MASTER's THESIS Space Radiation Analysis
2009:107 MASTER'S THESIS
Space Radiation Analysis - Radiation Effects and Particle Interaction outside Earth Magnetosphere using GRAS and GEANT4
Lisandro Martinez
Luleå University of Technology Master Thesis, Continuation Courses Space Science and Technology Department of Space Science, Kiruna
2009:107 - ISSN: 1653-0187 - ISRN: LTU-PB-EX--09/107--SE
Space Radiation Analysis: Radiation Effects and Particle Interaction outside Earth Magnetosphere using GRAS and GEANT4
Master’s Thesis For the degree of Master of Science in Space Science and Technology
Lisandro M. Martinez Luleå University of Technology Cranfield University June 2009
Supervisor: Johnny Ejemalm Luleå University of Technology
June 12, 2009 MASTER’S THESIS
ABSTRACT
Detailed analyses of galactic cosmic rays (GCR), solar proton events (SPE), and solar fluence effects have been conducted using SPENVIS and CREME96 data files for particle flux outside the Earth’s magnetosphere. The simulation was conducted using GRAS, a European Space Agency (ESA) software based on GEANT4. Dose, dose equivalent and equivalent dose have been calculated as well as secondary particle effects and GCR energy spectrum.
The results are based on geometrical models created to represent the International Space Station (ISS) structure and the TransHab structure. The physics models used are included in GEANT4 and validation was conducted to validate the data. The Bertini cascade model was used to simulate the hadronic reactions as well as the GRAS standard electromagnetic package to simulate the electromagnetic effects.
The calculated total dose effects, equivalent dose and dose equivalent indicate the risk and effects that space radiation could have on the crew, large amounts of radiation are expected to be obtained by the crew according to the results. The shielding comparison between ISS and TransHab indicate that a tradeoff between the two will have to be made, since the first has a higher protective ratio compared to the TransHab; on the other hand the second one is more flexible and could eventually become a larger structure. The GCRs effects upon the structure are found to be comparable to experimental data.
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Acknowledgements
The work presented in this thesis could not have been done without the help, support, and participation of many people as well as institutions.
First, I would like to thank my advisor, Dr. Jennifer Kingston, for her constant support during the development of this thesis. Her interest in the subject as well as her good will was always reassuring during this work.
I am grateful to the Space Master consortium for their support, not only during this work but also for the past year and a half of outstanding education, especially to the staff in Kiruna who have been very helpful. To the ErasmusMundus scholarship, that financed my dual MSc education.
I would like to thank all the staff from Cranfield University, specially Dr. Steve Hobbs and Dr. Peter Roberts, who always found time for advice and guidance in the difficult area of space science. To my colleagues who by intellectual and personal support made this work possible.
I want to mention the help and support from the developers of GRAS and GEANT4 projects, who always answered my questions, specially Giovanni Santin from ESA/ESTEC and John Allison from The University of Manchester
I want to thank my family, my future wife, and my friends who always supported me during this gratifying and challenging time.
