Inspiration Mars: Team JASPer

Alex Nikle [email protected]

Patrick Boyce [email protected]

Joyce K. Greene [email protected]

Sarah Pyle [email protected] TABLE OF CONTENTS Introduction ...... 3 Trajectory ...... 5 Reentry ...... 9 Launch and Assembly ...... 13 Chariot Assembly...... 15 Phase Zero ...... 16 Phase One...... 17 Phase Two ...... 17 Phase Three ...... 18 Phase Four ...... 18 Chariot Habitat Design ...... 19 Interior...... 19 Radiation Shielding ...... 21 Physiological Constrains ...... 26 Psychological Considerations ...... 27 Chariot Environmental and Life Support Systems (ECLSS) ...... 28 Engineering Busses ...... 34 Power ...... 35 Navigation ...... 36 Communication ...... 37 Selection ...... 40 Chariot Economics ...... 42 Conclusion ...... 45 References ...... 46

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INTRODUCTION Human and have come a long way in the last 60 years. The dream of man voyaging into outer space was merely science fiction until April 1961, when Cosmonaut Yuri Gagarin made the historic journey into space. Shortly thereafter, the race to reach the Moon began between the U.S. and U.S.S.R. In less than a decade, the question of which superpower would reach the Moon first was settled when the first footprints were forever imprinted on the lunar surface on July 1969, by two American . Two decades later, the fleet launched into low Earth orbit (LEO) and ferried equipment to help build the International Space Station (ISS). When the ended in 2011, private industries with the support from the National Aeronautics and Space Administration (NASA) emerged. This filled a gap in U.S. launch capability where the Space Shuttle left off. The question now is what’s next? Now that NASA has contracted International Space Station resupply missions to private industries, they have the capability to steer attention toward exploration beyond what they are already capable of. Although Mars is the next dream destination for manned exploration, the journey and survivability to get to the next celestial body and back will be the hardest feat to overcome in space efforts to date. A manned mission to Mars requires exposure to the elements of deep space for durations experienced by no other in human history. This mission would certainly test the capabilities and limits of modern space engineering to ensure the survival and safe return of the crewmembers aboard. The Mars Society continues to support this major step in manned space exploration through the creation of the Inspiration Mars Competition. This contest requires a proposal of a fly-by mission design to the red planet. The competition is open to all universities throughout the world in effort to inspire students to participate in space exploration and contribute ideas into future space engineering designs. We, the members of team JASPer, have elected to participate in this competition by submitting the following mission design. This proposal includes history, policy, economics, and of course engineering and science requirements to accomplish such a daunting challenge. It is our belief that this mission design proposal uniquely satisfies these requirements that will support the first real-life human mission to Mars whether that is in 2018 or anytime beyond. The members of Team JASPer are proud to be a part of this competition and hope that our design contributes to the future of human space exploration. Our is named “Chariot.” Chariots, used as ancient “battle taxies” represented mobility and strength. Centuries after they ceased being used in battle, they continued being used in competition. Often manned by two personnel, they vied for speed. It seems appropriate that we should compete for a manned mission around Mars in a modern day “Chariot.” Are we competing against anyone? Russia? China? Europe? Is there in fact a going on? Maybe or maybe not depending on which article one reads. The United States won the race to the Moon in 1969. China might meet that objective in the next decade or two. Arguably, this is a race with ourselves. It is a race for a personal best, for national pride, for the revitalization of our space industry, and for the inspiration of future generations.

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Our Chariot’s three core elements are a Bigelow Aerospace 330 (BA330) habitat, a Dragon capsule for reentry, and a J-2X booster engine that is being designed for NASA’s System. We will assemble our 55,000 kg spacecraft in low Earth orbit with the support of the International Space Station. SpaceX will provide the Falcon 9 and Falcon Heavy launch vehicles for this mission. NASA will support with J-2X booster technology, ISS activity, , access to its near Earth and deep space communications systems, and sharing in its public outreach program support. Our fuel tanks usefulness will not end after the trans- Mars injection boost. They will be built around our habitat to provide a redundant radiation shield for the BA330, the heart of our mission. It is designed predominantly as self-contained space station. It contains its own power systems, communications, thermal regulation, navigation and life support systems that our team will both use and augment to ensure success during the 501 day mission. It is built to withstand micrometeoroid and space debris strikes as well as a moderate amount of solar and galactic radiation. Its 330 m3 internal volume dwarfs the internal volume of an Orion capsule and will thus allow more supplies, more equipment, and more redundancy to make this mission a success. Furthermore its size will allow astronauts more room to spread out for more optimal psychological readiness. If all required programs maintain their current development progressions then our Chariot may be ready for the 2018 launch. This will require a proactive management team with a deep wallet. There should be no question that even if all systems are TRL 7-9 that this should still be considered a very high risk mission. The rewards are high though. The Apollo landing on the Moon in 1969 represented an achievement not just for the United States but for all mankind. As hard as it was, it was local. The next step will be to land on another planet. Once that happens, the others, such as landing on Jovian moons, will be milestone events but that “first” will not be lost. Those milestone events will owe their legacy to those that landed on the Moon and Mars. The Inspiration Mars mission is not about landing on Mars, this will not likely come until at least the 2030s if not decades later. The Apollo 8 lunar fly-by paved the way for the Apollo 11 landing just as the Inspiration Mars fly-by will pave the way for subsequent landing missions. Once proof of concept is established, the readiness and volume of future exploration may proceed. This Inspiration Mars mission is a test flight for our “Chariot”, the enabling technologies, and its effects on our astronauts’ ability to survive a long duration space mission beyond the sanctuary of Earth’s magnetosphere. Inspiration Mars is indeed an appropriate name but we do not believe it is about Mars at all. It is bigger than that. It is about inspiring a future of interplanetary travel and manned space exploration. It is about those generations that watched Apollo unfold on television sharing that same that same excitement and wonder. Team JASPer thanks the Greatest Generation and the Baby Boomers for the engineering miracles they performed. Though it would be an honor for our mission design to actually be flown, we are modest enough to know that many of the smartest people in the world have shared similar dreams. The synthesizing of our team has been rewarding and is a victory in itself. This comes despite the thousands of miles between us and is testament to the efforts that would be required for a mission such as this. We hypothesize that top universities in the world have had close-knit teams that have met regularly to go over their designs. Our team members live in North Dakota, Florida, South Carolina, and South Korea. We believe that this makes our team’s

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proposal unique. We have only had phone calls, emails, Skype, and chat rooms to coordinate our efforts. Our backgrounds are as varied as where we live. This lends itself to diverse life experiences that we have used in formulating our design. We are all part-time graduate students in University of North Dakota’s Space Studies Department. In our current job titles, one is a meteorologist, one is a commercial pilot, one is a flight instructor, and one is a prison warden. Three of the four of us have military experience. We are split in sexes: two females and two males. We have never met in person, yet we share a vision for interplanetary travel. We believe that our life experience combines with our academic experience and adds to the uniqueness of our team. We are not a group of engineers coming together to develop an engineering project. We are a team of widely different backgrounds that have come together from all over the world for this Inspiration Mars competition and our overcoming our diverse backgrounds is another victory we are savoring. Whether our design is chosen or not, Team JASPer has learned a tremendous amount about both about the details of designing a mission to Mars and about ourselves and our ability to complete this submission despite our distance and diversity. Trajectory

One might assume that the design of a habitat or spacecraft is the first requirement of a mission design. Team JASPer has taken a different approach. The trajectory we take is the primary factor driving our requirements. The length of time for our mission is important, but so is the actual timing of our departure. As we will address in the life support section, we will want to minimize the threat of solar flares by flying during solar minima. We will also want to launch at a time when the trajectories of the planets are optimal for both minimizing the time to travel and fuel required. Mars eccentricity is five times that of Earth’s. In fact, the perihelion of Mars is 43 million kilometers closer to the Sun than at aphelion. That is about 30 percent the distance between the Earth and the Sun! Therefore, the orbit of Mars is a much more significant driving factor in timing that the Earth’s orbit. There are two primary trajectory designs we considered. The first is a straight orbit-to- orbit “Hohmann transfer” maximum efficiency design and the second, our preferred trajectory, is a faster shortcut trajectory to Mars that flies interior to the Earth’s orbit for the initial outbound and final inbound portions of its journey. The final solution derived is a patched-iconic solution where our spacecraft’s velocity is reduced during its fly-by of Mars. This lost velocity is critical for its free-return arrival at Earth. Venus rendezvous trajectories have been eliminated from our Inspiration Mars mission design because of likely additional costs of mission operations due to precision tracking and trajectory requirements; and greater velocity change requirements due to those orbital corrections. The orbit-to-orbit design implies that the spacecraft is never closer to the Sun than Earth’s orbit. Miele and Wang (1999) made several calculations to show when Earth and Mars would be in position for these optimal trajectories. This designation as “optimal” implies that it is performed with a minimum change in velocity required (ΔV). They calculated this would occur when a spacecraft departs Earth with an accelerating impulse of ΔVLEO = 3.552 km/s and a phase angle of -61.85 degrees relative to Earth’s trajectory. This study calculates an entry into low Mars orbit so there is a corresponding braking velocity; however, it takes 257.88 days to arrive at

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low Mars orbit. Mars should be ahead of the Earth, in its phase angle compared to the Sun, by 43.86 degrees at the spacecraft’s departure from Earth. It should then be behind the Earth by 75.13 degrees at arrival at Mars. This calculation also assumed that the departure from Earth occurs during Earths perihelion and arrival occurs during the Mars aphelion. This optimal trajectory is rare so various suboptimal trajectories were also reviewed. Miele and Wang (1999) also made calculations in what they referred to as “suboptimal” trajectories. These suboptimal trajectories calculate orbital transfers but assume different Earth and Mars phase angles. Some suboptimal trajectories have reduced time of flight advantages. These reduced time of flight trajectories fly interior to the Earth’s orbit going to and from Mars. There are times of flight advantages but weaknesses as well. A change of +15 degrees in phase angle would come from departing 32 days after the optimal date. This orbital transfer calculation caused an eight percent needed increase in velocity and that extra required fuel led to a 13 percent decrease in payload mass. The advantage of this calculation was that it decreased flight time to Mars by 13 days. This data correlates with other research for similar types of trajectories. In 1998, Patel, Longuski, and Sims calculated Mars free return trajectories. Free-return trajectories with 1.4 year times occur every 15 years. This research also showed requirements for increased ΔV requirements. This orbital data continued to fit well with later calculations made by Miele and Wang (1999) for suboptimal trajectories. The next time Earth and Mars are in such a position will be in December 2015 and then December 2017. Additionally, 2018 coincides with a solar minimum increasing the benefits to launching during this window. The next opportunity for this 1.4 year trajectory will be in December 2030 and 2032 when dangers from solar radiation will be much more severe. These combined free return and solar minimum windows have led to the inspiration Mars competition.

Figure 1 Suboptimal Mars free return trajectory (Patel et al., 1998) An optimal trajectory is not possible during the Inspiration Mars launch window. For this suboptimal trajectory, V∞ Earth has been derived as 6.232 km/s through “Mission Analysis Environment” and “Astrogator” modeling software (Tito et al., 2013). Team JASPer’s Chariot will be assembled alongside the International Space Station. Using a nominal altitude of 400 km, a calculation is made to derive its orbital velocity at that parking orbit.

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Calculating Vpark at Earth: Given: 400 km parking orbit altitude (for International Space Station supported assembly) 5 3 2 µEarth = 3.986 x 10 km /s Rpark = radius of the parking orbit = 6778 km

Attached to the International Space Station for assembly, a plane change is needed prior to our hyperbolic departure to Mars. The ISS’s inclination is approximately 50 degrees. The ecliptic is 23.5 degrees. The plane change is thus approximately 50 degrees – 23.5 degrees = 26.5 degrees. For simplicity, and for the highest initial planning calculations, our preliminary calculations have been made by performing a simple plane change at 400 km in low Earth orbit and then begin our ΔVboost. This plane change velocity estimate is the most demanding of plane changes compared to other possibilities including combined plane changes in low Earth orbit or near Mars. Additional modeling is desired to find more optimal solutions to significantly reduce the total change in velocity requirements, and subsequently less fuel mass required to perform those changes.

