• Replaced RP-1 with liquid methane (CH₄) in the first stage while maintaining Falcon 9’s chamber pressure, nozzle area ratio (ε = 16), and equivalence ratio (Φ = 0.8).

• Used NASA CEA to simulate nozzle flow properties (pressure, temperature, Mach number, and velocity), as functions of area ratio and altitude.

• Calculated thrust and specific impulse (Isp) variations from sea-level ignition to Mach 10 stage separation (~75 km).

• Computed propellant mass flow rates, tank volumes, and turbopump power requirements for both stages.

• Simulated both CH₄- and RP-1-based configurations to evaluate design tradeoffs in mass ratio, power consumption, and tank sizing.

• Achieved a lower overall mass ratio: 22.98 (CH₄ design) vs. 25.75 (original Falcon 9 configuration).

• Modeled nozzle flow behavior using NASA CEA and ATMOS data across design altitudes

• Analyzed specific impulse (Isp) trends to inform trajectory-averaged performance metrics

• Calculated turbopump power requirements based on chamber pressure, fluid densities, and mass flow rates

• Iteratively balanced tradeoffs between Isp gains and tank volume increases to optimize first-stage performance

This project involved extensive performance modeling, including calculations for thrust, specific impulse (Isp), turbopump power, and propellant tank sizing. Key parameters such as chamber pressure (Pc), area ratio (ε), equivalence ratio (Φ), fuel/oxidizer mass breakdowns, and stage mass ratios are summarized in the final design configuration.

• First-Stage Isp: 310.4 s

• Second-Stage Isp: 359.2 s
  
  

• Final Mass Ratio: 22.98

• Baseline Mass Ratio: 25.75
  
  

• Thrust: 7,599 kN

• Chamber Pressure (Pc): 95.73 atm

• Nozzle Area Ratio (ε): 16 (Stage 1), 165 (Stage 2)

• Equivalence Ratio (Φ): 0.8

*Detailed values are available in the submitted Excel design spreadsheet. A PDF snapshot is available upon request.*

• Rebalancing engine performance, mass ratio, and thrust requirements required multiple design iterations

• Adjusted design point altitude, oxidizer-to-fuel ratios, and tank volume estimates to optimize tradeoffs

• Navigated complexity between CEA output assumptions and real-world physical design constraints

• Managed tradeoffs between specific impulse (Isp) gains and increased first-stage tank size and pump power

• Achieved a reduced overall mass ratio of 22.98, improving upon the Falcon 9 baseline of 25.75

• Increased specific impulse (Isp) for both stages, while managing tradeoffs in tank volume and turbopump power

• Demonstrated ability to iteratively optimize performance metrics through simulation, modeling, and engineering judgment

• Individual contributions included: baseline system analysis, CEA nozzle modeling, configuration summary, pump power/tank sizing, and report delivery

NASA CEA · Rocket Propulsion System Design · Nozzle Flow Simulation · Mass Ratio Optimization · Tank & Turbopump Sizing · Specific impulse (Isp) Trade Studies · MATLAB Modeling · Excel Engineering Tools · Technical Reporting & Communication

This project deepened my understanding of propulsion system design by challenging me to evaluate tradeoffs in stage performance, subsystem sizing, and propellant selection. Iterating between design parameters and output metrics strengthened my intuition for how rocket stages respond to changing conditions. Working with NASA CEA introduced me to industry-standard analysis tools and boosted my confidence in interpreting performance data, thermodynamic cycles, and system-level modeling of launch vehicles.