Airfoil Geometry
Aerodynamic Coefficients
CL — Lift
0.0000
CD — Drag
0.00000
L/D Ratio
—
Alpha (α)
0.0°
Flow Separation Detected
Attached flow predicted to separate above α ≈ 15°. CL values are post-stall approximations.
Simulation Parameters
−20°+25°
0.5M6.0M
0.5°2.0°
Profile
NACA 0012
Chord
1.00 m
Span
∞ (2D)
Max Thickness
12% chord
Method
Viscous Empirical v2.1
Stall Model
Re-dep. + Hysteresis
CL vs α polar — NACA 0012
Run sweep to generate full polar · current position shown as marker