After assuming position within the desired plane (to Mars) we calculate our ΔVboost. To calculate the ΔVboost, the velocity change required going from a parking orbit around Earth to its hyperbolic-departure trajectory, we subtract the velocity at the parking orbit from the velocity of the hyperbolic departure to Mars. Spacecraft energy on a hyperbolic escape trajectory is used to calculate the hyperbolic departure velocity. V∞ Earth is given, through previous modeling as 6.232 km/s (Tito et al., 2013).

= = = 19.42 km2/s2

This energy is then used to calculate Vhyperbolic at Earth:

From this now, ΔVboost can be calculated:

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ΔVboost = = 12.51 km/s – 7.669 km/s = 4.841 km/s

This, 4.841 km/s ΔVboost is the change in velocity needed for Chariot to go from its low Earth parking orbit to Mars and back via free-return trajectory. The amount of fuel that must be used for this trans-Mars injection (TMI) boost can be calculated by using the equation. The fuel used by the J-2X rocket has a specific impulse (Isp) of 448 seconds. We also calculated the initial Chariot dry mass to include: BA330 habitat, Dragon capsule, J-2X engine, fuel tanks, and supplies as 55,000 kg. The mass of the fuel tanks supporting this fuel load is estimated at 6.4 percent of the total fuel mass (Wertz & Larson, 1999).

The mass of the fuel required for this hyperbolic departure:

Massinitial – Massfinal = 110,478 kg Now that we know the total mass of our Chariot rocket on its hyperbolic trajectory, we know the mass of what we need to conduct our plane change with. The amount of fuel needed to perform the simple plane change can again be calculated by using the rocket equation:

The mass of the fuel required for this plane change = Massinitial – Massfinal = 202,726 kg

The total mass of fuel = Fuel MassTMI + Fuel MassPlaneChange = 313,204 kg Chariot final mass following ISS separation = 368,204 kg

Chariot Spacecraft Mass Component Mass (kg) Bigelow BA 330 20,000 SpaceX DragonRider 6,000 Habitat ECLSS/Supplies 5,000 J-2X Engine 2,430 Supporting Infrastructure 1,570 Fuel Tanks 20,000 Chariot Dry Mass 55,000 Fuel 313,204 Final Mass 368,204 Table 1: Chariot Spacecraft Mass After conducting the orbital change and trans-Mars Injection burns, the Chariot crew will begin the 227 day journey to Mars (Tito et al., 2013). Upon arrival, the red planet will approach

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from the rear as it overtakes the orbit of the Chariot spacecraft. Its gravitation force will then cause a reduction in our spacecraft’s velocity. Chariot will fly around Mars from the direction of Mars path, fly through the night side of Mars, and depart from the trailing side of Mars back to Earth. One point of Inspiration Mars advertises a fly-by at approximately 160 km. The flyby through Mars atmosphere at 160 km will be a factor to contend with. With limited fuel on board, the effects of the Martian atmosphere necessitates refinement in mission altitude. The Martian atmosphere at 160 km has a density between 10-9 and 10-11 kg/m2. Earth has an atmospheric density of 1.6x10-9 and 2x10-9 kg/m2, at those altitudes. The Martian atmosphere is denser at warmer temperatures and during dust storms (Fujita et al., 2014). Inspiration Mars should plan to account for those higher densities including those of dust storms that emerge during the mission immediately preceding our spacecraft’s arrival. In comparisons with known orbital decay around low Earth orbit, it is strongly recommended that an atmospheric density of no less than 10-11 kg/m2 be encountered. This correlates with an approximate Mars altitude of 250 km. Higher is preferred to ensure increased probability of mission success. This is in line with Mars orbits such as Mars Reconnaissance Orbiters 255 – 330 km altitude.

Reentry After the challenges of assembling Chariot in low Earth orbit, dodging micrometeoroids, and surviving 501 days in microgravity and solar radiation; one of the most challenging parts of the Inspiration Mars mission awaits - reentry. Humans have only returned to Earth via direct entry, such as in Apollo capsules, and gliding, such as in the Space Shuttles. The Inspiration Mars astronauts and capsule will have pushed the limits of space endurance so the severity of re- entry must be mitigated. The Bigelow 330 capsule will not survive reentry so our Chariot will jettison everything except for the Dragon capsule immediately prior to recovery. This will significantly reduce mass and leave us with a well modeled “Apollo like” blunted cone . The reentry mission design technologies are on track for success; however, additional testing on heat shielding is required.

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Figure 2: Re-entry The atmospheric reentry Velocity of our spacecraft back at Earth is calculated to be 14.2 km/s (Tito et al., 2013). This is a much faster velocity than manned recoverable spacecraft have ever attained. Slowing down from this velocity with available technology will entail both tremendous heating and deceleration “Gs.” Solving for structural strength for maximum deceleration will be critical for our mission engineers to plan for.

: Vehicle’s maximum deceleration (m/s)

: Vehicle’s re-entry velocity (m/s) : Atmospheric scale height = .000139 m-1 for Earth : Vehicle’s flight-path angle (deg) e: Base of natural logarithm = 2.7182…

G = amount of gravitational force the astronauts and capsule will endure. This calculation has been performed via spreadsheet and the results are presented in the figure below.

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A Comparison of Re-Entry Angle and Deleleration at Varying Re-Entry Velocities 600

520 528 500 496

457

400 404 14.23 km/s 369 375 352 339 325 12 km/s 300 287 264 10 km/s 245 256 260 241 225 200 8 km/s Deceleration(Gs) 187 199 180 167 164 166 144 156 6 km/s 128 130 127 100 107 91 89 83 81 88 92 93 4 km/s 65 57 60 71 45 46 39 41 41 28 32 20 26 31 36 0 716 14 1 2 3 4 5 6 7 8 9 Re-Entry Angle (x 10 degrees)

Figure 3: Aerocapture Re-entry Deceleration Over 90 Gs are experienced during a direct reentry from 14.2 km/s, even at a very shallow ten-degree reentry angle. This is well beyond safety margins for any crew and especially a crew that has been in microgravity for over 501 days. A reduction in deceleration can be made by using a combination aerobraking and reverse thrusting. Re-entry will be comprised of two phases. The first phase is an aerobraking descent that will “skip” off the atmosphere. The second phase will be the actual reentry and descent during our spacecraft’s next approach to Earth. The first phase will be a descent to a 56-60 km perigee of Earth. This aerobraking will slow our final re-entry velocity from 14.2 to 9.5 km/s and lead to a very shallow phase 2 recovery. This 56 km perigee comes with the cost of a stressful 9 Gs. This is within the bounds of Apollo mission’s 6-11 Gs during reentry. The Dragon capsule seats astronauts very similar to how the Apollo astronauts were seated. This seating arrangement puts most Gs along the astronaut’s x-axis, chest to back, where the body can tolerate Gs best. If doctors determine that the astronauts’ bodies have become too fragile for nine-Gs then a slightly higher perigee, up to 61.5 km may be used. These higher perigees will add several days to the mission since the outbound apogee may extend to the Moon’s orbital distance. Additionally, an extra 2,600 kg of habitat maneuvering thruster fuel can be planned to slow the capsule down to an additional 1 km/s or to further steer to a predetermined recovery point. Depending on the reentry angle, a 1 km/s reduction in reentry velocity can reduce G-loading by approximately 20 percent.

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As shown above, this deceleration causes a significant reduction of energy. That difference in energy is predominately transformed into heat energy. Dragon’s blunted cone will cause a lot of drag and heating. Assuming a coefficient of drag of 0.188 and a mass of 6,000 kg, Dragon’s ballistic coefficient is 2968 kg/m2. This can be used to determine how much heating our reentry phases may use or if one is more significant than the other. Per the following figure, the maximum heating occurs well below the skip altitude of 56-60 km. Our primary concern for re-entry heating will occur during the actual reentry, when much slower velocities will lead to lower temperatures.

: Altitude of the spacecraft’s max heating rate (m) : Atmospheric scale height = .000139 m-1 for Earth 3 : Atmospheric density at sea level = 1.225 kg/m : Spacecraft’s ballistic coefficient (kg/m2) : Re-entry angle

Altitude Where Dragon's Max Heating Rate Occurs 14000 12000 12521 10000 8000 7644 6000 4913 Altitude (m)

Altitude (m) Altitude 4000 3106 2000 1844 961 0 373 36 1 2 3 4 5 6 7 8 9 Reentry Angle (x 10 degrees)

Figure 4: Dragon Re-entry Dragon’s heat shield is described as a Phenolic Impregnated Carbon Ablator (PICA). As the carbon vaporizes, due to the heat, it absorbs up large amounts of heat energy and protects the spacecraft. This is the primary means of re-entry temperature control and has been successfully used on capsules including Apollo and Soyuz. PICA, as a technology, has performed very successfully in space. The Stardust mission used a PICA shielding and safely survived its 12.8 km/s re-entry velocity with a heat flux of 1,200 W/cm2 (Tito et al., 2013). This is shy of the maximum predicted re-entry velocity or heat flux for re-entry of the Inspiration Mars mission; however, its proven tests in space significantly bolster its readiness level. Perhaps, not

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surprisingly, the PICA material developed by SpaceX is called PICA-X. It has been tested with NASA researchers at the AMES research facility at temperatures of up to 3,450oF (National Aeronautics and Space Administration [NASA], 2014). Launch and Assembly designed to support any Mars mission must have three critical things in common. First, they must be able to launch very large masses to low Earth orbit, at a minimum, if not directly to an Earth escape orbit. Second, they must have very large payload bay dimensions to launch the human habitat and capsule systems. Lastly, they must be human rated so that the astronauts are able to carry on their mission to Mars with little interruption. There has been discussion regarding all-in-one designs and multiple rockets launches supporting a mission spacecraft that is assembled-in-space. The all-in-one single large rockets, from the Inspiration Mars mission perspective, would make the mission simpler; however the actual mission spacecraft is likely too small to be practical for a 501 day Mars mission. One significant advantage of this concept is that the all-in- one design would be assembled and well tested prior to mounting in the payload section of the rocket. Another advantage would be significantly reduced costs. The assembly-in-space design concept could use a series of already well tested rockets to separately preposition the habitat capsules and astronauts in low Earth orbit. The habitat would be fully linked together and prepared for the astronauts boarding. Once final systems checks are made, the mission would proceed to Mars. The advantage of this design is that the technology either currently exists or the bases of the technology already exist and are well tested. It does present more risk factors in that it will be very difficult to fix any assembly problems that occur in low Earth orbit. The costs of multiple rocket launches significantly drive up costs. Most importantly, our habitat and reentry vehicles that comprise our mission essential requirements are what our rockets must be designed to support. Team JASPer’s Inspiration Mars proposal calls for a Bigelow BA330 habitat as the primary habitat system and the Dragon capsule as the reentry vehicle and alternate supporting habitat. Orion’s J-2X rocket booster is what will accelerate Chariot to a trans-Mars injection orbit. Chariot is expected to have approximately 55,000 kg dry mass. It will take an estimated 313,204 kg fuel to support a plane change and 4.841 km/s boost to TMI. Chariots physical size and mass will require either an enormous “all-in-one” rocket or a mission design to link the Chariot’s component modules in space.