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Table of Contents Abstract...... i
Acknowledgements...... ii
Table of Figures...... v
Table of Tables ...... vii
List of Acronyms ...... viii
1. INTRODUCTION ...... 1
1.1. OBJECTIVES ...... 2 1.2. TRAJECTORIES...... 2 1.3. Propulsion…………………………………………………………… .…………………………………………………………. 5 1.4. Constrains…………………………………………………………………………………………………………………..……. 8 1.5. Mars Transit Vehicle ………………………………………………………………………………………………………….9 1.6. Physiological Risks ...... 10 1.6.1. Artificial Gravity ...... 10 1.6.2. Radiation ...... 11
2. SPACE PHYSICS AND RADIATION...... 14
2.1. Space Radiation ...... 14 2.1.1. The Sun and the Solar Wind ...... 14 2.1.2. Galactic Cosmic Rays ...... 17 2.1.3. Solar Proton Events ...... 18 2.2. Radioactivity and Radiation Protection ...... 19 2.2.1. Charged Particles ...... 20 2.2.2. X and γ Radiation ...... 20 2.2.3. Neutron ...... 20 2.3. Energy Absorption ...... 20 2.4. Radiation Effects ...... 22 2.4.1. Deterministic ...... 22 2.4.2. Stochastic Effects ...... 23 3. GEOMETRy AND SOFTWARE DEVELOPMENT ...... 25 3.1. GEANT4 ...... 25 3.2. GRAS ...... 28 3.3. Geometry ...... 30 3.3.1. ISS Model ...... 31 3.3.2. TransHab ...... 32 4. SOFTWARE VALIDATION ...... 37
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4.1. Stopping Power ...... 37 4.2. Total Dose ...... 43 4.3. Hadrons, Electromagnetic and Ion Validation ...... 44 4.4. Astronaut Phantom Test ...... 44 4.5. Validation Results ...... 46
5. SIMULATION OF THE INTERPLANETARY RADIATION ENVIRONMENT ...... 48 5.1. Data Normalization ...... 48 5.2. Computational Parameters ...... 51 5.3. Average Total Dose ...... 52 5.4. GCR Particle Fluence and Energy Spectra ...... 57 5.5. Equivalent Dose ...... 61
6. CONCLUSION ...... 63
7. RECOMMENDATIONS...... 66 7.1. Limitations...... 66 7.2. Further Work ...... 67
8. BIBLIOGRAPHY ...... 70
Appendix A ...... 73 Appendix B...... 74 Appendix C ...... 77 Appendix D ...... 80
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Table of Figure Figure 1: Mars mission trajectories...... 3 Figure 2: Hemisphere and the Interstellar Medium ...... 15
Figure 3: Sun Spot numbers for the latest five cycles...... 16
Figure 4: JPL‐91 Solar Proton Fluence from SPENVIS ...... 17
Figure 5: Solar Min GCR flux outside Earth's Magnetosphere ...... 18
Figure 6: SPE flux worst day case...... 19
Figure 7: Particle flux and energy provided to GRAS to perform particle analysis ...... 29
Figure 8: ISS structure representation with ICRU sphere ...... 32
Figure 9: ISS module structure ...... 33
Figure 10: TransHab layer structure composition...... 35
Figure 11: shows the layer structure of the TransHab...... 36
Figure 12 to Figure 21: Validation Stopping Power ...... 38 to 42
Figure 22: Solar Proton Fluence ...... 43
Figure 23: Dose analysis geometry ...... 44
Figure 24: Spenvis solar proton fluence ...... 45
Figure 25: Alpha particles energy distribution...... 46
Figure 26: Representation of radiation SPEs and Solar Fluence flux simulation...... 49
Figure 27: GCRs Isotropic flux representation ...... 50
Figure 28: GCRs particle flux and interaction with TransHab structure...... 51
Figure 29: Average dose equivalent for ISS and TransHab models...... 55
Figure 30: SPEs effects on ICRU sphere Dose Equivalent values...... 57
Figure 31a: GCRs particle fluence inside ISS model ...... 58
Figure 31b: GCRs particle fluence inside the TransHab model ...... 58
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Figure 32: H energy spectra entering the ISS model during Solar Max...... 59
Figure 33: H energy spectra entering the TransHab model during Solar Max...... 60
Figure 34: GCR and SPE equivalent dose effects in human tissue ...... 61
Figure A1: Solar Max particle flux outside Earth´s magnetosphere from CREME96...... 75
Figure A2: Solar Min particle flux outside Earth´s magnetosphere from CREME96 ...... 75
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Table of Tables
Table 1: Opposition and Conjunction class mission characteristics ...... 3
Table 2: Mars Conjunction Class Mission Opportunities ...... 5
Table 3: Propulsion system comparison ...... 6
Table 4: Mission Constrains ...... 8
Table 5: Mass budget for the transit Mars s/c ...... 9
Table 6: Radiation in the Space Environment...... 12
Table 7: Limits of exposure to sunlight in space ...... 12
Table 8: Organ Dose limits ...... 13
Table 9: Radiation weighting factors ...... 21
Table 10: Ten year human radiation dose limits ...... 24
Table 11: Indicates the different Geant4 physics models ...... 26
Table 12: Physics model used for the simulation ...... 27
Table 13: Indicates the ISS structure composition...... 31
Table 14: Indicates material composition, thickness and quantity for the TransHab structure ...... 34
