Lift Force Calculator

Lift on a wing or airfoil: L = C_L × ½ρV² × S. Find lift, lift coefficient, or stall speed.

Fluid Mechanics II

Altitude / fluid presets — click to auto-fill density

Inputs

Solve for:

Lift coefficient CL (dimensionless)

Compute C_L from angle of attack α (thin airfoil: C_L = 2π sin α)

Free-stream velocity V (m/s)

Air / fluid density ρ (kg/m³)

Enter span × chord instead of area

Wing planform area S (m²)

Include drag coefficient C_D (for L/D ratio)

Drag coefficient CD (dimensionless)

Theory

Lift and drag equations:

L = C_L × ½ρV² × S [N]

D = C_D × ½ρV² × S [N]

L/D = C_L / C_D

Stall speed (level flight, L = Weight W):

V_s = √(2W / (ρ × S × C_L,max)) [m/s]

Thin airfoil theory (CL from angle of attack):

C_L = 2π sin(α) ≈ 2πα (α small, in radians)

Valid for thin, uncambered airfoils at small angles. Real airfoils deviate, especially near stall.

Typical C_L and C_D values:

Airfoil / condition	C_L	C_D
Thin airfoil α = 2°	0.220	0.005
Thin airfoil α = 4°	0.439	0.008
Thin airfoil α = 6°	0.657	0.013
NACA 2412, cruise (α = 4°)	0.650	0.009
NACA 2412, climb (α = 8°)	1.050	0.018
Clark Y, cruise	0.700	0.014
Flat plate α = 5°	0.550	0.080
Near stall (CL_max ≈ 1.5)	1.500	0.070

ISA standard atmosphere densities:

Altitude	ρ (kg/m³)
Air — sea level	1.225
Air — 1 000 m	1.112
Air — 3 000 m	0.909
Air — 5 000 m	0.736
Air — 10 000 m	0.414

Where:

L = lift force [N]
C_L = lift coefficient (dimensionless, function of α and airfoil shape)
ρ = fluid density [kg/m³]
V = free-stream velocity [m/s]
S = wing planform reference area = span × mean chord [m²]
α = angle of attack [°]