✈️ NACA Airfoil — Lift, Drag & Pressure Distribution

Generate any NACA 4-digit wing profile and compute aerodynamic coefficients from thin airfoil theory. Adjust camber, camber position, thickness and angle of attack. Watch streamlines shift as lift changes.

Airfoil cross-section with flow (blue = low pressure, red = high pressure)

Pressure coefficient Cp(x/c) — upper surface (blue) / lower (red)

NACA 4-Digit Profile

Max Camber M 2% Camber Position P 0.4c Thickness T 12%

NACA 2412

Flight Conditions

Angle of Attack α 5° Aspect Ratio AR 6

⚠ STALL

Aerodynamics

Cl (lift coeff.)—

Cd (total)—

L/D ratio—

Zero-lift α_L0—

Stall angle—

Cm (moment)—

Thin airfoil theory:
C_l = 2π(α − α_L0)
C_d = C_d0 + C_l²/(πARe)
e ≈ 0.85 (Oswald efficiency)

How the Simulation Works

The NACA 4-digit designation encodes the wing geometry: the first digit is max camber as a percentage of chord, the second is its chordwise position in tenths, and the last two are the thickness ratio in percent. Coordinates are generated using the standard thickness distribution formula, and the camber line is computed analytically. Lift is calculated from thin airfoil theory: C_l = 2π(α − α_L0), where α_L0 is the zero-lift angle determined numerically from the camber distribution. Drag& uses the sum of profile drag (empirical, dependent on T) and induced drag C_di = C_l²/(πARe). The flow visualisation uses a point vortex at the quarter-chord point (Kutta–Joukowski): Γ = C_lV_∞c/2, producing circulation that deflects the streamlines.