NACA profiles · Angle of attack · Cl/Cd · Pressure distribution · Stall
Lift is not magic — it emerges from pressure differences above and below a wing. Explore how aerofoil shape and angle of attack determine lift coefficient, drag, and the stall boundary.
A wing generates lift because flow accelerates over the curved upper surface, reducing pressure (Bernoulli's principle), while higher pressure underneath pushes up. The lift coefficient Cl ≈ 2π(α + camber) grows linearly with angle of attack α until flow separates at the stall angle (~16° for NACA 0012). NACA 4-digit profiles encode the geometry: first digit = max camber % chord, second = position of max camber in tenths of chord, last two = max thickness % chord.
Increase the Angle of Attack slider to watch Cl grow and the suction peak on the Cp plot intensify. Push past 15° to trigger stall — the streamlines separate from the upper surface and Cl drops sharply. Switch to NACA 2412 (a cambered profile similar to light aircraft) to see that it produces lift even at 0° angle of attack. The L/D ratio tells you efficiency: gliders achieve L/D ≈ 60, airliners cruise at L/D ≈ 18.
NACA (the predecessor to NASA) developed their aerofoil series in the 1930s by systematically testing hundreds of wing shapes in wind tunnels. The NACA 2412 profile powers the Cessna 172 — the most-produced aircraft in history with over 44 000 built. Modern supercritical aerofoils used in airliners are designed to delay the onset of shock waves at transonic speeds, improving fuel efficiency by 30% over straight NACA profiles.