The “all-in-one” rocket concept: The first of these designs, the all-in-one design is perhaps the easiest to review because the technology simply does not exist yet. Man has never left the “gravity well” of Earth. Historically, the was as close as man has ever come. It assumed very high risk for every mission. Several Apollo astronauts died and there were many close calls. Of the 12 manned Apollo missions, Apollo 1 and 13 account for a one-in-six mission failure rate. The Soviet Union never even could get their manned Moon missions off the ground. The Space

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Shuttle regularly launched and recovered the largest all-in-one system but it never went beyond low-Earth orbit. The Space Shuttle program has been disestablished but technologies from within it should be considered in the Inspiration Mars mission. The Constellation program developed technologies for a large all-in-one system but that was cancelled. Constellation included development of the Ares V rocket. The Ares V was a precursor to the (SLS) design. The first Space Launch System launch called “Exploration Mission 1” is currently on track, as scheduled, for 2017. This launch design is expected to be able to launch 69,000 kg. A follow up more capable design “Exploration Mission 2” is expected to launch in 2021 can carry the Orion capsule and four American astronauts. This follow up design is expected to launch 129,000 kg to orbit (National Aeronautics and Space Administration [NASA], 2012). A few highlights of SLS readiness include: - The main engines, RS-25s, are on track to meet integration milestones in 2016. The J-2X engines, for trans-Mars injection, are completing developmental testing (NASA, 2013). - Boosters are on track for construction to begin in FY 2014 in support of Exploration Mission 1 (NASA, 2013). - The spacecraft and payload integration has completed Payload Fairing (5 m and 8.4 m) conceptual analysis. It has also had successes in completing two production Orion Multi- purpose Crew Vehicle Stage Adapters for Exploration Mission 1 and for structural testing (NASA, 2013). The Exploration Mission 1 design will be capable of launching a spacecraft to Mars; however, the additional requirements for long duration human life support are not supported. The primary advantage of SLS over all other rocket designs is its payload fairing size. It is the only design physically wide enough to support a Bigelow BA330 capsule. Its most significant negative attributes are its cost. At an estimated $500 million per launch, it is more than twice as costly as every other U.S. heavy lift rocket design. To support a manned Mars mission, either the larger Explorer Mission 2 design production must be accelerated or the assembly-in-space design must be pursued. Instead, Team JASPer recommends the assembly-in-space concept.

The assembly-in-space concept: The second mission design is the assembly-in-space concept. This is the chosen method for Team JASPer’s Inspiration Mars mission design. This concept must account for both heavy lift rockets and human rated rockets. These heavy lift rockets must be large and powerful enough to launch the habitat modules and powerful enough to launch the enormous masses of TMI rocket fuel to orbit. These include Atlas V and the Delta IV Heavy. It will also likely include the Falcon Heavy and SLS in the near future. The second is for human rated rockets to get the astronauts to orbit. This will include Falcon 9, Falcon Heavy, and SLS series rockets. Due to International Traffic in Arms Regulation (ITAR) complexities and short design to build schedule, as much as possible should be made in the United States. For this reason, only United States rockets are evaluated. - The Delta IV Heavy variant is capable of lifting 22,997 kg to low Earth orbit. Its payload bay is 4.5 m in diameter and approximately 13 m long. These dimensions are not symmetrical though as the nose section becomes narrower. For the lower end of this

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payload section, it is capable of carrying a centaur rocket (United Launch Alliance [ULA], 2007). This payload size is short both in length and width to launch a Bigelow BA330. Cost: FY2018 US$235 million.

- The Atlas V Heavy lift rocket is capable of lifting 29,400 kg to low Earth orbit. Its payload bay is 4.5 m in diameter and approximately 16 m long. Like the Delta IV, these dimensions are not symmetrical though as the nose section becomes narrower (ULA, 2010). This is long enough to launch a Bigelow BA330 but too narrow. Cost: FY2018 US$211 million.

- Falcon 9 can lift 13,150 kg to LEO. It can lift a Dragon spacecraft and trunk to LEO. It can also carry an alternate fairing that has dimensions of 5.2 m in diameter and 13.1 m in length. (SpaceX, 2013). Although the Falcon 9 fairing is wider than Delta IV and Atlas V, it is still too narrow and length too short to launch a Bigelow 330. Its payload mass is also the smallest. Its primary advantages are that it will be human rated by the proposed launch date and its relatively low cost. Cost: FY2018 US$62 million.

- Falcon Heavy is expected to launch 53,000 kg to LEO. It also advertises the ability to launch a 13,200 kg payload to Mars. It is expecting its first test launches as soon as 2014. Its payload fairing is the same as that for the Falcon 9 (5.2 m x 13.1 m) (SpaceX, 2013). Though wider than Delta IV and Atlas V, the Falcon Heavy is still too narrow and length too short to launch a Bigelow 330. It is a good balance of payload mass, cost, and expected human rating. Cost: FY2018 US$135 million

- Space Launch System-1. Per “all-in-one” design reference. It can launch 69,000 kg and has a payload size 8.4 m in diameter and 17.0 m in length. It is the only rocket that, as currently designed, can launch a Bigelow 330. Cost: $500 million. Due to its payload capabilities and relative low cost, the Falcon Heavy will be the primary used for the Chariot spacecraft assembly. However, in order to launch the Bigelow BA 330, the Falcon Heavy will require an expanded payload fairing in width by 1.5 m and its length by 0.6 m. This will allow for the launch for the BA330 as well as the two associated liquid hydrogen fuel tanks needed for the orbital change and TMI burns. Delta IV is being adapted to accommodate an Orion test launch so precedent for rocket modification in support of manned space requirements has been made. Chariot assembly: Any successful manned Mars mission will be an undertaking of the likes no space program has ever undertaken before. It was not until Apollo 11 before Apollo achieved mission accomplishment. It took 10 Saturn tests and test flights to prepare for the successful Moon landing. This Inspiration Mars proposal calls for a much more demanding feat. Specifically, two Falcon 9 and nine Falcon Heavy rockets are required for this mission. One standby Falcon 9 and one standby Falcon Heavy are also requested in order to ensure adherence to construction timelines. Costs of those standby rockets are not calculated into costs because they will be

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returned to SpaceX for other missions if unneeded. Assembly in space will be accomplished by automated docking systems and further supported by the International Space Station’s Canadarm2 and astronauts’ extra-vehicular activity (EVA). There are two primary objectives of assembly; manned habitat assembly and TMI booster and fuel tank assembly. The habitat assembly and trans-Mars injection booster assembly may be conducted simultaneously, vice sequentially. Simultaneous assembly will reduce rocket launch costs because they can each be loaded into payload modules if size and mass constraints are not exceeded. Chariot extends further than the reach of Canadarm2 if it is attached to ISS perpendicularly, as most modules are. It will be easier to assemble the Chariot if it is built parallel to the ISS. Habitat assembly modules will use ISS docking systems as tested by SpaceX’s resupply missions and with Bigelow’s tests on its Bigelow Expandable Activity Module (BEAM) capsule in 2015. Canadarm2 and EVA will support additional assembly and engineering check requirements. Canadaarm2’s maximum length is 17.6 m so as capsules rendezvous with ISS the Canadarm2 will assist in these modules being affixed to the Chariot Structural Support Element (CSSE). Canadaarm2’s maximum mass support of 116,000 kg is more than five times the dry mass of any module and will be of tremendous use in assembly. Canadarm2’s camera system will be used to perform visual checks of all module linkages and external systems. Astronauts onboard ISS will perform multiple EVA events to link modules together and to perform operational checks of all external systems. EVA missions will also be conducted to support the final “in-flight” fueling missions for the J-2X. Three assembly mechanisms will need to be developed to support this assembly project. This construction project’s spatial relationship that is parallel to the ISS means that the docking mechanism for BA330 will need some kind of additional tunnel support that will connect its entry point to the ISS for easy access. Second, a structural support element will need to be constructed from early on. This CSSE will extend from the front of the BA330 to the aft end by the engine exhaust. It is a rail that will serve to ease assembly by allowing massive equipment to make controlled movements from Chariot’s head to stern in conjunction with and/or beyond where the Canadarm2 can reach. Attached will be designed a mechanized motor that will drive along the CSSE to allow the astronauts to move massive equipment easily and safely. It will also serve as a base to mount additional equipment on and to ensure the BA330 maintains structural support during the TMI burn. The CSSE design must incorporate a remotely activated disconnect to detach the Dragon capsule from the J-2X engine modules and the BA330 prior to reentry.

Assembly Phase Zero: July 2017 (Prepositioning – No actual in space assembly) Requirements: - Ground operations ready. - 1st Falcon 9 launch with DragonRider and J-2X engine. - Canadarm2 checked for 100% mission readiness. - Phase One payloads pre-staged prepared to launch. - Manned Falcon 9 mission will be standing by with two astronauts and EVA gear to provide additional construction support if needed.

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Phase Zero will begin by ensuring all ground stations are fully manned, communications equipment is operating properly, and that all Phase modules are fully ready pre-staged for flight. Phase Zero will include launching a Falcon 9 with a crewed DragonRider to preposition 3,000 kg of crew, initial supplies, and BA330 preparation materials at the International Space Station. These supplies may include air, water, food, exercise equipment, Chariot engineering equipment, and additional support equipment required for ISS EVA missions. The J-2X engine will also accompany the DragonRider for use in assembly during phase 2. The Falcon 9 will be the only phase zero launch vehicle. At the end of Phase Zero, astronauts onboard ISS will have made ready two working areas for the simultaneous assembly of habitat modules and TMI booster modules. Additionally, all personnel and needed materials will be in place to link these modules to the ISS, immediately beginning the engineering procedures. The DragonRider used for this phase will be used for the Chariot spacecraft’s reentry capsule. Phase One should be pre-staged in payload fairings in rocket hangers and prepared for launch on 24 hours’ notice. The Habitat Assembly Phases will be referred to separately once Phase Zero is complete. Assembly Phase One: August 2017 (Initial assembly begins) Requirements: - 1st Falcon Heavy launches BA330 in a modified fairing. - Chariot Structural Support Element constructed and J-2X is linked - Phase Two prepared for launch. Habitat Assembly Phase One will be the launch of a BA330 habitat inside a modified Falcon Heavy rocket. The BA330 will rendezvous and link with the International Space Station through both internal guidance and use of Canadaarm2. Astronauts within ISS will inflate and inspect the BA330’s readiness for the Inspiration Mars mission. The DragonRider launched in Phase Zero will be attached to the BA330 along with the J-2X engine. Additional external structural support elements will be installed on the Bigelow capsule to ensure it will withstand the TMI boost acceleration and to add on additional external solar panels, radiation sensors, and communications equipment. This support element will be referred to as the CSSE. BA330’s mass and length is within the reach of Canadaarm2 to support all checks for the BA330. The TMI Boost Assembly Phase One will consist of the CSSE. This framing support will provide the holding mechanisms for the fuel tanks to be launched in Phase Two. Extensive EVA will be expected for the assembly of CSSE. Mounted to this CSSE will be a small motorized vehicle that will aid astronauts in the positioning of massive equipment. This vehicle may also have a camera mounted on it that can be used to examine the spacecraft’s structural integrity both during assembly and during the mission. Phase Two: October-November 2017 (Dates adjustable but defined by assembly of fuel systems) Requirements: - 2st Falcon Heavy launches with 26,861 kg of liquid hydrogen and the 6,500 kg tank in the modified fairing.

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- 3nd Falcon Heavy launches with 26,861 kg of liquid hydrogen and the 6,500 kg tank in the modified faring. - 4th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - 5th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - 6th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - 7th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - 8th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - 9th Falcon Heavy launches with 43,247 kg of liquid oxygen and the 1,166 kg tank. - Assembly of all tanks using the CSSE and EVA operations conducted by the ISS crewmembers. Phase Two consists of launching all required fuel for the Chariot orbital plane change and trans-Mars injection. Two separate fuel tanks containing a total of 53,723 kg of liquid hydrogen will be launched initially. Due to the amount of volume required, approximately 759,000 L, these tanks will be launched in the modified Falcon Heavy fairing designed originally for the BA330. Six fuel tanks containing a total of 259,482 kg, approximately 227,500 L, of liquid oxygen will be launched shortly after and assembled as indicated by the scaled Chariot diagram below. A 4.83 to 1 liquid oxygen to liquid hydrogen fuel ratio will be used to optimize the J-2X burn efficiency (Nortardonato, 2012). The combined fuel tank masses make up 6.4 percent of the 313,204 kg of fuel required. Note: At the end of Phase Two, Chariot will have all its primary components in place and enough fuel and power to complete operational checks.

Phase Three: December 2017 (Final operational checks) Requirements: - Final assembly of fuel tanks, fuel lines, and J-2X rocket. - Continue testing all systems and pre-launch systems checks and monitor cryogenic fuel. - A spare Falcon Heavy will be maintained and prepared to launch with fuel or other engineering equipment.

