A Near-Term Mars Sample Return Spacecraft Design Utilizing Solar Electric Propulsion

Details A Near-Term Mars Sample Return Spacecraft Design Utilizing Solar Electric Propulsion A Near-Term Mars Sample Return Spacecraft Design Utilizing Solar Electric Propulsion

Jan 2002

adsabs.harvard.edu/cgi-bin/nph-data_query?bibcode=2002iaf..confe.702g&link_type=abstract

IAF abstracts, 34th COSPAR Scientific Assembly, The Second World Space Congress, held 10-19 October, 2002 in Houston, TX, USA.,

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In support of advanced planning activities for a potential Mars sample return (MSR) mission, a spacecraft design concept was developed and various mission implementation options were assessed. The baseline 760 kg wet solar electric propulsion (SEP) spacecraft was configured to transport a 20 kg sample canister from low Mars orbit (LMO) to low Earth orbit (LEO) in approximately 2 yr. A shuttle orbiter carrying a specially engineered containment vault would then be used to bring the sample and spacecraft to a safe and controlled Earth runway landing. The spacecraft architecture is largely dual-string with high projected reliability to address planetary protection concerns. Off-the-shelf or near-term system components and technologies were assumed. A nominal 10 kWe (at 1.0 AU) solar array powers two 30 cm or 40 cm xenon ion thrusters, providing efficient primary propulsion and modest propellant need. Two innovative features of the design include reaction control via xenon resistojets, allowing a single inert propellant to serve both primary and auxiliary propulsion, and use of the internal xenon tank as primary load-bearing structure. The above results in an efficient and low mass solution for return of Mars samples, as well as a compact spacecraft (1.2 m dia x 2.3 m stowed) capable of being returned to earth entirely contained within a shuttle- carried vault. Alternative mission architecture options were also explored. In Option 1, a heavy Delta II was used to inject a 1300 kg wet variant of the baseline to a slightly positive C3, allowing the spacecraft to be "self- delivered" to Mars in 2 yr. Option 2 emplaced a 1600 kg wet variant into LEO, allowing reduction in launch vehicle size to a smaller "Med-Lite" Delta II, at the expense of a now 3 yr outbound leg to Mars. Other options explored the ability of the SEP vehicle to ferry the 2900 kg Mars lander and ascent vehicle. It was found that SEP could enable an entire MSR mission to be injected on a single Delta IV class launch vehicle, though power, thruster, and propellant requirements were significantly higher than the original baseline system. This activity was supported by the NASA Jet Propulsion Laboratory, and was performed as an element of the Mars Program Systems Engineering Team (MPSET) in year 2001.

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