The Air Force Institute of Technology is studying a new upper-stage rocket engine architecture: the dual-expander aerospike nozzle. The goal of this research is to provide the maximum thrust-to-weight ratio in an engine that delivers a minimum of 50,000 lbf vacuum thrust with a vacuum specific impulse of 464 s. Previous work focused on developing an initial design to demonstrate the feasibility of the dual-expander aerospike nozzle architecture. That work culminated in a design exceeding the requirements, delivering an estimated 57,000 lbf thrust with a specific impulse of 472 s by using an oxidizer-to-fuel ratio of 7.03, a total mass flow of 121 lbm=s, and an engine length of 38 in. These results were computed in a numerical model of the engine. Current work expands the model in preparation for optimizing its thrust-to-weight ratio. The changes to the model are designed to support running automated parametric and optimization studies. Parametric studies varying oxidizer-to-fuel ratio, total mass flow, and chamber length show that a dual-expander aerospike nozzle engine can achieve 50,000 lbf vacuum thrust and 489 s vacuum Isp with an oxidizer-to-fuel ratio of six, a total mass flow of 104 lbm=s (a reduction of 14%), and an engine length of 27.9 in. (a reduction of over 25%), which should equate to a significant weight savings over the original design.