Lift Coefficient Calculator

The lift coefficient is a dimensionless number that describes how effectively a wing or airfoil turns airflow into lift. It lets you compare wings of different sizes and speeds on equal footing: lift equals one half times air density times speed squared times wing area times CL.

CL = 2L / (ρ × v² × A), where L is lift force, ρ (rho) is air density, v is airspeed, and A is the wing reference area. The equivalent lift equation is L = 1/2 × ρ × v² × A × CL. This calculator solves the equation in all five directions: enter any four values and it finds the fifth.

As a rule of thumb for subsonic airfoils: cruise CL is often 0.2 to 0.5, the typical operating range is about 0.3 to 1.5, and most airfoils stall near CL 1.6 to 1.8. A computed CL above about 1.6 is flagged as a stall warning here — it means the wing would need an angle of attack past where airflow normally separates. These bands are approximations, not a stall prediction for any specific wing.

Select the ISA altitude option and enter altitude: the tool applies a simplified ISA-1976 model — 1.225 kg/m³ at sea level, temperature falling 0.0065 K per meter up to 11 km, then an isothermal layer above that. You may also enter a custom temperature; pressure still comes from the standard column and density is computed as p / (R × T). Humidity and real weather deviations are not modeled, so for actual flight planning use certified performance data.

Dynamic pressure, q = 1/2 × ρ × v², is the kinetic energy per unit volume of the airflow, measured in pascals. The lift equation can be rewritten as L = q × A × CL, so lift is simply dynamic pressure times wing area times the lift coefficient. The calculator shows q for every solution as a quick sanity check on your inputs.

In steady level flight, lift equals weight. Setting L = W in the lift equation with the lift coefficient at its maximum gives v_stall = √(2W / (ρ × A × CLmax)). Enter weight, air density, wing area, and CLmax — the default 1.5 is a common rule-of-thumb value. Real stall speed also depends on load factor, flap configuration, and the actual wing, so treat the result as an educational estimate.

Lift and weight accept newtons, pound-force, or kilogram-force, and weight also accepts kilograms as mass. Density accepts kg/m³, g/m³, slug/ft³, or lb/ft³. Speed accepts m/s, km/h, mph, knots, or ft/s. Area accepts m² or ft², and altitude accepts meters or feet. Every input is converted to SI before the math runs, so mixing units is safe.

Yes. A negative CL means lift points downward — inverted flight, or a downforce wing on a race car. Values above 2 exist on high-lift configurations with large flaps, slats, or blown flaps, but they sit far beyond the 1.6 to 1.8 stall region of a plain subsonic airfoil, so this tool flags them as outside the normal envelope.

The three classic ones: forgetting the 1/2 in the lift equation, which doubles the answer; mixing km/h with m/s without converting, an error of 3.6 squared, about a factor of 13; and using the wrong reference area — the CL of a whole wing uses the planform area. For a rectangular wing that equals chord times span, but for tapered or swept wings use the true planform area.

No. A 2D airfoil-section CL comes from wind-tunnel or CFD data at a given angle of attack, while a finite 3D wing produces a lower CL at the same angle because wingtip vortices create induced downwash — the difference grows as aspect ratio shrinks. Use section data for airfoil analysis and the whole-wing reference area for aircraft performance work.