Table 15: Total dose results from GRAS and SPENVIS ...... 43
Table 16: Validation result from Mulassis and GRAS...... 44
Table 17: Results of astronaut phantom test with proton radiation ...... 45
Table 18: Results of astronaut phantom test with alpha particle interactions...... 46
Table 19 a and b: Total average dose in Solar Max and Solar Min ...... 53
Table 20: Total average dose by Solar Fluence...... 53
Table 21: SPE average total dose effect during worst day case scenario ...... 54
Table 22: ISS Columbus module radiation environment ...... 55
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List of Acronyms
AG: Antigravity
ALARA: As Low As Reasonably Achievable
ASTAR: stopping power and range tables for helium ions
BERT: Bertini Cascade Model
BNTR: Bimodal Nuclear Thermal Rocket
CERN: European Organization for Nuclear Research
CHIPS: Chiral Invariant Phase Space
CPU: Central Processing Unit
CREME96: Cosmic Ray Effects on Micro Electronics
DOSTEL: DOSimetry TELescopes
ESA: European Space Agency
GCR: Galactic Cosmic Rays
GDML: Geometry Description Markup Language
GEANT4: toolkit for the simulation of the passage of particles through matter
GNC: Guidance Navigation and Control
GRAS: Geant4 Radiation Analysis for Space
HEP: High Energy Parameterized
ICRU: International Commission on Radiation Units and Measurements
IMF: Interplanetary Magnetic Field
IMLEO: Initial Mass in Low Earth Orbit
INFI: Instituto Natzionale di Fisica Nuclear)
ISP: Specific Impulse
ISS: International Space Station
JPL: Jet Propulsion Laboratory
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LEO: Low Earth Orbit
LEP: Low Energy Parameterized
MPD: Magneto Plasma Dynamics
MULASSIS: MUlti‐LAyered Shielding Simulation Software
NASA: National Aeronautics and Space Administration
NCRP: National Council of Radiation Protection
NEP: Nuclear Electric Propulsion
NERVA: Nuclear Engine Rocket Vehicle Application
NIST: National Institute of Standards and Technology
NTR: Nuclear Thermal Rocket
PSTAR: stopping power and range tables for protons
QGS: Quark Gluon String
S/C: Spacecraft
SPE: Solar Proton Events
SEP: Solar Electric Propulsion
SPENVIS: Space Environment Information System
TEPC: Tissue Equivalent Proportional Counter
TMI: Trans Mars Insertion
TPS: Thermal Protection System
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Chapter 1
1. Introduction
Throughout the years human space exploration has been an important part of the evolution of society, since the beginning of the space programs man and women have had the opportunity to experience life in space as we move into the 21st century. The future holds new mission and difficult challenges that should be addressed in advance in order to provide continuous support to the evolution of space exploration; missions to the Moon and Mars are becoming a reality pushing today’s technological limits one step forward.
Scientist and engineers should study and understand the dangers of the upcoming journeys and explore to the deepest extent the hazardous situation that will make the voyage a safe one. The space environment pose a threat to astronauts, the Mars transit times for a mission will expose the crew to the longest periods for which humans have ever been in space, pushing their bodies and minds to the limit. Different physical issues should be addressed previously to the beginning of these exiting trips, long exposure to zero gravity, radiation, and psychological issues are the most significant ones.
Current propulsion technology imposes a setback to the progress of deep space exploration; while rocket technology needs improvements new development such as
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nuclear propulsion are still under investigation. The most significant work has been done by NASA in the early years of the human space program in which the problematic of using a different propulsion technology (i.e. nuclear) have been addressed; the selection of the propulsion system has a significant impact on the mission changing transit and surface times in Mars making it a crucial and critical decision to be made in the upcoming years.
The piloted Mars interplanetary transfer orbit (Earth‐Mars) will be a high energy transfer orbit utilizing a fast trajectory to limit the exposure of the crew to radiation and zero gravity effects, where same procedure will be applied to the return transfer orbit (Mars‐Earth). The limiting factors for the transfer time are the entry velocities, in which case a decrease in fly time of one of the legs leads to a higher entry velocity. Therefore, tradeoffs between time and performance will have to be taken into consideration in the design phase.
