Add RocketCEA integration for propellant thermochemistry

[cmflannery]

Summary

Integrates NASA CEA (via RocketCEA) for automatic combustion thermochemistry calculations. Engineers can now specify propellants directly instead of manually looking up Tc, gamma, and molecular weight.

Changes

• New propellants.py module

  • get_combustion_properties() - get Tc, MW, gamma, c* from CEA
  • get_optimal_mixture_ratio() - find optimal O/F for max Isp
  • Propellant name normalization (LOX/LO2/O2 all work)

• EngineInputs.from_propellants() factory method

  • Specify oxidizer, fuel, thrust, chamber pressure
  • Thermochemistry auto-calculated from CEA

• RocketCEA is a required dependency

  • Requires Fortran compiler: brew install gcc (macOS) or apt install gfortran (Linux)

• Bug fix: Fixed convergent section discontinuity in nozzle contour generation

Example

from openrocketengine import EngineInputs, design_engine
from openrocketengine.units import kilonewtons, megapascals

# Before: manually specify Tc=3620K, gamma=1.17, MW=23.3
# Now: just specify propellants
inputs = EngineInputs.from_propellants(
    oxidizer="LOX",
    fuel="RP1",
    thrust=kilonewtons(100),
    chamber_pressure=megapascals(7),
    mixture_ratio=2.7,
)

performance, geometry = design_engine(inputs)

Testing

• Added tests/test_propellants.py with CEA integration tests
• Added examples/propellant_design.py demonstrating LOX/RP-1, LOX/CH4, LOX/LH2 comparison