The ISS crew will ensure pre-mission checks are completed and that spacecraft assembly is being completed per schedule. At the completion of this phase, the CSSE will be extended past the BA330 habitat and Dragon capsule and will support the J-2X engine and its fuel tanks. Once installed, the CSSE will reinforce the integrity of the Bigelow capsule to withstand the TMI boost as well aiding Canadarm2’s maximum extension limitations during assembly. At the end of this phase, all aspects of life support systems, laboratory equipment, navigation, power, and communications should be operational. Note: At the end of Phase Three, Chariot will be fully assembled, evaluated, and ready for its mission crew.

Phase Four: Late December – Early January 2018 Requirements:

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- 2nd Falcon 9 launches with Inspiration Mars crew, remaining necessary supplies. - All internal systems pre-flight checked by Inspiration Mars crew. - External systems pre-flight checked using a combination of Inspiration Mars and ISS crew EVA. Canadarm2’s cameras will also be used. Phase Four is the final phase assembly phase. Chariot’s two astronaut crew will arrive on board for the first time and will conduct their pre-mission checks of the spacecraft. This will include both internal equipment checks as well as EVA engineering inspections. Note: The end of Phase Four will lead directly to the order to proceed to Mars. Chariot will separate from ISS under its own power. TMI boost will begin. Chariot Habitat Design Interior The interior design of the habitat will be designed to not only take care of basic life functions but, also a space where mission productivity is stimulated. Our design aims to achieve this by making the habitat not just a place to live but a place where the crew enjoys living. The BA330 provides the best platform for our crew’s habitat due to its volume of space, which will allow for a comprehensive interior design. The 330 m3 of space is a luxury no space crew has seen to date and will allow us to provide all the necessities, redundancies and comforts for the two-manned crew. The habitat locations will be referenced as Overhead, Deck, Port, Starboard, Forward, Aft and Core (Imhof, et. al, 2010). We further divide the space into living and working areas. As with previous space habitats, we assume that the spaces will be designed to conform to rack positions. This will ensure ease of installment and flexibility to change and evolve the design. The Overhead section will be primarily the living area for the crew, including their crew quarters, toilet, galley and communal living space. The Deck section will be the work area, to include the lab, gym, greenhouse and maintenance/medical area. While specific purposes, such as the Crew Quarters or toilet, must have permanent locations, our habitat will be designed to allow the crew to modify dimensions through the use of collapsible walls and screens, in order to allow the crew to modify their space dimensions for the certain task at hand. The remainder of this section will discuss the design and purpose of the living, work and storage areas.

Crew Quarters (CQ) Each crew member will have their individual CQ unit. The CQ design will be modified from NASA’s latest CQ design found on the ISS. CQ1 and CQ2 will be located in the overhead living area. One unit will be placed forward and one aft to provide personal space and privacy. Each CQ will be a personal home that ensures privacy for sleep and personal time. Structurally, the CQ will provide radiation protection, adjustable lighting, noise isolation, ventilation, electronic connections for laptop and music use, as well as storage for personal items. Communal Area and Galley The communal living area, which will be a shared space with the galley, will be located in the overhead section between the two CQ units. In this area the crew can enjoy their downtime

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together, utilize Earth communications, media library and games, as well as have their meals. This space will have adjustable walls and screens to adjust the space to the crew needs and activities. Toilet The design will be modified from the latest model found on the ISS. If weight and space allow, there should be a toilet located in the overhead and deck sections with the plumbing located in the core section. The toilet in the overhead section will have a larger volume to provide needed space for the crew’s hygiene time, while the toilet in the deck section will be for immediate bodily functions only. Lab Scientific experiments and studies carried out during the mission are not only necessary to further space exploration, they will be essential to provide the crew with daily stimulating activities. The lab area will have the resources and be adjustable to accommodate the predetermined experiments to be executed, as well as basic experiments created en-route by the crew.

Greenhouse Fresh foods and a green living space will provide the crew physiological and psychological benefits, as well as improve the air quality of the habitat. The greenhouse will be located in the deck section to ensure the lighting is not a nuisance for the crew during their downtime. It is our hope that the current Lada greenhouse design will evolve to a system that will provide space to grow a variety of plants and vegetables. Lada is the current greenhouse model currently hosted on the Russian section of the ISS. It has had significant success in providing the ISS crews vegetables, down-time activity and an improvement in air quality and smell. The Space Dynamics Laboratory reports its specification as: - A control Module (center): 9.5” x 7” x 9.5” and approx. 14.75 lbs. - Two independent vegetation modules: 9” x 21.5” x 6.5” and approx. 11.5 lbs. - Two water reservoirs. With both chambers and lights on, the system uses approximately 200 Watts. We recommend expanding the vegetation modules for our Chariot habitat as space and power limits allow. A greenhouse has many benefits, as discussed in our Psychological Design section. The crops are regularly tested to ensure microorganisms do not contaminate the air or water supply. If possible the greenhouse should be able to expand into the surrounding storage units as space becomes available. With minimal windows, having a green wall for the crew to enjoy and look at will be an invaluable benefit to their quality of life.

Gym The multi-functional exercise equipment is essential to the health of the crew during this long-term mission. At least three free-floating pieces of equipment will be located in the deck/forward section of the habitat. The Interim Resistive Exercise Device (iRED) will provide the crew muscle strengthening exercise options. The Cycle Ergometer Vibration System

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(Bicycle) and Treadmill Vibration Isolation System (TVIS) will provide the crew their cardiovascular exercise. The total mass to house these three sets of equipment will be 943 kg (iRED – 408 kg, CVIS – 27 kg, & TVIS – 408 kg). Maintenance & Medical Area Having a space where the crew can fix equipment or themselves is very important during this remote, self-sufficient mission. The deck section, between the lab and gym, will have a space that can be easily sanitized to carry out sensitive procedures and adjusted to accommodate the required space for execution. While most information and manuals will be accessed electronically, this area will maintain emergency hard copy checklists for crew reference. An emergency kit will also be available that will be able to hold over 2 people. The kit will include lifesaving medication and equipment (Clement, 2005) such as sutures, bandages, and an automated external defibrillator (AED).

Storage Maximizing the utility of every space and surface in the habitat is essential. Every location not designed for a specific purpose will be utilized for storage. Storage areas should be well labeled and adjustable to accommodate the needs and resources of the crew. Personal items to the maximum extent will be kept in the CQs. One item that poses a significant storage problem is trash. For other space missions, trash is stored and returned home for disposal or stored and eventually detached to burn up in the Earth’s atmosphere. Burning trash will not be an option for this mission. Detaching trash is also not an option to ensure the is not violated. Decontamination and storage of trash in the habitat is the safest option. The trash storage area will be located in the habitat’s core. We recommend that a contingency option to launch trash into space be available to the crew, should sanitary and health constraints require such action.

Core The core of the habitat will primarily be the location for the mechanics of the habitat: electronic panels and connections, toilet plumbing, and life support systems such as air and water recycling. In addition, it will be used to store crew gear and trash. The EVA lock is also located in this section. If space and weight allow, access to the few windows, which are only located at the forward and aft of the core should be made for the crew to comfortably access them. Radiation Shielding One of the major reasons all but a select few astronauts, from the Apollo program, have not ventured beyond low Earth orbit is the risk of radiation exposure. Close to home, they are protected within Earth’s magnetosphere. Even if there is a radiation burst, astronauts can quickly return to the Earth’s surface for protection and medical treatment. Inspiration Mars projects must overcome the effects of radiation exposure through: 1) an understanding of the sources and types of radiation, 2) understanding the effects of radiation on human biology, and 3) applying that aforementioned knowledge with proper engineering and medical practices to protect the

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astronauts from that radiation. Our primary habitat’s, the BA330’s, hull is 40 cm thick and made of Kevlar, Vectran, and water. The Kevlar and Vectran are strong materials that will maintain the airtight integrity of the spacecraft as well as to mitigate the impacts of meteoroids (Fildes, 2006). The water will protect from the effects of radiation. One unique aspect of our mission design is to use our “dead” or used up post TMI J-2X engine and its fuel tanks as much as practical for additional passive shielding. The most common radiation threat is from ionization, where radiation imparts energy to an atom within the body and that energized atom then emits an electron. The higher the energy in the radiation, the more damaging it can be. Increased energy in this radiation causes increased potential for cell damage. Compared to the photon levels of energy, such as what we might wear sun block for at the beach, the protons and heavy nuclei present in solar flares and galactic cosmic radiation (GCR) are much larger and more massive. They have much greater ionization capability (Buckey, 2006). There are three most likely sources of radiation that astronauts will be exposed to. These are solar particle events (SPE), galactic cosmic rays, and Earth’s Van Allen Belts. The most significant aspect of our radiation mitigation plan is through passive shielding and by avoiding it as much as possible by launching during a solar minima (NASA, 2013). Flying our missing during a minima could lead to a 5-10 percent increase in GCR but the severity of the increased 501-day GCR dose is much less than that of even a single SPE (University of Delaware, 2013). Solar radiation is emitted from our Sun. There is a steady stream of this radiation in the solar wind. It is historically consistent and thus a factor that engineers and doctors can anticipate. Bursts of radiation from solar particle events are more challenging to plan for. These events are often associated with terms more commonly referred to as solar flares or solar storms. These storms’ radiation consists mostly of protons that arrive at Earth after approximately eighteen minutes and Mars in twenty-six minutes. Once the Protons arrive, their levels can be expected to increase for four to six hours (Buckey, 2006). Electromagnetic radiation arrives at Earth as many as ten minutes faster and can be used as a preceding warning indicator of potential proton bombardment. Current manned space missions receive ample shielding from the Earth’s magnetic field and the spacecraft. The Inspiration Mars mission must account for proper shielding in the absence of the Earth’s magnetosphere. Strong solar particle events can still penetrate to the interior of the spacecraft. This implies that radiation levels within the inside of the spacecraft should be monitored (Buckey, 2006). In addition to shielding, avoiding solar storms is an even more efficient proactive approach. Inspiration Mars proposed 2018 launch occurs near a solar minima. This historically reduces the probability of hazards from solar particle events. Unfortunately, the lessened intensity of the solar winds allows more galactic cosmic rays to threaten our astronauts.

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Figure 5 (An update on the predictions for solar cycle, as of 11 November, 2013. http://solarscience.msfc.nasa.gov/images/ssn_predict_l.gif)

Figure 6 (A correlated inverse relationship between solar cycle activity, as measured with sunspot numbers, and intensity of galactic cosmic rays. http://neutronm.bartol.udel.edu/modplot.html) Galactic cosmic radiation emanates from distant stars and galaxies and from all directions. GCRs are threatening because they represent a steady stream of more massive particles than solar flares. Eighty-five percent of these GCR particles are protons, the same as in solar particle events. Fourteen percent are alpha particles, and the remaining one percent represents nuclei as heavy as iron. As GCRs are halted within the skin, “they transfer their

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energy to the skin’s atoms. The high energy iron nuclei, for example, can be absorbed by aluminum in the skin. These aluminum atoms then fragment into lighter nuclei that penetrate the cabin as a secondary radiation. Generally, more dense materials are poorer shields due to this secondary radiation effect and lighter materials such as hydrogen or water are more effective” (Buckey, 2006). The energy levels and flux of radiation is important; however, even more important to survival needs is how much is absorbed by the body. Sieverts (Sv) are the units that scientists use to describe and measure radiation exposures and exposure limits. These Sv limits are cumulative, not necessarily the affects from a single dose. For example, depending on age, female astronauts are limited to 0.4 to 1.7 Sv over a 10 year span while males are limited to a range from 0.7 to 3.0 Sv (Buckey, 2006). These long-term effects of radiation are difficult to determine. The effects of DNA damage could lead to cancer but that could occur 2 – 20 years after the exposure. “Large temporary radiation doses, as would be expected from a solar flare, produce immediate, dose- related effects. Most of the effects are due to failure of cells to reproduce. Therefore, the effects of radiation occur initially in those organ systems whose normal functioning required active cell reproduction, such as bone marrow and the gastrointestinal tract” (Buckey, 2006).