A common concern throughout the literature indicates the necessity of having the crew the least amount of time in space due to the fast deconditioning that the astronauts will suffer during flight. This is why important attention should be paid to the tradeoff during trajectory analysis.
1.1. Objectives
The work done in this thesis tries to addresses the current issues in deep space exploration, as mentioned in the previous section there are many areas, which need review and a deeper understanding before taking this endeavor. Currently, there are many efforts among the scientific community to tackle some of these problems; this is why the work done in this paper intends to explore the effects of radiation exposure to humans inside different structures in deep space conditions.
1.2. Trajectories
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There are two main trajectory types for human mission to Mars, conjunction class and opposition class (figure 1). Both of this trajectories have direct implication on the time
zspent in the red planet upon arrival, the first one, conjunction class, proposes a longer stay in the planet while opposition a longer flight time.
Figure 1: Mars mission trajectories. (Drake, 2007)
According to the NASA Exploration Blue Print Data Book (Drake, 2007), mission times ranges between 365‐661 days for opposition class and 892‐945 days for conjunction class with surface stay times ranging from 30 to 596 days accordingly. Each of these mission architectures has advantages and disadvantages that will be fundamental to the decision of the final mission planning. Some of these differences are expressed in table 1.
Table 1: Opposition and Conjunction class mission characteristics
Opposition (Short Stay) Conjunction (Long Stay)
Transit Vehicle Larger vehicle to The use of advance propulsion accommodate the crew enhances the mission implies larger mass architecture decreasing the requiring advance mass and shortening the trans propulsion for a reasonable mars injection (TMI) mass at Earth orbit
Trajectory Venus Swing‐by, distance Direct from Sun ~0.7 AU
Departure and Larger delta V and large Shorter delta V throughout the Arrival Velocities propulsive requirements mission
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Human Health Long zero‐g space mission Zero‐g gravity exposure similar and longer radiation to ISS experience and less time exposure to the crew. exposure to the radiation Venus Swing‐by increases environment the exposure to radiation
Abort Mission Propulsive abort capabilities Free return abort capability, Trajectory two‐year free return, and three‐year free return.
In addition to the normal consideration for such a trip, abort mission considerations shall be made. There are many options ranging from free return mission to propulsive transfer times that will affect the flight trajectory around Mars. The selection of the abort mission characteristic will depend on the orbit energy and delta v capabilities at Mars and the Earth. Of course the selection of the abort mission will commit the crew to that specific orbit, therefore in‐depth trade studies should be made to provide the best option for the crew in case a problem arises during the Trans Mars Insertion (TMI).
Due to the implication of a human mission to Mars and to the fact that the space transit times between transfer orbits are not very different, a longer surface stay should be intended for the first mission, decreasing the human exposure to the space environment until a full understanding of the implications of long exposure to zero‐g and radiation.
According to ESA and NASA exploration programs 2030 is the decade in which a human mission to Mars will take place. Therefore, according to the conjunction class transfer orbits specific dates will be available for departure during that time. Table 2 shows different opportunities obtained from a NASA Technical Memorandum (Young, 1984). The dates are in Julian date format.
