Propulsion System Design for a Martian Atmosphere Breathing Supersonic Retropropulsion Engine

摘要

Design and analysis was performed on an atmospheric breathing propulsion system to land large-scale spacecraft on Mars. Initial feasibility of the engine was investigated analytically by employing equilibrium combustion and finite rate kinetics simulations in addition to 1st order propellant mass and inlet sizing. I_(SP) values (based on total propellant usage) were determined to be on the order of 120s-160s for onboard subsystems having a 10-to-1 oxidizer compression ratio. This corresponds to an I_(SP) of 600s-800s based on fuel consumption. While Mg-CO_2 mixtures have significant ignition constraints, favorable conditions were found, yielding ignition delay times of less than 1ms, 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 1st order sizing calculations confirmed that atmospheric breathing supersonic retropropulsion has the potential for significant mass savings over traditional retropropulsion architectures. Engines sized with an oxidizer-to-fuel ratio of 4 require half the propellant consumption for an equivalent change in velocity. Inlet capture areas of the examined atmospheric breathing propulsion systems were on the order of the corresponding entry vehicle projected area. Therefore, this study envisioned an annular inlet design, which encircled the vehicle forebody. The aforementioned analyses address some of the challenges that need to be solved in order to ultimately obtain a practical atmospheric breathing supersonic retropropulsion system for Mars descent.