⟿ NACA 4-DIGIT // THIN AIRFOIL THEORY

Airfoil Lab

AIRFOIL DESIGN & AERODYNAMIC ANALYSIS

NACA

2412

4 DIGIT

5 DIGIT

MAX CAMBER (m) 2%c

CAMBER POSITION (p) 4×10%c

THICKNESS (t) 12%c

ANGLE OF ATTACK (α) 5°

INSTANTANEOUS AERODYNAMICS

LIFT COEFFICIENT Cl

—

α = 5°

ZERO-LIFT AOA

—

α_L=0

L/D RATIO

—

estimated

MOMENT Cm c/4

—

quarter-chord

CENTER OF PRESSURE

—

x_cp

AERODYNAMIC CENTER

—

x_ac

PERFORMANCE CURVES

ⓘ Cl–α, zero-lift angle, and Cm_c/4 are calculated from thin-airfoil theory (potential flow, small angle approximation). Cl = 2π(α − α_L=0), valid in the linear region. Stall (~12–16°) is empirically smoothed. Drag Cd is APPROXIMATELY modeled with a thickness-dependent parabolic polar (Cd = Cd₀ + k·Cl²) — actual values depend on Reynolds number, surface roughness, and viscous effects.