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Table 2: Mars Conjunction Class Mission Opportunities (NASA, 1984)
Mission Earth‐ Leave Arrive Leave Arrive Outbound Mars Inbound Total Mars Earth Mars Mars Earth Trip Time Stopover Trip Mission Year Opposition J.D. (Days) Time Time Time J.D. 2460000 (Days) (Days) (Days) 2460000 (EMOS)
2031 2989 2860 3142 3642 3858 282 500 216 998 (.1107) (.1165) (.0826) (.1202)
2033 3775 3706 3906 4456 4656 200 550 200 950 (.1016) (.1111) (.0996) (.1013)
2035 4584 4508 4712 5242 5512 204 530 270 1004 (.1091) (.0876) (.1202) (.1129)
2037 5382 5290 5646 5986 6276 356 340 290 986 (.1363) (.0947) (.1046) (.0957)
2040 6157 6052 6392 6732 7036 340 340 304 964 (.1193) (.0833) (.0922) (.0968)
2042 6920 6812 7130 7470 7802 318 340 332 990 (.1060) (.0836) (.0870) (.1127)
2044 7688 7568 7874 8214 8564 306 340 360 996 (.1003) (.0940) (.0866) (.1318)
1.3. Propulsion
There are many different propulsion systems that could be potentially developed to perform the transfer trajectory to Mars. The development of new technologies is a requirement for the completion of this mission and it will be an advantage for the crew due to the shorter transit times given by new technologies. Griffin et al (B. Griffin, 2004) provides a quick comparison table (table 3) that gives a rough idea of the advantages and disadvantages of each propulsion system.
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Table 3: Propulsion system comparison (B. Griffin, 2004)
Propulsion Description Advantages Disadvantages Options
Chemical Conventional cryogenic rocket ‐Mature technology ‐Low performance engines. ‐High thrust, short burn times leads to high IMLEO Insulated tanks with vapour‐ ‐Ballistic interplanetary except for cooled shields to reduce boil transfers facilitate conjunction profile off. Start T/W 0.1 to 0.25 implementing artificial gravity with long transfer Isp ~ 460s times ‐Cryogenic with hydrogen, low density, needs heat leak control ‐Expendable system
Chemical/Aeroc Same as chemical except ‐Reduces IMLEO by replacing ‐Performance still apture aerocapture used for MOI. one major manoeuvre with marginal for "hard Large aeroshell needed aerocapture year" opportunities requiring either intact launch ‐Aerocapture risk: or in‐space assembly. Lander TPS/thermal, GN&C may capture separately to ‐Mars Vhp limited to simplify configuration. ~ 6 Km/s for safe aerocapture ‐Expendable system
NTR Nuclear thermal rocket ‐Known technology ‐Nuclear costs and engine, hydrogen propellant, ‐Twice the Isp of chemical risks Isp ~ 900s. propulsion reduces IMLEO ‐Engine test protocols Usually drop tanks utilized for and sensitivity to not resolved (how to each major manoeuvre. opportunity contain radioactive Insulated tanks as ‐High thrust, short burn times products) above; start T/W <= 0.1 to ‐Ballistic interplanetary ‐Cryogenic with transfers facilitate hydrogen, low
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reduce nuclear engine size. implementing artificial gravity density, needs heat leak control ‐Expendable system
SEP Large (multi‐megawatt) solar ‐Known technology with flight ‐Large size may electric propulsion system, experience in small size require more space performs all major ‐High Isp reduces IMLEO and assembly than other manoeuvres. sensitivity options Isp typically 3000s; ‐No hydrogen propellant ‐High‐power electric MPD or comparable thrusters. ‐Reusable system thrusters not mature Achievable power‐to‐ mass ratio may not permit opposition‐ class profiles
NEP Large (multi‐megawatt) ‐Known technology (no space ‐Nuclear costs and nuclear electric propulsion experience or experimental risks system, probably Brayton or prototypes except ‐Large size may liquid metal Rankine power thermoelectric and require more space generation performs all major thermionic conversion) assembly than other manoeuvres. ‐High Isp reduces IMLEO options Isp typically 3000s; ‐No hydrogen propellant ‐High‐power electric MPD or comparable thrusters. ‐Potentially reusable system. thrusters and space configuration power conversion not mature ‐Achievable power‐ to‐mass ratio may not permit opposition‐ class profiles
During the past decade many different propulsion systems have been considered for this mission; each of whom has pros and cons as mentioned in table 3, therefore in order to fulfill all the mission requirements extensive research is needed in this area.
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Nuclear thermal propulsion is a possible solution as pointed out by Borowski in his paper (Borowski, et al., 2001), emphasizes the extensive research which was conducted during the 60’s and 70’s in the NERVA program (Nuclear Engine for Rocket Vehicle Application) leading to a few certain answers. On the other hand, nuclear electric propulsion (NEP) has a large advantage over the NTP since it does not use LH2 propellant, simplifying the issues of storing cryogenic liquids. According to research performed in NTP and in BNTR the amount of propellant (LH2) needed for the entire mission is 50 t (tons), using the data obtained from the NASA Exploration Blueprint Data Book (Drake, 2007) the expected total mission mass for a long stay (conjunction class transfer) in Mars ranges between 400 and 700 t.