Radiation type Energy Fluence Composition Dose delivered Solar particle 10 – 100 MeV 1010 particles/cm2/sec Mainly protons Variable, can event (no be 1 – 5 Sv shielding) Galactic cosmic 300 – 3000 4 protons/cm2/sec; 85% Protons, 0.5 Sv/year radiation (no MeV 0.4 helium 14% helium shielding, solar ions/cm2/sec nuclei, 1% minimum) heavier nuclei Table 2(Characteristics of the predominant types of radiation, Buckley, 2006)

Dose (Sv) Predicted physiologic effects 0.1 – 0.5 No obvious effects, some minor changes in blood counts 0.5 – 1 Fatigue, transient reduction in lymphocyte and neutrophil counts; 5 – 10 % experience nausea and vomiting for approximately 1 day; no deaths anticipated 1 – 2 50% reduction in lymphocytes and neutrophils; 25 – 50% experience nausea and vomiting for 1 day; no deaths anticipated 2 – 3.5 75% reduction in circulating blood elements; most experience nausea and vomiting; loss of appetite, diarrhea, and minor hemorrhage also seen; death in 5 – 50% of those exposed 3.5 – 5.5 Nearly all experience nausea and vomiting on first day followed by fever, hemorrhage, diarrhea, and emaciation; death of 50 – 90% within 6 weeks; survivors convalesce for about 6 months Table 3 (Direct "deterministic" effects of radiation, Buckley, 2006) The challenge in shielding against lies herein: the success of the mission may depend more on protection against solar flare types of events; however, the long term health of the crews

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depends more on protection against galactic cosmic radiation. Protection measures for the crew fall into three general categories; passive shielding, active shielding, and medicine. Passive shielding requires no additional power. It predominantly relies on the spacecraft’s skin and insulation. It is usually described as having “thickness” defined by grams per square centimeter (Buckey, 2006). “A thin layer of a material such as lead might provide the same density as a thick layer of a less dense material like water. This does not always hold true though. For example, lead might protect a human from x-rays; however, heavy nuclei from galactic cosmic rays would likely strike the lead and cause secondary radiation in the form of neutrons and gamma rays. One shielding strategy for high-energy particles is to alternate heavy and light materials to provide protection. Low-density materials that contain a high percentage of hydrogen molecules (such as polyethylene or water) are ideal choices as shielding materials for interplanetary flight” (Buckey, 2006). The BA330 will be surrounded by 2.5 inches of water as radiation shielding (Bigelow, 2011). Aluminum is a common metric for shielding protection so water’s protection must be converted to an equivalent thickness of aluminum. Water’s radiation shielding is 1.64 times as effective as aluminum (Churchill, 1997). Therefore, 2.5 inches of water is equivalent to an aluminum shield 9.8 cm thick. This equates to 65-70% effective shielding against in heavy ions and protons from both SPEs and GCRs. Per Tables 3 and 4 above, 2 Sv from the most severe solar particle events would affect the astronauts. They would be temporarily sick but would recover and complete the mission. Two additional means of passive shielding could be used. An additional “storm shelter” could be constructed within the BA330 habitat that the astronauts could stay for the hours required for the solar storm to . This shelter could be made of aluminum and water. Also, the J-2X engines and its fuel tanks used to boost Chariot to its transfer orbit can remain attached and also be an effective shield. More calculations on further reducing radiation levels will be expected when engineers design the fuel tanks and fittings. Active shielding incorporates some type of magnetic field that surrounds and protects all or a portion of the spacecraft against charged particles. This mimics the shielding the Earth’s magnetosphere does for astronauts in low earth orbit. GCRs have such high energies that even active shielding does little to mitigate them. Electromagnetic shields do appear to have strong capability to shield against SPEs (Eckart, 1996 and Spillantini, 2010); however their technology readiness level (TRL) for a spacecraft of this size is very low. Long-term support would require significant development and engineering that is likely beyond the scope of this Inspiration Mars proposal. Finally, it would also draw valuable power. We will not be employing active shielding. So far we have addressed mitigating radiation exposure to our astronauts; however, we cannot escape the reality that our astronauts WILL be exposed to radiation. There are two means of monitoring astronauts’ radiation exposure: physical dosimeters and biodosimetry. Film badges are one type of a physical dosimeter. They can measure dosages of different types of radiation. The weakness of physical dosimeters is that they do not actually measure the effects on the body. For this, some type of biodosimetry tests are needed. One type of test that we will seek to include is for chromosomal aberrations in lymphocytes in peripheral blood. “The presence of chromosomal aberrations may also be correlated with the subsequent development of cancer. Chromosomal aberrations are measured in blood samples before and after spaceflight” (Buckey,

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2006). Sensors adapted from other spacecraft will also monitor localized radiation for comparing the space radiation environment to the internal habitat radiation environment. These sensors may include MAVEN’s Solar Energetic particle, Langmuir Probe and Waves, and Solar Wind Electron Analyzer instruments. Once bodies are exposed to radiation the science of medicine comes to play. “Radiation causes biological damage by producing free radicals and other reactive molecules in tissues” (Buckey, 2006). Antioxidants can be taken via pill or consumed as food in order to minimize the damage. “Antioxidants react with damaged molecules, repair them chemically, or react with chemical intermediates before they damage key biological molecules” (Buckey, 2006). The naturally occurring antioxidants may only protect against very low doses of radiation. These antioxidants include cysteine, glutathione, vitamin C, vitamin E, and selenium. Drugs, often taken via intravenous injection, tend to have significant side effects and are less tested; however, they may be more effective against stronger doses of radiation. They include amifostine, diethyldithiocarbamate, and tempol (Buckey, 2006). Radiation protection is the first line of defense for protecting the bodies of our astronauts. It requires little of them yet provides that vital protective shell that protects their cells against damage that would likely kill them if not accounted for. Chariots radiation shielding is sufficient to ensure survival through the mission against short-term solar storms and long term GCR effects. There is much more required though. Our astronauts’ physiology and psychology are addressed with a similar level of concern. The end state of all this is that our astronauts survive the mission and go on to continue living a healthy life. Physiological Constrains In addition to radiation protection and life support systems, the physiological and human psyche go hand in hand to ensure the mental health and well-being of the crew are met. In this section, muscle atrophy and bone deterioration will be the main focus followed by the psychological preparedness of the crew. The attention will be placed on the countermeasures Team JASPer will be implementing. Muscle atrophy occurs when activity in the muscle groups that require opposition to gravity are reduced (Kanas & Manzey, 2010). Typically the lower extremities such as the quadriceps, calves (tibialis anterior, gastroenemius, and soleus) and gluteus medius and maximus are affected the most due to the need for gravitational resistance to remain strong. Similarly, bone deterioration or demineralization occurs when there is a decrease in mechanical load affecting the weight bearing bones in the legs and hips (Kanas & Manzey, 2010). Demineralization of calcium, where bones, “derive their structure and strength” (Clement, 2005) are hindered as this process can be compared to the elderly suffering from osteoporosis (Kanas & Manzey, 2010). Countermeasures will be to implement an exercise regime and potentially ingesting medication. First, a similar workout regime will be utilized as the ISS where at least two hours of exercise are executed for both aerobic and anaerobic exercising. The Cycle Ergometer Vibration System (CEVIS), Interim Resistive Exercise Device (iRED) and Treadmill Vibration Isolation System (TVIS) will be the exercise equipment housed inside the BA300 habitat. The workout

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regime will consist of at least 2-3 hours of exercise per day, which can be altered by the physical trainers on Earth. This will be dependent on their physical progress. Medication is another possible countermeasure that could be used by ingesting Vitamin E and Amino acids. Vitamin E can also benefit with minimizing radiation damage of low dose rates, while Amino acids promote the stimulation of net protein synthesis that are decreased during long-duration spaceflight (Buckey, 2006). Since the biggest factor for bone deterioration is calcium loss due to a lack of stress on the weight bearing bones, ingesting calcium not more than 1000 mg/day is recommended. To caveat this, too much calcium can produce kidney stones so deciding to take the risk is something to consider. Monitoring the two-crewmember team’s rate of muscular and bone loss will be a challenge that will need specific attention to ensure rapid deterioration does not occur. There are a few physical traits that are tall tell signs of significant muscle/bone loss, which are weight change, changes in anthropometric measurements, and changes in leg circumference (Buckey, 2006). Also a high level of calcium in the urine will also indicate the bones are demineralizing. Chemical and physical changes in the body will be easier to monitor; however, the mental health factor can also play a role against the musculo-skeletal systems. Stress, under-nutrition, hormonal and oxidative stresses are some examples that have been observed. Psychological Considerations

Providing an environment conducive to the psychological needs of the crew will be essential for mission success. Ensuring the crew is taken care of, psychologically, is equally as important as providing their basic life support needs. Providing for the crew’s psychological needs will allow them to not only complete the mission but, enjoy the journey, returning home happy and healthy. As we stated earlier, the BA330 has been chosen due to its large volume which can accommodate our crew’s activities and resources. This space and availability of resources is critical to the crew’s psychological well-being. While stress and problems during a long-term, isolated mission cannot be avoided, our mission design will ensure the crew is provided with the motivation, communication, entertainment and relaxation needed to overcome the psychological difficulties they may encounter. In order to prepare for this mission, we must apply the lessons learned from past long- term, isolated and dangerous missions. Previous long-term space missions, polar expedition teams and Mars 500 provide noteworthy examples in order to anticipate the needs and challenges of our crew. These examples are analogous to our mission in that our crew will also be self- reliant with delayed and limited communications back to home base and rescue is most likely not an option. Previous missions have demonstrated that eventually the newness of the mission will fade and a crew can become subject to boredom, anxiety and decreased productivity. Having a purpose and being productive daily is very important for crew motivation and morale. Having space and resources to enjoy downtime is also necessary for the crew. We have applied these lessons to design a mission, habitat and crew which will ensure overall success. Psychology of the design

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Having a view which the crew can use to enjoy, note progress and which stimulates thinking will be important for their mental health and can be provided in a variety of ways. While windows are limited in the design, access to an external view should extend beyond video camera feeds. It is important this access be convenient and comfortable for the crew. Use of light colors and the ability to easily manipulate the dimensions of spaces will help provide a sense of having a large space to live in. It is important that the crew feel in control and comfortable with the habitat and its resources. The crew should be able to provide personal touches for the engineers to incorporate into the CQ design to make the habitat feel like home. Having a green space, provided by the greenhouse area, will also be a great resource for the crew. Plants not only improve the quality of their diet and air supply, they also provide a visual stimulus of life and the passing of time as a variety of harvests cycle through.

Recommended Resources:

The most important addition to the habitat to ensure mission success and psychological stability is a well screened and adaptable crew. The crew must be physically and mentally prepared to encounter mission and health issues which they can overcome on their own. Having technical training in medicine or engineering would be invaluable. Personalization and making the habitat feel like home has been discussed throughout this section. The crew should be able to bring pictures and personal items up to a designated weight that will not pose logistical or sanitation risks. A selection of media items, which can be updated, will provide the crew entertainment, resources to carry out daily activities and platforms which ground control can better communicate with and monitor the crew behavior. Media platforms should include flat screen TVs, personal laptops/tablets and mp3 players. The use of virtual reality and Google Glass technology could also greatly benefit the crew recreationally and with their daily tasks. Multiple and easily accessible connections throughout the habitat are required for crew-crew, crew-ship and crew-Earth communication. Each crew member should also have a psychological support person or team back on Earth. This is not a psychotherapist but an aid that can assist with communications and logistics back at home and provide an overall assurance that the crews needs on and off the ship are being taken care of. Developing facilities that provide the communication, entertainment, stable work environment and personal space necessary for crew morale is imperative, and will greatly lower opportunities for conflict or boredom. However, even the best facilities will go to waste if the selected crew does not possess the traits necessary for such a demanding mission. It will be dependent on the crew to properly utilize the resources, develop techniques for conflict resolution and maintain motivation from launch to recovery back on Earth.

