Airfoil & Wing Aerodynamics, Design & Simulation with Python

Identify key airfoil parameters such as camber, thickness, and leading edge radius. Explain how airfoil shapes vary with speed across subsonic, transonic, and supersonic regimes and visualize airfoils with Python.

Explore the four forces of flight—lift, drag, thrust, and weight—and how wings, airfoils, and engines balance them to optimize lift and the lift-to-drag ratio.

Define airfoils as the two-dimensional cross sections of wings that generate lift via pressure differences, with shape choices influencing drag across subsonic, transonic, and supersonic speeds.

Explore the two main airfoil types: symmetric and cambered (including reflex cambered), and how their shapes create lift at zero angle of attack, with notes on typical uses.

Learn the key aerofoil parameters, including the chord line, mean camber line, maximum thickness, and camber ratio, and use unitless thickness-to-chord ratios for scalable airfoil design in Python.

Explore how NACA four-digit airfoils encode maximum camber, camber position, and thickness, and how five- and six-digit series relate to lift coefficient and camber details.

Learn how to construct airfoil geometry by using camber line, chord line, and thickness to compute x and y positions for upper and lower surfaces via theta and y_t.

Learn to compute the slope of the camber line (theta) from dyc/dx using arctan, applying NACA four-digit and five-digit camber-line formulas with a 25% chord example.

Learn to calculate the camber line and slope for an airfoil using Python, with x between 0 and 1 and fixed m and p, including arctangent to degrees.

Explore thickness distribution for airfoils using the y_t equation, theta, and camber line. Compute maximum thickness t as a percent of the chord with a 0.5 chord example.

Learn how to compute airfoil half-thickness with Python by implementing a y_t function, using thickness t and position x, and coefficients a0–a4 to reproduce the equation.

Use non-dimensional coordinates to make airfoil design scalable, by dividing x and y by chord length c and working with bar x and y bar in equations and drawings.

Determine the leading edge radius by placing the circle center on the camber-line, applying t/c, and ensuring the circle passes through the origin to shape the airfoil surfaces.

Draw the symmetric NACA 0018 airfoil in Python with numpy and matplotlib, computing y_t and plotting upper and lower surfaces. Note: camber line coincides with the chord for symmetric airfoils.

Learn to draw a cambered naca 2412 airfoil in python by setting thickness 12%, maximum camber 2%, and camber position 0.4, then compute camber, thickness, and upper/lower surfaces.

Generate NACA airfoils with an interactive Streamlit app. Adjust thickness to chord ratio, maximum camber, camber position, chord length, and point density to view symmetric and cambered airfoils updating.

This video demonstrates generating airfoil data using a Nagavali generator, selecting parameters, computing x, y coordinates, camber, and theta, and downloading the result as a csv with pandas.

Analyze aerodynamic forces on airfoils and compute lift and drag coefficients, lift-curve slope, maximum lift, aerodynamic center, center of pressure, flow separation, Reynolds and Mach numbers, and Python.

Explore how lift and drag arise from pressure and shear on airfoils. Understand pitching, rolling, and yawing moments and how angle of attack, Reynolds, and Mach numbers shape wing behavior.

Explore how boundary layer shear stress and pressure distribution generate lift and drag on an airfoil, including skin friction drag, pressure drag, and the resulting total profile drag.

Explore how finite wing span creates wingtip vortices and downwash, contrasting with infinite span assumptions that simplify two-dimensional flow and enable wind tunnel simulations.

Compute dynamic pressure using the q = 1/2 rho v^2 formula, relate rho and velocity to freestream conditions, and apply results to lift and drag on airfoils.

The lecture demonstrates calculating dynamic pressure in Python by defining a function that uses density and velocity, applying the equation q = 0.5 rho v^2, and printing results in pascals.

Explore aerodynamic forces on an airfoil, including lift per unit span and drag per unit span, and decompose normal and chord forces with the angle of attack using trigonometry.

Learn to compute lift per unit span and drag per unit span in python by converting degrees to radians and applying cosine and sine to the angle of attack.

Understand non-dimensional coefficients such as lift coefficient CL and drag coefficient CD, and learn to compute CL, CD, and the lift-to-drag ratio LD from dynamic pressure, chord, and span.

Compute dynamic pressure from 0.5*rho*v^2 and derive lift and drag coefficients using Cl = L/(q_infinity*c) for a chord of two, yielding q ≈ 245, Cl ≈ 1.8, Cd ≈ 0.06.

Learn how to use XFOIL to compute lift and drag coefficients for airfoils, set up the tool parameters, run viscous simulations, and prepare coordinate data for four-digit airfoils.

Design airfoils using simplified camber and thickness equations, export coordinates to a dat file, and compute coefficients with XFOIL, validated against the analytical design using Python and NumPy.

Master running XFOIL with Python by automating input and polar file creation, including loading airfoils, setting viscosity, Reynolds number, and Mach, and saving results for analysis.

Learn to extract lift, drag, and moment coefficients from a polar text file with Python by reading lines and splitting to floats.

Combine the NACA airfoil generator with streamlit to interactively adjust thickness, camber, Reynolds number, Mach, and angle of attack, then calculate lift coefficient, drag coefficient, moment coefficient, and lift-drag ratio.