Nuclear propulsion shielding provided by (Borowski, et al., 2001) mentions a minimum required 2.84 kg/MWt of reactor power. Therefore for a 335 MW reactor proposed by (Gandini, 2003), the total nuclear reactor shielding is
1.4. Constrains
Table 4 provides final constrains for a typical mission.
Table 4: Mission Constrains
Parameters Constrain Crew size 6 Orbit Transfer Conjunction (Long Stay) Total Mission to Mars (days) 892‐1004 Abort Mission Scenario 2 years free return trajectory Total Mission Mass (mt) 400‐700 % Vehicle Mass 31% Radiation GCR and SPE studies Zero‐Gravity Partial AG with exercise Earth Departure June 29th 2035 Mars Arrival January 19th 2036 Mars Departure July 12th 2037 Earth Arrival March 29th 2038
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1.5. Mars Transit Vehicle
According to the NASA Exploration Blueprint (S. Hoffman, 1997), the transit spacecraft should provide supporting capabilities for a crew of six for up to 200‐day transitions to and from Mars, provide zero‐g countermeasure and deep space radiation protection, return propulsion stage integrated with transit system, and provide return to Earth abort capability for up to 30 hours post Trans Mars Injection (TMI) (Drake, 2007). Table 5 provides a table indicating an ideal mass budget for the transit Martian spacecraft.
Table 5: Mass budget for the transit Mars s/c (Drake, 2007)
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1.6. Physiological Risks
During the mission to Mars astronauts will be exposed to a harsh environment that will affect different areas of their bodies. Long exposure to microgravity will generate bone loss, cardiovascular alterations, muscle atrophy and neurovestibular disturbances (Drake, 2007); while exposure to radiation will generate short term issues (physiological effects) such as cataracts, acute radiation sickness and damages to central nervous system and long term issues such as cancer, hereditary effects and neurological disorders.
In order to get the most out of the crew and to accomplish the mission objectives, the astronauts should be protected against the physiological effects imposed by space exploration. In many of the proposed architectures for the human mission to Mars the leading method for solving the zero‐g issue is the artificial gravity generation; where for radiation mitigation the proposed technique is to use shielding in the s/c structure; the main difficulty with this is to find materials that will provide crew protection for the broad range of environmental effects caused by space radiation.
1.6.1. Artificial Gravity
W. Von Braun and C. Clarke introduced the idea of artificial gravity many years ago; it provides a countermeasure and acts as a mitigation technique addressing a series of physiological risks during human space flight. This device should maintain the crew’s level of physical performance without affecting psychological aspects created by the constant transition between gravity scenarios.
Artificial gravity should provide benefits and eliminate the effects associated with the zero‐g environment using s/c spinning (entirely or partially). Many different investigators indicate that a combination between AG, exercise, and a strict diet is likely to be the optimal mitigation technique for certain health risks faced by the crew (C. Allen, 2003).
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In the NASA Guidelines and Capabilities for Designing Human Missions (NASA, 2003) report some AG design considerations are addressed.
Intermittent AG Continuous AG
Short arm centrifuge: During S/C rotation or crew compartment
this approach crews are exposed rotation: This may affect the to AG during their sleep time or nervous system with the constant
during power exercise times changes of sensor stimulation introducing a number of g induced by the rotation. Coriolis
transitions between the desired forces created by this rotation will level of g and the zero‐g give the feeling of motion; this environment. may cause nausea and vomiting during the entire mission. Also long duration AG might affect how subjects readapt to the Mars‐ Earth environment.
1.6.2. Radiation
Radiation is the greatest risk for the astronauts’ health; during spaceflight the crew will be exposed to two different types of radiation, ionizing and none ionizing. The radiation environment during a mission to the red planet will fluctuate considerably over location and time; this means that a crew will require different protection techniques over the course of their flight. Table 6 shows a table from a NASA case study (C. Allen, 2003) that indicates the different types of radiation, their frequencies and their source; this table helps understand the main issues and provides guidelines for designing countermeasure techniques.
