Space and rocket-related projects are notoriously expensive, with costs that often exceed initial budgets, especially for government-funded projects. These high costs pose a substantial barrier to expanding the capabilities of space exploration, satellite deployment, and related scientific research. To overcome these challenges, this study presents the optimization of a two-stage rocket for low Earth orbit (LEO) at a maximum altitude of 450 km and a 300 kg payload, focusing on aerodynamic and propulsion optimization to enhance fuel efficiency.

A multidisciplinary framework was applied, incorporating propulsion system analysis, and thermodynamic properties of the rocket nozzles using NASA’s Chemical Equilibrium with Applications (CEA) code. The Computational Fluid Dynamics (CFD) simulations were conducted in SolidWorks to refine the rocket's aerodynamic profile by optimizing fin geometries. The nozzle geometry was refined using Rao’s method, with MATLAB calculations applied to determine the optimal nozzle expansion ratio. NASA’s General Mission Analysis Tool (GMAT) was utilized to perform trajectory optimization and propellant mass calculations. The nozzle performance was validated through benchmarking against comparable-class small-lift launch vehicles.

The optimized launch vehicle achieves a total lift-off mass of 13,000 kg, including 11,200 kg of propellant and a 300 kg payload. The first stage produces 235 kN of thrust with a burn time of 155 s and a specific impulse of 265 s, while the second stage delivers 25 kN of thrust over 350 s with a specific impulse of 340 s. The combined two-stage propulsion system provides a total ΔV of 13.71 km/s, enabling a final orbital altitude of approximately 400 km. These findings contribute to more sustainable two-stage propulsion systems by enhancing fuel efficiency and sustainability in small-lift launch vehicles.