Chariot Environmental Control and Life Support Systems The Chariot Environmental Control and Life Support System (ECLSS) design must provide the required human physiological needs for the entire duration of the 501 day mission, as well as adapt to any unforeseen emergency situations to ensure the survival and safe return of the two Chariot crewmembers. The following section describes the Chariot life support system design using the, Advanced Life Support Research and Technology Development Metric – Fiscal

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Year 2005, as a basis for the included state of the art and advanced system designs. This metric has been used by the National Aeronautics and Space Administration as a tool to assess the Agency’s progress and will be the primary document to support the Chariot spacecraft sub- system design (Hanford, 2006). Assumptions Bigelow Aerospace is a privately owned and operated corporation developing inflatable habitats to provide affordable spaceflight to national space agencies and corporate clients (Bigelow, 2013). Due to the limited information published by Bigelow Aerospace regarding the BA330 spacecraft, the Chariot sub-system designs will be based off of technology currently in use onboard the International Space Station. These sub-systems are assumed to be similar in design to the BA330 ECLSS and will be included in the 20,000kg estimated mass. Therefore only the additional supplies required for this specific mission will be included in the final projected Chariot mass. According to Bigelow Aerospace, the BA330 is designed to support six crewmembers for three month duration (Bigelow, 2011). Adjusting this data to account for a two person crew, rather than a crew of six provides enough life support supplies to last approximately nine months. This nine month duration consumes 54% of the total 501 day mission duration. Based off the metabolic requirements listed in table 1, we can calculate the estimated components assumed to be included within the BA330 and the additional supplies required for the total 501 day Mars Free-Return Mission.

Required Crew Metabolic Input Interface Average Crew-Member per Day Total 2CM 500d Crew Body Mass 70 kg 140 kg

Air (Hanford, 2006) Oxygen Consumed 0.835 kg/CM-d 835 kg

Water (Bobe, 2007) Drinking & food preparation 2.20 kg/CM-d 2200 kg Water in food supplied 0.5 kg/CM-d 500 kg Personal hygiene 0.2 kg/CM-d 200 kg Flush water 0.3 kg/CM-d 300 kg Water for Oxygen generation 1.0 kg/CM-d 100 kg Shower/laundry water 6.0 kg/CM-d 6000 kg

Food (Hanford, 2006) Daily food consumed 1.086 kg/CM-d 1086 kg Table 4 (Crew Metabolic Requirements)

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Air The metabolic consumption of oxygen required throughout the mission for the two crewmember occupants is approximately 835 kg as indicated by Table 1. The oxygen will be supplied to the habitat primarily from the electrolysis of water using solid polymer technology. Additionally, high pressure tanks will deliver a secondary method of oxygen and nitrogen supply as well as provide backup in case of emergency operations and cabin leakage indicated in Table 2. The other byproduct of electrolysis is hydrogen (H2). This hydrogen will be supplied to a Sabatier carbon dioxide (CO2) assembly that will remove unwanted levels of CO2 from the environment. The Sabatier reaction of CO2 and H2 produces methane (CH4), which will be vented overboard, and water which will be routed to the Oxygen Generation Assembly (OGA) for use in the electrolysis reaction (Zubrin, 1996). A Four Bed Molecular Sieve will also be incorporated into the design to remove additional CO2 form the Chariot atmosphere. The Trace Contaminant Control System (TCCS) uses activated carbon to remove trace gases from the environment, which pass through High Efficiency Particulate Air (HEPA) filter assemblies for particulate removal. The Major Constituent Analyzer (MCA) and fire detection/suppression system provide monitoring for air contaminants and combustible products (Hanford, 2006). These systems work in unison to ensure all unwanted trace contaminant toxic gases are removed from the habitat atmosphere.

ECLSS Atmosphere Control (Hanford, 2006) Parameter Unit Value Cabin Volume [m^3] 330 Maximum Leak Rate [kg/d] 0.01344 Internal Atmosphere Nitrogen Partial Pressure [kPa] 77.1 Oxygen Partial Pressure [kPa] 21.3

Water Vapor Partial Pressure [kPa] 1.2

Carbon Dioxide Partial Pressure [kPa] 0.4 Table 5 (ECLSS Atmospheric Control) Water The Chariot spacecraft design requires approximately 9300kg of water to supply the daily metabolic needs and associated spacecraft subsystem operations throughout the 501 day mission. In order to significantly reduce the amount of water supply carried onboard the Chariot spacecraft, a Water Recovery System (WRS) will be incorporated into the sub-system design. All wastewater will be collected in a wastewater tank where it will pass through a liquid separator (MLS), a particulate filter, and two Multi-Filtration (MF) beds in order to remove any unwanted contaminants within the water. The water then passes through a Volatile Removal Assembly (VRA) to remove additional volatile gasses from the filtered water. The final phase of the water regeneration involves the purification of the water through the use of an Ion Exchange

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(IX) bed. The water then travels through a Microbial Check Valve (MCV) to return any unfiltered contaminants back to the wastewater tank and finally into the potable H2O tank where it can be distributed throughout the spacecraft. The use of the Water Recovery System reduces the initial required water mass to approximately 595 kg. This TRL 9, flight proven, equipment has been incorporated into the International Space Station and has resulted in an overall 93% water recovery rate (Bell, 2006). Although the WRS has proven to be efficient, the Chariot spacecraft will carry an additional 717 kg of water using 18 Contingency Water Containers (CWCs) for unforeseen circumstances. Based off of the 2.2 kg/CM-d of water required to support the basic physiological needs, the crew would be able to survive an additional 163 days in the event of a complete WRS failure.

Food The crewmember metabolic energy requirement for daily mission activities is 11.82 MJ/CM-d (Miller, 1994). This metabolic requirement corresponds to a daily crew food allowance of 1.086 kg/CM-d consisting of a mix of fresh, dehydrated, and full-water preserved food (Hanford, 2006). The total, as shipped, food weight is approximately 1415 kg including 262 kg of packaging. A 5% food reserve was added to account for any unforeseen contingencies and to ensure a slightly excess supply of food upon mission completion. Assuming a stocked BA330 includes the four month supply of food for six crewmembers, an additional 717 kg will be included onboard the Chariot spacecraft prior to the trans-Mars injection. Waste The Chariot waste sub-system will use storage to contain all discarded products from the spacecraft to include: trash, used filters/cartridges, fecal matter, and brine from the urine and water processing system. Crewmember urine will be recycled through the use of a Vapor Compression Distillation (VCD) processor. The unusable brine will be sent to the waste storage, while the liquid water will pass through a Gas/Liquid (GL) separator and enter the water filtration system where it can be processed and returned to system as potable water.

Thermal The Chariot internal thermal control system includes an Avionics Air Assembly (AAA) responsible for removing heat from electrical components, a Common Cabin Air Assembly (CCAA) which cools, dehumidifies, and circulates the cabin air, and a Thermal Control Subsystem as a primary method of cabin temperature regulation through the use of cold plates and heat exchangers (Hanford, 2006). Human Accommodations The human accommodations required for the 501 day mission are provided in Table 3. These values have been included by the Inspiration Mars Feasibility Analysis and account for one male and one female crewmember (Tito et al., 2013). However, in order to drastically reduce the clothing required for the crewmembers throughout the mission, a form of laundry

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washing/drying equipment has been included, thus reducing total human accommodation supply to a combined total of 306.02kg.

Human Accommodations (Tito, 2013) Item Mass (kg/CM-d) Total 2 CM 500d (kg) Clothing 0.03 30.02 Toilet Paper 0.028 28 Gloves 0.007 7 Dry Wipes 0.013 13 Detergent 0.058 58 Disinfectant 0.056 56 Paper 0.077 77 Tape 0.033 33 Feminine Health (1 female) 0.008 4 Total: 0.31 306.02 Table 6 (Chariot Human Accommodations) Master Supply List & Sub-System Design The master supply list for the Chariot spacecraft is included in Table 4. As mentioned previously, approximately 54% of the two crewmember mission duration supplies are assumed to be included in the BA330 habitat as well as the sub-systems required for spacecraft operation. The air component listed as 1708 kg includes N2/O2 tanks as well as the electrolysis equipment. Therefore the required additional supply column includes: contingency air/water, the remaining 46% food supply, and crewmember clothing/personal provisions required for the 501 day mission. The subsystems listed in Table 8 and illustrated in figure 7 are a combination of State of the Art (SOA) and Advanced technologies currently in use on the ISS or being designed for use prior to the 2018 launch date in order to meet the demands of the Inspiration Mars Mission.

Chariot Master Supply List Component BA330 Estimated Supply (kg) Required Additional Supply (kg) Air (Tanks & SPE) 1708.6 300 Water (Potable Storage) 595.54 717 Food (Supply & Packaging) 754 651 Clothing Not Included 30.02 Personal Provisions Not Included 276 Greenhouse Not Included 43.75 Total: 3058.14 2017.77 Table 7 (Chariot Master Supply List)

Chariot Sub-System Design Technologies (Hanford, 2006) Subsystem Technology TRL (NASA, 2012) Mass (kg) Power (W) Air (Advanced & SOA) 2,633.35 4,343.36 Gas Storage

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Nitrogen Storage High Pressure 9 1,028.73 0.00 Oxygen Storage High Pressure 9 300.17 0.00 Contingency High Pressure 300.00 0.00 Atmospheric Control System Atmospheric Pressure Control ISS 9 119.40 70.50 Atmosphere Revitalization System Carbon Dioxide Removal Molecular Sieve 9 179.14 536.06 Carbon Dioxide Reduction Sabatier 7 143.53 148.59 Oxygen Generation Electrolysis 9 378.86 3,288.88 Gaseous Trace Contaminant Control ISS 9 85.81 194.35 Atmosphere Composition Monitoring ISS 9 54.30 103.50 Sample Delivery System ISS 9 35.11 0.00 Fire Detection and Suppression Fire Detection System ISS 9 1.50 1.48 Fire Suppression System ISS 9 6.80 0.00 Water (SOA) 4,070.03 1,126.62 Water Recovery System Urine, Waste Water Collection System ISS 9 4.55 4.00 Water Treatment Process ISS 9 2,463.74 919.74 Urine, Hygiene, & Bring Storage Tanks ISS 9 181.57 17.80 Microbial Check Valve ISS 9 5.72 0.00 Process Controller ISS 9 36.11 156.18 Water Quality Monitor ISS 9 14.07 4.72 Product Watery Delivery System ISS 9 51.73 3.44 Water Storage Potable Water Storage ISS 9 595.54 20.74 Contingency ISS 9 717.00 0.00 Food (SOA) 1,415.50 0.00 Food Supply ISS 9 1,140.40 0.00 Food Packaging ISS 9 275.10 0.00 Human Accommodations (Advanced) 435.04 833.33 Clothing ISS 9 30.02 0.00 Laundry Equipment Washer/Dryer NA 4 80.00 633.33 Detergent NA 4 5.27 0.00 Greenhouse ISS 9 43.75 200.00 Miscellaneous Items ISS 9 276.00 0.00 Thermal (SOA) 329.08 913.74 Temperature and Humidity Control Common Cabin Air Assembly ISS 9 118.08 530.52 Avionics Air Assembly ISS 9 12.40 175.00 Atmosphere Circulation ISS 9 9.80 61.00 Atmospheric Microbial Control ISS 9 100.00 0.00 Internal Thermal Control System ISS 9 88.80 147.22 Waste (SOA) 213.35 14.00 Solid Waste Collection ESDM 9 36.36 14.00 Solid Waste Treatment Storage 9 176.99 0.00 TOTAL: 9,096.35 7,231.05 Table 8 (Chariot ECLSS)

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Figure 7 (Chariot Sub-System Design: Hanford, 2006) Engineering Busses Supporting engineering busses onboard Chariot covered within this mission design include power, communications, and navigation. TRL level 9 programs are desired for critical elements of our spacecraft. Though Chariot’s individual components may be through TRL 9, the spacecraft as a whole integrated system will be TRL 8, at best, as evaluated during the assembly. Dragon and the Bigelow BA330 capsules may conduct final TRL readiness evaluations during the assembly phase. In addition to these core systems, several lesser developed TRL readiness programs will be evaluated throughout this mission. One complicating factor for mission design is the limited proprietary information released (or more likely not released) that drives many estimations in Team JASPer’s Inspiration Mars mission design. Chariot will be capable of autonomous operations as directed by onboard computers and/or astronaut input. Chariot will

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also be capable of being completely controlled by ground control operators. None of these busses can exist without power, making the power generation and distribution busses top priority. Next in importance is Chariot’s internal navigation and attitude control system. Finally, redundant communications architecture is required for continuous monitoring of Chariot status as well as command guidance. Thermal regulation is not addressed here but is a critical factor for follow on examination due to our trajectory’s outbound and inbound path that flies inside 1 astronomical unit (AU). BA330 has its own internal thermal regulation system. We have flown several spacecraft, such as Messenger closer to the Sun than our mission trajectory is intended to go, so there are several successful missions to study from.