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Table 6: Radiation in the Space Environment. (C. Allen, 2003)
Ionizing radiation is a particular concern among flight physicians since it damages the genome as it travels through the cell’s nucleus. “It also affects the genome by producing free radicals and by transducing signals between adjacent cells within the body.” (C. Allen, 2003)
Currently there are constrains imposed by the National Council on Radiation Protection and Measurements (NCRP) on crew radiation exposure, this are based on past studies of similar earth jobs and LEO astronauts exposure; this should only be use as a guideline for the time being but future consideration and new constrains should be provided.
Non Ionizing radiation
Table 7: Limits of exposure to sunlight in space (C. Allen, 2003)
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Ionizing Radiation
Table 8: Organ Dose limits (C. Allen, 2003)
One of the mitigation techniques proposed by the NASA Radiation Shielding Materials Workshop (NASA , 2000) are hydrogen‐based liquids such as liquid hydrogen and recycled water, these could be implemented in the spacecraft structure to protect the crew against space radiation.
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Chapter 2
2. Space Physics and Radiation 2.1. Space Radiation
The space radiation environment outside Earth’s Magnetosphere consist of Galactic Cosmic Rays (GCRs), Solar Proton Events (SPE) and Solar Fluence. The variation of these phenomenon are directly linked to the solar activity, the Sun’s changes and its cycles pose a direct effect on the flux and time variation of the space radiation environment. Whilst at solar maximum a higher number of sun spots are present and a higher probability of the occurrence of an SPEs the GCRs’ flux decreases accordingly and inversely to the solar wind strength, on the other hand, when the sun is at solar minimum the opposite happens, the decrease of the intensity in the solar wind allows the increase in the GCR’s flux.
2.1.1. The Sun and the Solar Wind
The Sun is a class 2 star at the center of the Solar System which accounts for 99.8 % of its mass; the distance to the Earth is about ~150 million km or 1AU, and like most stars the sun is made of 74 % Hydrogen which accounts for 92% of its volume, 24 % Helium as well as other quantities such as oxygen, carbon, and neon to name a few. The
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solar wind is a continuous flow of ionized plasma from the Sun’s corona, it consist mainly of H with a small percentage of helium. The solar wind is a supersonic flow with speeds around 400 km/s; therefore the plasma travels from the Sun to the Earth in approximately four days, this plasma extend to the Heliosphere decreasing intensity as it gets farther and away from the Sun (ERSMARK, 2006).
Figure 2: Heliospehre and the Interstellar Medium (Eastman, 1990)
Due to the rotation of the Sun the fields form a spiral effect and the plasma flows rapidly away carrying the frozen‐in magnetic field.
Many aspects of the Sun behavior are periodical and follow an 11 year cycle which affects the interplanetary medium, during times of high activity there may be release of energy and matter in the form of solar flares and coronal mass ejections which impose a threat to spacecrafts. A fair indication of sun’s activity is the number of sunspots, which can be observed from earth and are an indication of the phases of the cycle; during solar minimum the Sun’s surface is almost spotless, the number o spots increases reaching the maximum value during solar max. Figure 3 indicates the number of sunspots per cycle.
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Figure 3: Sun Spot numbers for the latest five cycles (2009)
The variation of the solar cycles affects the solar wind and the Interplanetary Magnetic Field (IMF) exerting a large influence to the dynamics of the Solar System making a fundamental effect in any model or simulation.
