Propulsion System Design for a Martian Atmosphere-Breathing Supersonic Retropropulsion Engine

Abstract

Design and analysis of an atmospheric-breathing propulsion system to land large-scale spacecraft () on Mars was performed. propulsion feasibility was analytically investigated by employing equilibrium combustion simulations, finite-rate kinetics simulations, and first-order propellant mass and inlet sizing. values (based on total propellant usage) were determined to be on the order of 120–160 s for onboard subsystems having a 10-to-1 oxidizer compression ratio. This corresponds to an of 600–800 s based on onboard fuel consumption. Although mixtures have significant ignition constraints, favorable conditions were found, yielding ignition delay times of less than 1 ms, by simultaneously employing designs exploiting both large reentry Mach numbers () and modest compression ratios. These combinations allow for combustion to occur within moderately sized combustion chambers. The first-order sizing calculations confirmed that atmospheric-breathing supersonic retropropulsion has the potential for significant mass savings relative to traditional architectures. Designs with higher oxidizer-to-fuel ratios were more mass efficient. The largest benefit was seen for small inlet area vehicles that leveraged deceleration from a terminal instantaneous burn over higher thrust throughout the trajectory.