Power Bigelow has not released details on its power system other than to note its independent power system will be comprised of solar arrays and batteries (Bigelow Aerospace, 2014). These power sources will provide power to all systems, including: communications, navigation, and life support systems. Several important estimation factors can be deduced by examining the designs posted by Bigelow Aerospace and comparing those to the given dimensions of the habitat. The BA330 never has more than two solar arrays, each appearing to be approximately 15 m2. These arrays appear to be on a rotatable swivel, essential for maintaining optimal solar energy. SpaceX has not released data on Dragon’s solar panels; however, there are also two panels that each appear to be approximately 12 m2. Mars will be near perihelion during our fly-by at approximately 1.4 astronomical units (AU). Solar intensity decreases with the radius squared so a doubling of solar panel surface area is required to ensure equivalent power production at Mars as is generated at Earth. Another possibility, though Bigelow has not released any information is to increase the efficiency of its solar panels. We make the assumption that additional panels are possible but redesigning them for increased efficiency is not possible for the purposes of this Inspiration Mars mission. Total power generation within the BA330 habitat is not provided. The MAVEN spacecraft has 12m2 of solar panel areas generating 1,150 watts of power. The Mars Reconnaissance Orbiter has 20 m2 of solar panel area generating a total of 1,000 watts. Using these as guides, we estimate that the BA330 is capable of producing between 1,500 - 2,875 watts at a distance of Mars. These same estimations calculate that Dragon may be able to generate 1,200 - 2,300 watts. The DragonRider Factsheet states it will produce 1,500 – 2,000 watts so this estimation method appears valid (SpaceX, 2013). This combined BA330 and Dragon power generation is estimated to be 2,700 – 5,175 watts. This is much lower than the Space Shuttle’s fuel cell power generation of 7,000 watts so further coordination with Bigelow is needed to confirm its total power generation (NASA, 2014). Additional solar panels will be designed into Chariot to provide redundancy and additional power for mission requirements. Dragon must maintain its power source for reentry so it will need to be further studied whether the power advantage of using the array throughout the mission is more advantageous that keeping it folded and protected until needed. Four additional MAVEN type solar arrays will be attached to produce an additional 4,600 watts of power. This

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will result in a total of approximately 7,300 – 9,775 watts of power produced. Using an estimate of 4,600 additional watts required and approximately 110 W/kg of solar power as a planning estimate, we estimate our additional solar panels will have a mass of 42 kg (Wertz & Larson, 1999). Using that same 4,600 watts of additional panels and approximately $2,000 $/W, we estimate that that these additional solar arrays will cost $9.2 million. Batteries will provide a secondary power source. They will be essential while Chariot is being shadowed by Earth while it is being assembled and shadowed by Mars during the fly-by. The time that batteries will need to carry power is dependent on the planet’s angular radius and orbit period.

Given: = Planet’s angular radius viewed from space h = Orbital altitude REarth = Earth’s radius / RMars = Mars’s radius

Knowing Earth’s and Mars’s angular radius, we can solve for the maximum time the spacecraft is in each planet’s shadow and dependent on batteries. Given: TE = Maximum time of eclipse (minutes) = Earth’s angular radius P = Orbital period (minutes) = Earth at 400 km = 93 minutes / Mars at 250 km = 112 minutes

It is evident that the fly-by of Mars will place a greater demand on batter power than the assembly phase in low Earth orbit. This becomes the driving factor on our battery life design. Nickel-cadmium have less specific energy densities than nickel-hydrogen designs so our heaviest mission design planning calls for the nickel-cadmium designs for mass calculations. We estimate that it will take 283 kg of batteries to store 8,500 watts of power for use during the fly-by, assuming 30 W hr/kg (Wertz & Larson, 1999).

Navigation The BA330 design will have its own independent system to support navigation, docking, and other maneuvering activities (Bigelow Aerospace, 2014). Bigelow has not released the true

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design and capabilities of its navigation system. Team JASPer thus assumes that its navigation system will both need to be augmented and provide an opportunity to test emerging interplanetary navigation technologies. Supporting attitude control systems will be modeled after MAVEN’s avionics systems. The MAVEN spacecraft uses ring laser gyros to detect spacecraft inertial motion and four reaction wheels to provide fine attitude control for proper spacecraft orientation (NASA, 2013). Most Mars missions that have failed in the past two decades have come from being lost in route or a failed approach to the planet, therefore only the most proven systems will be considered. The navigation bus on Chariot is modeled after MAVEN’s navigation bus. In addition to the navigation systems built into the BA330, Chariot will have two star tracker cameras that feed data into a computer which will then provide spacecraft attitude information to other navigation systems. Additionally, because power is one of the most critical parts of this mission, Chariot will have two Sun sensors. These Sun sensors will help point its solar panels toward the Sun in case of an error in its primary navigation systems (NASA, 2013). Precise navigation around Mars is critical because fine adjustments to altitude will be required to keep Chariot from missing its window and venturing too far from Mars into deep space or too close to Mars atmosphere and losing too much velocity for a safe return to Earth. For this, there will be one additional navigation system that is critical during the Mars fly-by. This is the Optical Navigation Camera. This camera was successfully used in Mars Reconnaissance Orbiter (MRO). It takes pictures of the Martian moons Phobos and Deimos during the month prior to arrival. The positions of the moons are compared to background stars and allow engineers to accurately determine the position of the orbiter in relation to Mars (Jet Propulsion Laboratory, 2014). Additional experimental navigation systems such as pulsar based navigation systems may be flight tested as well.

Communications Communications with ground control communications will be important to ensure a safe mission for the Inspiration Mars crew. It is also important for morale and outreach programs. NASA’s Near Earth Network (NEN) and Deep Space Network (DSN) will be instrumental to supporting this mission as will Chariot’s redundant and easily maintained radio suite. These communications systems will support an array of functions but the most important ones will be: acquiring telemetry data from our spacecraft, transmitting commands to our spacecraft, tracking our spacecraft’s position and velocity, and voice and video transmissions with ground control. The interactions between our astronauts and ground control will be that of mutual support and redundancy. Chariot will have autonomous flight control ability; however, ground control will also be able to seamlessly assume complete control of the spacecraft’s various engineering busses. Ground control will also provide oversight of proper spacecraft operations, trajectory, and habitat viability. Time delays, up to 20 minutes will drive an increased requirement for the astronauts’ autonomy; however, the link to ground control stations is critical. The Apollo missions required extensive support from ground control. The fate of the crew from Apollo 13 would surely have been disastrous had good communications with ground control been maintained. Priority should be added to Near Earth Network and Deep Space Network systems in the order of: 1) supporting manned missions and 2) critical event communications including

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spacecraft separation, orbit insertion, powered flight, entry/descent/landing, critical maneuvers, and fly-bys (NASA, 2010). Radios carried onboard Chariot will include S-, X-, and Ka- band transceivers. Communications planners should consider this a Category B mission (for distances over 2 x 106 km). Primary communications should be in the 7/8 or 7/32 GHz bands due to congestion in the 2 GHz deep space band. NASA also directs deep space missions to operate in the Ka-band at 31.8 – 32.3 GHz. NASA’s Space Communications and Navigation Office recommended that all missions launching beyond 2016 should use Ka-band. Deep space missions should use X- and Ka-bands (32 GHz) and near Earth missions should use S- and Ka-bands (26 GHz) (NASA, 2010). DSN antennas are limited to two simultaneous spacecraft communications. DSN antennas can receive both spacecraft but only one uplink frequency can be transmitted. If sharing antennas with other spacecraft, transitioning uplink frequencies back and forth at the is time consuming. It takes approximately 30-minutes to transfer two-way operations from one spacecraft to another. It is unlikely that all deep space missions will be on hold through the duration of the Inspiration Mars mission so these delays are something that should be expected as the norm. The Mars Reconnaissance Orbiter transmitted data back to Earth for up to eleven hours per day for 700 days. MRO was a non-manned mission so our team assesses that it is reasonable to request 24-hour links to be maintained for this manned deep space mission. The total data storage for the MRO was approximately four terabytes or what would need to be stored on over 6,500 compact disks (NASA, 2014). This Inspiration Mars mission is a test flight for follow on manned missions so every detail will be evaluated by engineers. This is too much data to have hard copies stored onboard Chariot. It will need to be transmitted to Earth for archiving. Use of NASA’s antennas is also very costly. It will cost approximately $557.00 per NEN pass (<= 30 minutes per pass) in any of the S-, X-, or Ka-bands. Standard multiple access rates for the Space Network are $29.00 per minute for forwarding and $14.00 per minute for return services. Deep Space Network costs are estimated according to the below equation. AF = RB [AW (0.9 + FC / 10)] Where: AF = weighted Aperture Fee RB = contact dependent hourly rate, adjusted annually ($1057/hr. for FY09) AW = aperture weighting: = 0.80 for 34-meter High-Speed Beam Waveguide (HSB) station. = 1.00 for a single 34-meter station (i.e., 34 BWG and 34 HEF). = 2.00 for a two 34-meter station array. = 3.00 for a three 34-meter station array. = 4.00 for a four 34-meter station array (70-meter equivalent). = 4.00 for 70-meter stations. FC = number of station contacts, (contacts per calendar week). Thus, a cost estimate can be had for use of NASA’s ground based radio systems. A total of approximately $30 million will be the costs for sharing the NEN and DSN for the 501-day mission plus the costs during assembly and low Earth orbit testing.

Mission Phase Subnetwork Hours No. Tracks No. Weeks Cost Total Costs

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Requested Per Per Week Required Per Segment Track 1 Assembly NEN .5 112 12 $557 per $748,608 pass 1 System 1x34-meter station 1 56 12 $2,220 / $1,491,639 Testing hr. 2 TMI Boost NEN .5 1 1 $557 per $557 pass 3 Cruise to 1x34-meter array 1 168 32.5 $2,220 / $12,119,562 Mars hr. 4 Fly-by of 1x34-meter station. 1 168 1 $2,220 / $372,960 Mars Ka-band hr. 4 Fly-by of 1x34-meter station. 1 168 1 $2,220 / $372,960 Mars X-band hr. 5 Cruise to 1x34-meter station 1 168 39.5 $2,220 / $14,731,920 Earth hr. 6 Reentry NEN 1.5 112 1 $557 per $62,384 pass Total Costs $29,900,590 Table 9: Communication Cost Estimates Continuous communications requires mitigation planning against likely degradation. Assessed degradation factors here include radio reliability, line of sight limitations, and the ability to transmit through dust within the solar system. Each of these bands has their own advantages and disadvantages. The X-band and S-band are commonly used for deep space communications and many antennas within the deep space network receive and transmit on these bands. X-band is most commonly found on spacecraft with limitations to a single band. Their high frequencies provide faster transmission of data than S-band. S-band frequencies have the advantage of a low space loss of transmission. Ka-band, with the highest frequencies, is recommended for command and emergency telemetry. It is also the least susceptible to solar effects. Ka-band disadvantages are that it is very directional and must be aimed at where the target will be when the signal arrives. It also has the highest space loss and this precludes use of it as part of an omni-directional antenna for emergency transmission (JPL, 2013). Team JASPer assumes that radios will fail on this 501-day mission and will make redundancy essential. Redundancy will include multiple radios available at all times for both voice and data from the near-Earth to the deep space environment. Chariot will have one S-band radio and one Ka-band radio for operations in the near Earth environment and one X-band and one Ka-band radio for operations in the deep space environment. One additional Electra UHF Transceiver (EUT) relay radio will also be installed for flight test purposes. This EUT would be a copy of one currently installed in the MAVEN spacecraft. Its mission is to communicate with and relay information between other Mars spacecraft, ground rovers, and Earth. This is a minimum total of five radios: one S-band, one X-band, two Ka-bands, and one EUT. While antenna will be mounted external to the spacecraft, their components such as amplifiers, will be maintained and monitored inside. This will allow astronauts easier access in troubleshooting faults. Optical communications equipment may be tested during the mission. One additional positive aspect of this mission’s timing is that Mars will not be opposite the Sun, in conjunction, during this window. Similar to other Mars mission discrepancies, this could have led to several weeks of communications gaps. The preceding conjunction is July 27, 2017 and the proceeding one is September 2, 2019.

