manned mission. Both of these missions can benefit greatly from ISMU by reducing the landed mass required at the Martian surface, since the propellant for the return portion of the mission would not have to be transported to Mars. A great deal of work has been done on mission analysis with methane ISMU, but two specific missions are considered below for the sake of comparative discussion of propellants. A sample return mission might return a 1 to 10 kg sample. With a 1 kg sample return as a baseline, Ash et al. [1] determined a mass of approximately 250 kg for the Mars sample return vehicle. They found that when the mass of all the necessary tankage and propulsion systems for the return trip to LEO are included, the total vehicle dry mass landed on the Martian surface becomes approximately 750 kg. Using the rocket equation, and assuming a single stage vehicle, we can find the initial mass required on the Martian surface to return the vehicle to LEO using a single stage. Here mj is the initial mass, mj is the dry vehicle mass (750 kg), AV is the velocity increment for the mission (6454 m/s from the surface of Mars to LEO [6]), ISp is the specific impulse of the fuel (340 s with O/F=3.5, assuming methane will be used for return), and go is the Earth’s gravitational acceleration (9.81 m/s2). Hence, the mass required on Mars for the return mission is 5190 kg (4440 kg of this is propellant). If all the propellant were terrestrially derived, the entire 5190 kg would have to be landed on Mars. Assuming hydrogen fuel for the outgoing mission (I of 454 s at O/F=5.5), this requires a launch mass of 13,450 kg in LEO (using AV = 4242 m/s from LEO to the surface of Mars [6]). However, if the propellants for the return mission are produced on Mars, only 750 kg would need to be landed, and the launch mass required at LEO would be 1940 kg. This represents a launch mass savings of 85% for the ISMU case. Note that we assume all propellant is Mars derived, and nothing is brought from Earth. A manned mission with its added complexities and larger required landed mass would require a much greater initial mass, and would benefit more from utilization of local resources. Again following the lead of Ash et al. [1], a manned mission might utilize a lander similar in size and mass to the Apollo lunar ascent module. The mass of the lander, without propulsion or tankage systems, would be 2700 kg. Adding 1500 kg for the propulsion and tankage systems would bring the estimated dry landed mass on Mars to 4200 kg. Here, unlike the case considered by Ash et al. [1], a rendezvous orbit around Mars is not considered. Rather, a single stage vehicle is directly launched to LEO. Consequently, using a similar analysis to the sample return case discussed previously, the mass required on the Martian surface to return the 4200 kg vehicle to LEO using methane fuel for the return mission is 29,100 kg (24,900 kg of this is propellant). To land this entire mass on Mars the
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