Design and Optimization of Hybrid Propulsion Systems for In-Space Application

Abstract

Due to their potential for high performance, inherent safety, throttling and restart capability and low development costs, hybrid rocket motors are believed to be good candidates for in-space propulsion applications. In order to demonstrate the potential of hybrids, a design/optimization tool for in-space, single stage hybrid rocket propulsion system is developed. H2O2/paraffin-based propellant combination and a lunar mission is selected as the reference case. Spacecraft gross mass to payload mass ratio has been minimized over key propulsion variables such as chamber pressure, oxidizer to fuel ratio (O/F) and nozzle area ratio. Sensitivity analysis on geometrical constraints, fuel regression rate and nozzle erosion rate is performed. Calculations are also made with the hydrogen peroxide/hydroxyl terminated polybutadiene propellant combination, considering both single port and multiport geometries for this slow burning fuel. Major results indicate that fuel regression rate has the highest influence on the motor size and gross mass. Nozzle erosion rate also has a high influence on the performance due to the decreasing nozzle expansion ratio during the motor burn, resulting in a lower delivered specific impulse. It has been determined that oxidizer to fuel ratio shift caused by the variability of the fuel mass flow during combustion does not have a strong effect on the system performance.