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Some of the alternatives considered here contemplate the use of the Falcon Heavy (FH) as well. FH is not used to launch any components of the crew transit vehicle itself, but is used only for water refueling operations in Earth orbit or at Mars. SLS can deliver 40 t to Mars,13 while FH can deliver 16.8 to Mars and 63.8 t to LEO.23 Although the FH is projected to be considerably less expensive than the SLS, it has not flown yet at the time of writing, and official cost figures for a FH launch with maximum payload are not available. Therefore, FH availability is not assumed to be as a cost effective alternative, and this study includes architecture options both with and without it. B. Mars Design Reference Architecture 5.0 In-Space Propulsion Challenges We address challenges to the DRA-5 mission that are direct consequences of the multi-staged propulsion approach. First is that the propulsion modules and their payload are placed into orbit separately on five SLS launches. This architecture requires several months (120 days) of on-orbit assembly time. Due to the long period spent in LEO, a reboost module is also required, to perform stationkeeping during this time.1 This module is jettisoned before Earth departure, and therefore does not hamper the Mars transfer directly. However, it masses over 48 t and, thus, fills a significant amount of payload capacity. If the crew were able to depart immediately or at least much sooner than 120 days, that mass could be eliminated or used for other purposes. Second, the multi-stage approach also increases the total dry mass for the crewed vehicle. For example, each of the three TMI modules has five RL10-B2 engines, which combined contribute 1 t per module.17 These and other sources of redundant dry mass and cost could be eliminated by using a single propulsion module. Launching a single module with as much propellant as the five propulsion modules in the DRA-5 is impossible for SLS, but it is possible instead to launch an empty or partially empty propellant tank to be filled by another launch, or even a full propellant tank that that is refueled at key points during the mission. The third challenge the DRA-5 architecture faces concerns the shelf life of the engines and propellant in the existing chemical propulsion technology. The study identifies this as a critical problem: Cryogenic fluid management (CFM) is a critical technical area that is needed for the successful development of the Mars architectures. The first and foremost challenge is the storability of LH2, CH4, and O2 propellants for long durations. Note that the longest flight of stored cyrogens is Titan Centaur-5, where the propellants were stored in orbit for a 9 hours.1 It is this third challenge that is principally behindthe decision to use water electrolysis propulsion rather than sending LOX/LH2 tanks for refueling of the RL10-B2 engines. In contrast, using water completely avoids these challenges. The storage life of water is almost indefinite, and water fueling operations are safe in comparison to cryogenic fueling. C. Electrolysis Propulsion Overview An electrolysis propulsion thruster uses hydrogen and oxygen as propellant, but stores the propellant as liquid water instead of as separate cryogenic fluids.18 The water is then electrolyzed into hydrogen and oxygen gas, which can be temporarily stored together as an already combustible mixture, or separately, for later mixture and combustion. In our experiments, this technology has demonstrated a specific impulse of close to 300 s. Other developers have achieved similar performance.19 Though the aforementioned projects are designed for CubeSats, the technology can scale up with the additional power available from solar panels on larger spacecraft. The specific impulse of this technology is significantly lower than that of cryogenic LH2/LOX engines, which can reach 450 s.20 In our experiments, the gap in performance is because electrolysis is much slower than the combustion process. So, the thruster must fire in small pulses, when it would be more efficient in a steady state burn of a longer duration. However, liquid water can be stored at lower pressures and with less insulation than cryogenic fluids (or none), and electrolysis propulsion requires no turbopumps or other complex apparatus to operate, saving considerable dry mass. Also, as detailed in the previous subsection, this technology addresses several specific concerns with LH2/LOX that are present in the crewed Mars mission examined in this paper. The electrolysis propulsion vehicle in this paper is refueled using tanks of liquid water that are sent ahead of the crew vehicle prior to the mission. This reduces the spacecraft mass during maneuvers preceding the refueling operation, and thus reduces the total propellant consumed. Together with the dry mass savings of consolidating five stages down to one that is refueled, and the dry mass saving of electrolysis propulsion over LH2/LOX engines,the proposed technology significantly reduce the total launch mass and hence the number of launches required for the mission. The next section provides an overview of several alternative mission architectures made possible by this approach. 3 American Institute of Aeronautics and Astronautics Downloaded by NASA LANGLEY RESEARCH CENTRE on January 30, 2018 | http://arc.aiaa.org | DOI: 10.2514/6.2018-1537PDF Image | Water Electrolysis for Propulsion of a Crewed Mars
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