Supersonic retro-propulsion (SRP) is defined as rocket propulsion used to decelerate aerospace vehicles at supersonic speed. SRP is often used as a method of high-speed deceleration on space vehicles. The main method of propulsion used in the application of SRP is rocket propulsion. Rocket engine thrust and performance changes with altitude and expansion ratio. Changing altitudes across the trajectory of a rocket affect how the exhaust plume shock waves expand. Being able to identify how different expansion ratios affect the exhaust plume flow fields would provide useful data on how SRP performance can be predicted. This research projects aims at developing a computational model for existing physical test data on SRP and extrapolating data from the model to assess how SRP would perform with different nozzle expansion ratios.

For the past four decades, Mars missions have relied on Viking heritage technology for supersonic descent. Extending the use of propulsion, which is required for Mars subsonic deceleration, into the supersonic regime allows the ability to land larger payload masses. Wind tunnel and computational experiments on subscale supersonic retropropulsion models have shown a complex aerodynamic flow field characterized by the interaction of underexpanded jet plumes exhausting from nozzles on the vehicle with the supersonic freestream. Understanding the impact of vehicle and nozzle configuration on this interaction is critical for analyzing the performance of a supersonic retropropulsion system, as deceleration will have components provided by both the aerodynamic drag of the vehicle and thrust from the nozzles. This investigation focuses on the validity of steady state computational approaches to analyze supersonic retropropulsion flowfield structures and their effect on vehicle aerodynamics. Wind tunnel data for a single nozzle and a multiple nozzle configuration are used to validate a steady state, turbulent computational fluid dynamics approach to modeling supersonic retropropulsion. An analytic approximation to determine plume and bow shock structure in the flow field is also developed, enabling rapid assessment of flowfield structure for use in improved grid generation and as a configuration screening tool. Results for both the computational fluid dynamics and analytic approaches show good agreement with the experimental datasets. Potential limitations of the two methods are identified based on the comparisons with available data. Six additional geometries are defined to investigate the extensibility of the analytical model and determine the variation of supersonic retropropulsion performance with configuration. These validation geometries are split into two categories: three geometries with nozzles located on the vehicle forebody at varying nozzle cant angles, and three geometries with nozzles located on the vehicle aftbody at varying nozzle cant angles and number of nozzles. The forebody nozzle configurations show that nozzle cant angle is a significant driver in performance of a vehicle employing supersonic retropropulsion. Aerodynamic drag preservation for a given thrust level increases with increasing cant angle. However, increasing the cant angle reduces the contribution of thrust to deceleration. The tradeoff between these two contributions to the deceleration force is examined, noting that performance improvements are possible with modest nozzle cant angles. Static pitch stability characteristics are investigated for the lowest and highest cant angle configurations. The aftbody nozzle configuration results show that removing the plume flow from the region forward of the vehicle results in less interaction with the bow shock structure. This impacts aerodynamic performance, as the surface pressure remains relatively undisturbed for all thrust values examined. Static pitch stability characteristics for each of the aftbody nozzle configurations are investigated; noting that supersonic retropropulsion for these configurations exhibits a transition point from static stability to instability as a function of this center of mass location along the axis.

Supersonic deceleration has been identified as a critical deficiency in extending heritage technologies to the high-mass systems required to achieve long-term exploration goals at Mars. Supersonic retropropulsion (SRP), or the use of retropropulsive thrust while an entry vehicle is traveling at supersonic conditions, is an approach addressing this deficiency. The focus of this dissertation is aerodynamic and performance evaluation of SRP as a decelerator technology for high-mass Mars entry systems. This evaluation was completed through a detailed SRP performance analysis, establishment of the relationship between vehicle performance and the aerodynamic-propulsive interaction, and an assessment of the required fidelity and computational cost in simulating SRP flowfields, with emphasis on the effort required in conceptual design. Trajectory optimization, high-fidelity computational aerodynamic analysis, and analytical modeling of the SRP aerodynamic-propulsive interaction were used to define the fidelity and effort required to evaluate individual SRP concepts across multiple mission scales.