AeroProp-X

Open-source rocket nozzle design, Method-of-Characteristics (MOC) solver, and aerothermodynamic performance analysis — in pure Python.

AeroProp-X turns chamber conditions and a target expansion into a mathematically-grounded converging–diverging (de Laval) nozzle contour. It generates minimum-length MOC bells, Rao thrust-optimised bells, and conical baselines; maps the supersonic flow field; analyses over/under-expansion and flow separation; and exports CAD (CSV / DXF / FreeCAD-STEP) you can open in Fusion 360, SolidWorks, or FreeCAD. No paywalled toolbox required.

Scope and honesty. AeroProp-X models steady, inviscid, adiabatic, shock-free flow of a calorically-perfect gas. Results are ideal preliminary-design estimates — excellent for learning, sizing, and early trade studies, but not a replacement for CFD or hot-fire testing. Every approximation is documented in docs/theory.md.

Architecture

            ┌─────────────────────────────────────────────┐
            │     High-pressure combustion chamber (p0,T0) │
            └───────────────────────┬─────────────────────┘
                                    ▼
                          Converging section
                       (subsonic acceleration)
                                    │
                                    ▼
                          Throat  (M = 1, sonic)
                                    │
                                    ▼
              ┌──────────────── Diverging section ───────────────┐
              │           supersonic expansion, M > 1            │
              │                                                  │
              │   gasdynamics ──► moc ──► nozzle ──► flowfield   │
              │   (isentropic +   (char.   (MOC/Rao/  (Mach/p/T   │
              │    Prandtl-Meyer)  mesh)    conical)   field)     │
              └──────────────────────┬───────────────────────────┘
                                    ▼
                 performance  ──►  cad_export / plotting
            (C_F, c*, Isp, thrust,   (CSV / DXF / FreeCAD-STEP,
             expansion regime)        contour & flow-field figures)
                                    │
                                    ▼
                    Optimal contour: max thrust, shock-free,
                            uniform axial exit

Note on the original project brief. The brief that seeded this repo contained pipeline labels copied from an unrelated speech/NLP template — "Linguistic & Acoustic Decomposition", "Neural Synthesis Module", "Linguistic/Acoustic Expansion Lane". Those concepts have no meaning in nozzle aerodynamics and are not implemented. The architecture above reflects the actual gas-dynamics pipeline.

Installation

git clone https://github.com/your-org/aeroprop-x
cd aeroprop-x
pip install .            # core (numpy, scipy)
pip install .[plot]      # add matplotlib for figures
pip install .[dev]       # add pytest for the test-suite

Requires Python ≥ 3.9.

Quick start (Python)

from aeroprop_x import moc_nozzle, evaluate, to_freecad_macro

# 1. Minimum-length MOC contour for exit Mach 2.6, 40 mm throat radius.
contour = moc_nozzle(exit_mach=2.6, n_characteristics=80, throat_radius=0.04)
print(contour.area_ratio, contour.length, contour.meta["area_ratio_error"])

# 2. Ideal performance at 10 km (pa ~ 26.5 kPa).
perf = evaluate(exit_mach=2.6, p0=5e6, pa=26_500, gamma=1.4,
                R=287, T0=3000, throat_area=3.14159 * 0.04**2)
print(perf.report())          # C_F, c*, Isp, thrust, expansion regime

# 3. Export a FreeCAD macro that revolves the contour and writes STEP.
to_freecad_macro(contour, "nozzle.FCMacro")

Quick start (CLI)

# Design a minimum-length MOC contour and save CSV + DXF + a PNG.
aeroprop-x design moc --exit-mach 2.4 --throat-radius 0.05 \
    --csv nozzle.csv --dxf nozzle.dxf --plot contour.png

# Design an 80% Rao bell from an area ratio.
aeroprop-x design bell --area-ratio 25 --throat-radius 0.05 --csv bell.csv

# Evaluate performance at altitude.
aeroprop-x analyze --area-ratio 25 --p0 7e6 --pa 26500 \
    --gamma 1.22 --R 320 --T0 3500 --throat-area 0.01

# Export an MOC design straight to a FreeCAD STEP macro.
aeroprop-x export moc --exit-mach 3.0 --throat-radius 0.04 --freecad nozzle.FCMacro

What it does

1. Nozzle contour generation. A rigorous planar minimum-length MOC solver (sharp-corner centred expansion, shock-free wall cancellation), plus a Rao thrust-optimised parabolic bell and a conical baseline. Enter exit Mach or area ratio.
2. Supersonic flow-field plotting. Extract Mach, pressure, and temperature over the characteristic mesh and interpolate onto a regular grid for colour-mapped field plots.
3. Expansion & separation analysis. Classify perfectly / under / over- expanded operation against ambient pressure and flag likely flow separation (Summerfield criterion).
4. CAD export. Write CSV, ASCII DXF polylines, or a FreeCAD .FCMacro that revolves the contour to a solid and exports STEP for SolidWorks / Fusion 360 / FreeCAD.

Repository layout

aeroprop-x/
├── src/aeroprop_x/
│   ├── gasdynamics.py     isentropic + Prandtl-Meyer relations
│   ├── moc.py             Method-of-Characteristics min-length solver
│   ├── nozzle.py          MOC / Rao bell / conical contour generators
│   ├── flowfield.py       mesh -> Mach/p/T field extraction & gridding
│   ├── performance.py     C_F, c*, Isp, thrust, expansion analysis
│   ├── cad_export.py      CSV / DXF / FreeCAD-STEP export
│   ├── plotting.py        contour, mesh, and field figures (matplotlib)
│   └── cli.py             `aeroprop-x` command-line interface
├── tests/                 pytest suite (gas dynamics, MOC, nozzle, perf)
├── examples/              quickstart.py, design_ssme_like.py
├── docs/theory.md         equations, derivations, assumptions
├── validation_cases/      comparison vs tables & historical engines
├── pyproject.toml         packaging + console-script entry point
├── CHANGELOG.md  CONTRIBUTING.md  LICENSE (MIT)

Validation

• Isentropic and Prandtl–Meyer outputs match standard compressible-flow tables to table precision.
• The MOC solver self-tests: for planar flow the exit height ratio converges to the exact isentropic A/A* (relative error ≈ 3×10⁻⁴ at n = 160 for M = 2.4), the peak centreline Mach equals the design Mach, and the exit flow leaves parallel to the axis.

See validation_cases/README.md for the full tables and reproduction commands.

Key limitations (read before trusting a number)

• Planar MOC. The MOC solver is 2-D. Revolving its contour as axisymmetric overshoots the design area ratio by √(A/A*); for accurate axisymmetric area ratios use rao_bell. A true axisymmetric MOC is on the roadmap.
• Ideal gas, inviscid. No boundary layers, finite-rate chemistry, heat transfer, or real-gas effects; Isp and thrust are ideal upper estimates.
• Approximate defaults. Rao wall angles are coarse chart interpolations; the Summerfield separation threshold is an engineering estimate.

License

MIT — see LICENSE. Contributions welcome; see CONTRIBUTING.md.

References

Anderson, Modern Compressible Flow (MOC, minimum-length nozzle) · Sutton & Biblarz, Rocket Propulsion Elements · Rao, "Exhaust Nozzle Contour for Optimum Thrust" (1958).