The data used for the simulations was obtained using the JPL model from SPENVIS (Space Environment Effects and Information System); SPENVIS is a web interactive tool that provides scientists and engineers the required information on space environment, it has the capabilities of generating orbits and allowing the study of radiation effects on the spacecraft (Belgian Institute for Space Aeronomy, European Space Agency, 2009). The data used by the JPL‐91 model consists of continuous records of daily average fluxes, and the model assumes that there were no significant proton fluence events during these periods. The JPL model uses a 7 year period of the active sun in order to consider for the statistical events occurring during that period. Figure 4 shows a plot of the results obtained from SPENVIS
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Figure 4: JPL‐91 Solar Proton Fluence from SPENVIS at near Earth Interplanetary conditions
2.1.2. Galactic Cosmic Rays
The dominant flux of GCRs is believe to come from supernova events, its thought that Fermi acceleration is the cause of particles accelerating near the speed of light. 90% protons, 9% alpha particles, and the remaining 1% of various ions compose the GCR flux. These particles are trapped inside the galactic magnetic field and travel around the galaxy various times before reaching the Solar System. (ERSMARK, 2006)
The GCRs’ energy varies with quantity, more particles are in the lower energy of the spectrum whether high energetic particles have very small flux, and therefore the analysis performed in this investigation was truncated to 102 GeV energies due to the low flux of the remaining particles. Figure 5 shows the data obtained from the CREME96 (Cosmic Ray Effects on Micro‐Electronics) Model for GCRs flux outside Earth's magnetosphere. The CREME96 model allows the creation of numerical representation of
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the radiation environment in near‐Earth orbits making it an important tool for this study (NRL, 2007).
Figure 5: Solar Min GCR flux outside Earth's Magnetosphere
The charge particles within the GCRs are affected by the electromagnetic environment present in the Solar System (IMF); the IMF affects the flux of particles depending of its intensity, for instance when the Sun is in the solar maximum period it produces a much stronger IMF decreasing the GCR particle flux. The opposite happens when the sun is at solar minimum, this modulation is very important for human mission design since affects directly the s/c environment.
2.1.3. Solar Proton Events
Solar Proton Events are related to the solar maximum period and closely associated to the solar‐flare events; protons and ions, which interact with the
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interplanetary medium to travel through the Solar System, compose the particle flux. It is thought that the acceleration of the particles is done by the recombination of the Sun's magnetic field with its surface producing the flux indicated in figure 6.
Figure 6: SPE flux worst day case
The analysis performed in this thesis was completed using the particle flux representing the worst day scenario using CREME96; figure 10 shows the energy spectrum of this event. The main issue when studying the SPEs effects on humans in interplanetary space is that the lack of protection from the earth's magnetosphere, this is an unavoidable problem which has harmful consequences during the entire mission.
2.2. Radioactivity and Radiation Protection
In 1896 Becquerel discovered that some naturally occurring elements were radioactive by observing the blackening of photographic films in the vicinity of Uranium (Martin A., 1996). This radiation process and transformation is called radioactive decay
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that usually is manifested in the form of charged particles and gamma rays. These types of radiation interact with matter producing different effects.
2.2.1. Charged Particles
Alpha and beta particles loose energy by interacting with the electrons in the medium causing them to raise their energy levels (i.e. excitation) or by separate them from the atom (i.e. ionization). A very important effect when the charged particles are slowed rapidly is the emission of X‐rays. (Martin A., 1996)
2.2.2. X and γ Radiation
There are three main processes for which this radiation interacts with matter: photoelectric effect, Compton scattering, and pair‐production (Martin A., 1996). In the first one the total energy is transferred to the electron that is kicked out of the atom. The second one, only a partial amount of the incidence energy is lost during the particle interaction with the electron, then the incidence particles continues with lower energy. The pair production is the resulting case in which gamma rays may interact with an intense electric field producing an electron‐positron interaction (Martin A., 1996).
2.2.3. Neutron
Since the neutron is an uncharged particle it cannot produce ionization, ultimately neutrons transfer their energy to charged particles that will produce ionization.
2.3. Energy Absorption
Absorbed Dose: “is a measure of energy deposition in any medium by any type of ionizing radiation.” (Martin A., 1996) The unit is the gray [Gy] that represents the energy deposited in a volume by the mass of that volume [J/kg]
Dose Equivalent: In order to evaluate the risk of the absorbed dose many characteristics should be taken into consideration such as particle types, species, and
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MASTER’S THESIS
energy. The dose equivalent (ICRP 92, 2003) can be calculated as the sum of the energy deposition in combination with the radiation quality factor Q.