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Astronaut Selection NASA along with foreign space agencies, have developed a process to select the best candidates for short duration (SDM) and long duration missions (LDM). Due to the extreme environment of space and long duration isolation crews experience, the ability to work as a team player, be resilient when times are tough, and present maturity are some of the key factors that are taken into account when building a crew. Once the crew is selected, they need to become one cohesive team to ensure the mission is completed (team compatibility). The process NASA and other space agencies have used since the Mercury program is called select-in and select-out. Select-in and out is the process of testing and finding positive and negative traits of potential astronauts. Examples of select-in are “relevant operational skills, training, maturity, stress tolerance, and the ability to get along with others” (Kanas & Manzey, 2010). Select-out characteristics are, “predisposition for mental illness, psychopathological characteristics, and difficulties with interpersonal relationships” (Kanas & Manzey, 2010). These present the foundation of qualities NASA is looking for before delving into the compatibility of crew dynamics. In the past, NASA mainly focused on negating candidates that possessed the select-out traits until the Shuttle era (Kanas & Manzey, 2010). Due to the increase in duration in space, crew compatibility became a growing factor of importance especially in the Salyut and eras (Kanas & Manzey, 2010). In addition to crew composition, select-in traits became and still are the methods of selecting an astronaut crew. For this mission, Team JASPer is interested in sending a male and female for the 501 day mission to Mars and back. The best approach is adopting the current method utilized by NASA in using the select-in and some select-out traits. The current NASA approach is “based on analyses of available research and anecdotal information from analog environments, as well as expert ratings from 20 Russian, European, and American astronauts and mission support experts” (Kanas & Manzey, 2010). Reference the table below for the critical psychological factors required for long-duration and short-duration space missions (1 - most important).

Selected Sample Factor Proficiencies Criticality for LDM Criticality for SDM

Mental/emotional stability Freedom from mental disorder, self-control, 1 2 self-confidence Performance under Ability to perform under stressful conditions threat to life stress and stressful conditions, flexibility and adaptability, ability to 2 1 cope with limited personal space

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Group living skills Group living and interaction skills, adaptability to crew 3 7 diversity, multicultural adaptability Teamwork skills Conflict resolution and cooperation, priority of 4 4 team over personal goals, followership skill Family Issues Ability to cope with prolonged separation from 5 6 family and friends Motivation Achievement motivation, intrinsic work motivation, perseverance, goal orientation 6 8

Judgment/decision Exercising sound making judgment, situational awareness and vigilance 7 3 Conscientiousness Responsibility, attention to detail, integrity 8 5 Communication Interpersonal communication skills 9 9 Leadership capability Team leadership, effective resource management, 10 10 accountability

Table 10: Astronaut Selection Another key point in crew selection will involve the mission training and operations practice drills after initial and alternate crew selection is completed. As found in basic personality and social desirability testing norms, the true test of social compatibility lies in the cooperation and interaction with coworkers under stressful conditions. Candidates tend to exaggerate characteristics they believe are favorable for selection, a reaction called Social Desirability (SD) (Sandal et. al, 2005). In a review of 55 human interaction and teamwork studies, researchers “noted that every study found the more time team members spent training together, the fewer conflicts and conflict-related performance deficiencies the team members experienced. Also, in a review of simulation based training practices, observers noted that more benefits can be accrued from team performance if teams are encouraged to practice complex and emergency simulations together than if team members are trained in simulations in random groups.” (Schmidt et. al, 2009). For long duration missions, it is imperative crew members have sufficient time to train and operate with each other to build up a cohesive team effort before departing.

Training On average, crews train for their specific mission 2-3 years prior to spaceflight. Due to the nature of the mission to Mars, choosing a crew by late 2015 is ideal to ensure they are fully capable to provide mechanical maintenance, medical maintenance, and have knowledge of

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troubleshooting measures for example. An alternate team of a male and female will train simultaneously with the primary crew in the case they need to step up. If something did prevent the primary group to go into space, the swap will not be one-for-one; the swap will be the whole team to ensure the cohesiveness of the team remains. The beginning of their training will consist of reading the Chariot flight manuals and completing computer-based training, as was done for the ISS and Space Shuttle (Clement, 2005). Then they will be prepared for in-flight procedures and the timeline of the Mars flyby. During this training, the 2-person crew will also need to maintain a certain level of health. The main focus for medical support is prevention (Clement, 2005), and applying countermeasures to ensure the astronauts are healthy when they go into space. Previous human missions, to include Antarctic missions, will be used to identify the common in-flight medical issues to ensure the crew knows how to react should a medical issue arise. These training measures will aid the crew in being flexible and reactive when unfavorable or emergency conditions arise.

Chariot Economics The following cost model analysis for Team JASPer includes all Chariot spacecraft launch, assembly, and operational hardware to include: launch vehicles, habitat structure/sub- systems, upper stage rocket design, fuel supply, crewmember provisions, and contingency equipment. Due to the similarity of mission operational requirements for all Inspiration Mars participants, and mission costs to include: support facilities, Earth based hardware/software, employee salaries, communication architecture, and associated ground mission operations will not be included in this analysis. Therefore, the following cost model will be specifically tailored to the components providing the foundation of the Chariot spacecraft. The costs included have been adjusted to account for inflation to Fiscal Year (FY) 2018. Additionally, estimated Technology Readiness Levels have been applied to each component of the spacecraft and adjusted appropriately based off of the published Space Mission Analysis and Design standard deviation chart (Wertz and Larson, 1999).

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Chariot WBS

Launch Segment Ground Space Segment Segment

Ground Operations and Habitat Design Trajectory LEO Assembly Launch Vehicles Facilities & Maintenance Support

Equipment & ECLSS TMI Burn ISS Docking Payload Mass Mainenance Software

Communication Internal Mars Orbital Habitat Inflation Fairing Design & Navigation Spares Configuration Maneuver & Configuration Architecture

Mission Physiological Navigation & Fuel & Tank Crew Docking Logistics Command Constraints Orbit Correction Design Operations

Safety & Radiation Earth Re-entry Management Mission Protection Assurance

Habitat Contingency Public Subsystems Planning Interaction

Figure 8: Chariot Work Breakdown Structure

Space Segment Cost modeling for the space segment of the Team JASPer proposal will primarily consist of habitat design, subsystem integration, and additional crew provisions to meet the demands of the 501 day Mars mission. A SpaceX DragonRider attached to a Bigelow BA330 combine as the primary habitat for the Chariot crew, which accounts for the majority of the space segment costs. The necessary supplies and equipment to support the crew will also be included in the following cost estimate. The Bigelow BA330 inflatable habitat is designed to include all Environmental Control and Life Support Systems and associated sub-system components. The projected cost of the BA330 is approximately FY2018 US$139.33 million (Wang, 2013). Due to the current unpublished developmental status of the BA330, the resulting Technology Readiness Level of the habitat is estimated to be in the 3-5 range, which is associated with a moderate level of relative risk. This TRL is supported with the successful launch and operation of the Genesis 1 and 2 habitats that have been in orbit since 2006 and 2007. The standard deviation for cost at this Technology Readiness Level is approximately 10-25% (Wertz and Larson, 1999). Using the

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conservative 25% standard deviation estimate, the BA330 can be projected to cost approximately FY2018 US$174.16 million. The SpaceX DragonRider crew transport and support vehicle will combine with the BA330 to form the habitat of the Chariot spacecraft. The projected cost of a crewed SpaceX DragonRider is estimated to be FY2018 US$156.05 million (Seedhouse, 2013). As with the BA330, the DragonRider is still in the developmental phase of design; however, a pressurized Dragon did successfully deliver Cargo to the International Space Station in May, 2012 demonstrating a Technology Readiness Level of 7. Two additional resupply missions have been successfully conducted, which has further demonstrated the reliability of the Dragon spacecraft. The standard deviation for TRL 6-8 is 10% resulting in a projected cost of FY2018 US$171.65 million for the DragonRider spacecraft. The internal configuration of the Chariot spacecraft will contain all crew provisions as well as additional contingency life support supplies. The custom interior components included within the Chariot design consist of individual crew quarters, exercise equipment, crew entertainment, and all other provisions necessary to satisfy the physiological and psychological demands of a 501 day confinement. The associated budget allocated to these supplies is approximately FY2018 US$35 million. These supplies, along with an assembly crew, will be delivered to the ISS via a DragonRider spacecraft, where they can then be transferred into the BA330. Six 38,000 L liquid oxygen and two 379,499 L liquid hydrogen tanks will need to be researched and developed specifically for this mission. Once in orbit, the tanks must be assembled and linked in order for the propellant transfer to occur during the orbital plane change and TMI burns. The total allocated budget set forth for the development of the tanks and assembly systems has been set at FY2018 US$146.43 million. This estimate, adjusted for the 20,000 kg combined tank weights, is based off of the Space Shuttle External Tank Contract granted to Lockheed Martin in 2007. Launch Segment The launch segment and assembly is the most complex and costly aspect of the Chariot design primarily due to the amount of propellant needed to conduct the orbital plane change and trans-Mars injection. As discussed in the assembly section, the Chariot spacecraft launch segment requires two Space X Falcon 9 and nine Falcon Heavy launches to deliver the needed supplies to a 400 km LEO orbit for assembly. The SpaceX Falcon 9 has been chosen as the launch vehicle to deliver the phase zero DragonRider supply spacecraft and J-2X engine as well as the phase four DragonRider crewed capsule. The Falcon 9 has a 10,450 kg payload to LEO capacity and is estimated to cost approximately FY 2018 US$62.10 million, averaging US$5,942 per kg (Wertz and Larson, 1999). The SpaceX Falcon Heavy will deliver the Bigelow BA330 and eight required propellant tanks. The Falcon Heavy will have a payload of 53,000 kg to LEO and is estimated to cost FY2018 US$135 million, averaging US$2,547 per kg (Seedhouse, 2013). As mentioned previously, the BA330 and liquid hydrogen tanks require a modified Falcon fairing design,

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which has been allocated US$20 million for research and development. Due to the relative low weight of the BA330 (20,000 kg) and liquid hydrogen tanks (28,861 kg/tank), the modified fairing design should not cause issues with the payload capabilities of the Falcon Heavy. Ground Segment As discussed previously the ground segment and mission costs to include: support facilities, Earth based operational hardware/software, employee salaries, communication architecture, and associated ground mission operations will not be included in this analysis. The ground segment mission costs will be provided upon request with the final JASPer mission design report.

FY 2018 Chariot Hardware Cost Estimation Segment Cost (US million) Space Bigelow BA330 $174.16 SpaceX DragonRider Cargo $171.65 SpaceX DragonRider $171.65 Chariot Supplies $35.00 LOX & LH2 Fuel/Tanks $146.43 J-2X Rocket $20.00 Additional Solar Panels $9.20 Chariot Total $728.09 Launch Modified Falcon Fairing $20.00 (SpaceX Falcon 9) X 2 $124.20 (SpaceX Falcon Heavy) X 9 $1,215.00 Launch Total $1,359.20 Ground NA Mission Total $2,087.29 Table 11: Chariot Hardware Costs Conclusion The Team JASPer Chariot spacecraft has been predicated on crew physiological and psychological accommodations required for the duration of this monumental adventure. We hope that contents within this proposal have provided a unique design that will assist in the future of manned space flight. We would like to thank the generations that have gone before us and inspired us and hope that this competition inspires many others to follow. We also thank The Mars Society and NASA for sponsoring this competition. Finally, we thank our professors within University of North Dakota’s Space Studies Department for all their mentorship and assistance throughout this project.

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