NACA Wing Creator

A Python library for generating NACA airfoil coordinates with support for 4-digit and 5-digit series. This tool calculates the precise geometric profiles used in aerospace engineering for wing cross-sections.

Features

• ✈️ NACA 4-Series: Analytical calculations for symmetric and cambered airfoils
• ✈️ NACA 5-Series: Support for more complex airfoil shapes with reflex camber
• 📊 Flexible Point Distribution: Configurable number of points with optional cosine spacing for enhanced leading edge resolution
• 🎯 Ready-to-Plot Output: Returns complete airfoil outlines as NumPy arrays
• 🧪 Well-Tested: Comprehensive unit test coverage
• 📈 Visualization Examples: Includes plotting examples for common aircraft airfoils

Installation

Install from GitHub:

pip install git+https://github.com/JeffHanna/NACA_Wing_Creator.git

Or clone and install locally:

git clone https://github.com/JeffHanna/NACA_Wing_Creator.git
cd NACA_Wing_Creator
pip install .

For development (editable install):

pip install -e .

Quick Start

NACA 4-Series Airfoil

from naca import NACA_4

# Create a NACA 2412 airfoil (2% camber, 40% location, 12% thickness)
airfoil = NACA_4("2412", num_points=200, half_cosine_spacing=True)
x, y = airfoil.points

# Plot
import matplotlib.pyplot as plt
plt.plot(x, y)
plt.axis('equal')
plt.grid(True)
plt.xlabel('x/c')
plt.ylabel('y/c')
plt.title('NACA 2412 Airfoil')
plt.show()

NACA 5-Series Airfoil

from naca import NACA_5

# Create a NACA 23016 airfoil (Cl=0.3, 15% camber location, 16% thickness)
airfoil = NACA_5("23016", num_points=200)
x, y = airfoil.points

Usage

Parameters

All NACA classes accept the following parameters:

Parameter	Type	Default	Description
naca_number	str	Required	NACA designation (e.g., "2412", "23016")
num_points	int	200	Number of points per surface (upper/lower)
half_cosine_spacing	bool	True	Use cosine spacing for better leading edge resolution

Understanding NACA Numbering

4-Digit Series (e.g., NACA 2412)

• First digit (2): Maximum camber as % of chord (2%)
• Second digit (4): Location of max camber in tenths of chord (40% from leading edge)
• Last two digits (12): Maximum thickness as % of chord (12%)

5-Digit Series (e.g., NACA 23016)

• First digit (2): Design lift coefficient × 2/3 (Cl = 0.3)
• Second & third digits (30): Location of max camber in half-percent of chord (15%)
• Fourth digit (1): Reflex camber indicator (0=simple, 1=reflex)
• Last two digits (16): Maximum thickness as % of chord (16%)

Examples

The repository includes example_plot.py which demonstrates plotting airfoils used in famous aircraft:

• Boeing B-17: NACA 0018 (root), NACA 0010 (tip)
• Lockheed P-38: NACA 23016 (root), NACA 4412 (tip)
• Cessna 172: NACA 2412 (root and tip)

Run the examples:

python example_plot.py

Comparing Multiple Airfoils

from naca import NACA_4, NACA_5
import matplotlib.pyplot as plt

# Create different airfoils
symmetric = NACA_4("0012")
cambered_4 = NACA_4("2412")
cambered_5 = NACA_5("23012")

# Plot comparison
plt.figure(figsize=(12, 6))
plt.plot(*symmetric.points, label="NACA 0012 (Symmetric)")
plt.plot(*cambered_4.points, label="NACA 2412 (4-series)")
plt.plot(*cambered_5.points, label="NACA 23012 (5-series)")
plt.axis('equal')
plt.grid(True)
plt.legend()
plt.xlabel('x/c')
plt.ylabel('y/c')
plt.title('NACA Airfoil Comparison')
plt.show()

Export to CSV

import numpy as np
from naca import NACA_4

airfoil = NACA_4("2412", num_points=100)
x, y = airfoil.points

# Save coordinates to file
np.savetxt('naca2412.csv', np.column_stack([x, y]), 
           delimiter=',', header='x,y', comments='')

API Reference

NACA_4

Generates NACA 4-digit series airfoil coordinates using analytical equations.

Properties:

• points: Returns (x_array, y_array) tuple with complete airfoil outline
• naca_number: The NACA designation being used
• x_points: X-axis points along the mean camber line

NACA_5

Generates NACA 5-digit series airfoil coordinates using analytical equations with tabulated mean line data.

Properties:

• points: Returns (x_array, y_array) tuple with complete airfoil outline
• naca_number: The NACA designation being used
• x_points: X-axis points along the mean camber line

Supported 5-digit configurations: 210, 220, 221, 230, 231, 240, 241, 250, 251

Testing

Run the test suite:

python -m unittest test_naca.py

The test suite includes:

• Point count validation
• Data structure verification
• Symmetric airfoil validation
• Coordinate range testing
• Leading/trailing edge position checks

Technical Details

Coordinate System

All coordinates are normalized by chord length (c), so x and y values range from 0 to 1 (or ±1 for y depending on airfoil shape).

Point Distribution

The outline path goes: trailing edge (upper) → leading edge → trailing edge (lower)

With half_cosine_spacing=True, points are concentrated near the leading edge where curvature is greatest, providing better geometric fidelity.

Calculations

The library uses the standard NACA equations as documented:

• Wikipedia: NACA Airfoil
• Airfoil Tools: 4-Digit
• Airfoil Tools: 5-Digit

License

This project is open source. See the repository for license details.

Contributing

Contributions are welcome! Please feel free to submit pull requests or open issues for bugs and feature requests.

References

• NACA (National Advisory Committee for Aeronautics) - the predecessor to NASA
• Standard airfoil equations and coefficients from NACA technical reports
• Validated against published airfoil coordinates

Author

Jeff Hanna (@JeffHanna)

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For detailed usage instructions, see USAGE.md