM05L/ LanderStudyRFI.doc. Nealy, J.E., L.C. Simonsen, J.W. Wilson, L.W. Townsend, G.D. Qualls, B.G. Schnitzler, and M.M. Gates. 1991. Radiation exposure and dose estimates for a nuclear-powered manned Mars SPRINT mission. American Institute of Physics Confer- ence Proceedings 217(2):531-536. NRC (National Research Council). 2006. Priorities in Space Science Enabled by Nuclear Power and Propulsion. The National Academies Press, Washington, D.C. Simonsen, L.C. 1997. Analysis of lunar and Mars habitation modules for the space exploration initiative. Pp. 43-77 in Shielding Strategies for Human Space Exploration, eds. J.W. Wilson, J. Miller, A. Konradi, and F.A. Cucinotta. NASA Conference Publication 3360. Washington, D.C. Stafford, T. 1991. America at the Threshold: Report of the Synthesis Group on America’s Space Exploration Initiatie. Report of the Synthesis Group On America’s Space Exploration Initiative. U.S. Government Printing Office, Washington, D.C. Wilson, J.W., B.M. Anderson, F.A. Cucinotta, J. Ware, and C.J. Zeitlin. 2006. Spacesuit Radiation Shield Design Methods. Paper presented at the International Conference on Environmental Systems, Norfolk, Va., July. Document ID No. 20060046504. Available from NASA Center for AeroSpace Information, Hanover, Md. Zapp, N., E. Semones, and P. Saganti. 2003. A comparison of model calculation and measurement of absorbed dose for proton irradiation. Pp. 97-104 in Radiation Protection Studies of International Space Station Extraehicular Actiity Space Suits, eds. F. Cucinotta, M. Shavers, P. Saganti, and J. Miller. NASA/TP-2003-212051. NASA, Washington, D.C. Zeitlin, C., L. Heilbronn, J. Miller, and J.W. Wilson. 2002. Accelerator-Based Validation of Shielding Codes. Paper LBNL 51264. Lawrence Berkeley National Laboratory, Berkeley, Calif. Available at http://repositories.cdlib.org/cgi/viewcontent. cgi?article=1483&context=lbnl.