CalculatorHub Search

Lift Coefficient Calculator

Reviewed by the CalculatorHub team, edited by James Graham. Standard lift equation per NASA Glenn Research Center. Last verified 14 June 2026.

The lift coefficient (CL) is a dimensionless measure of a wing's ability to generate lift. It is derived from the fundamental lift equation: L = 0.5 x rho x V^2 x S x CL, where L is lift, rho is air density, V is velocity, and S is wing area. This calculator solves for CL given the other variables, or computes lift given CL. Understanding CL is essential for wing design, stall speed analysis, and flight performance calculations across all phases of flight.

In level flight, lift equals aircraft weight in Newtons (weight in kg x 9.81)
ISA sea level = 1.225 kg/m3; decreases with altitude
True airspeed in metres per second (1 kt = 0.5144 m/s)
Planform wing area including fuselage carry-through
0.71
2,205.00 Pa
3,125.00 N/m2

Lift equation and coefficient formula

q = 0.5 x rho x V^2 (dynamic pressure, Pa)
L = q x S x CL (lift force, N)
CL = L / (q x S) = 2L / (rho x V^2 x S)
Wing Loading = L / S (N/m2)

In steady level flight, lift equals weight, so L = m x g where m is mass in kg and g = 9.81 m/s^2. The dynamic pressure q = 0.5 x rho x V^2 is in Pascals when rho is in kg/m3 and V is in m/s. CL is dimensionless. This formulation is from NASA TP-1120 and standard aerodynamics textbooks.

Typical lift coefficient values

  • Clean cruise (airliner at altitude): CL = 0.3 to 0.6.
  • Light aircraft cruise: CL = 0.4 to 0.8.
  • Takeoff with flaps: CL = 1.4 to 2.5.
  • Landing with full flaps: CL = 1.8 to 3.5.
  • Stall (CL_max for clean wing): typically 1.2 to 1.8.
  • CL above CL_max causes stall; the wing can no longer maintain attached flow.

Lift coefficient calculator: frequently asked questions

What is the lift coefficient?

The lift coefficient (CL) is a dimensionless number that describes how effectively a wing generates lift relative to the dynamic pressure and wing area. It captures the combined effects of wing shape, angle of attack, and airflow conditions. A higher CL means more lift for a given dynamic pressure and wing area.

How is the lift coefficient calculated?

CL = 2L / (rho x V^2 x S), where L is the lift force in Newtons, rho is air density in kg/m3, V is airspeed in m/s, and S is wing reference area in m2. This is the standard lift equation rearranged to solve for CL. It is equivalent to CL = L / (q x S), where q = 0.5 x rho x V^2 is the dynamic pressure.

What is a typical CL for aircraft wings?

Clean cruise CL for commercial airliners is typically 0.3 to 0.6. At takeoff with flaps extended, CL can reach 2.0 to 3.0. Stall occurs when CL reaches the maximum value (CL_max), which is typically 1.2 to 1.8 for a clean wing and 2.5 to 3.5 with high-lift devices. Below CL_max, flow remains attached; above it, the wing stalls.

How does angle of attack affect lift coefficient?

For most wings, CL increases approximately linearly with angle of attack at a rate of about 2*pi per radian (0.11 per degree) in subsonic flow, per thin airfoil theory. This linear relationship breaks down near stall. The exact CL versus alpha curve depends on wing geometry and must be determined by wind tunnel testing or computational analysis for a specific design.

What is the relationship between lift coefficient and drag?

Lift and drag are related through the lift-to-drag ratio (L/D = CL/CD). Induced drag coefficient (CDi) = CL^2 / (pi x e x AR), where e is the Oswald efficiency factor and AR is the wing aspect ratio. Higher CL increases induced drag rapidly (quadratic relationship). Maximum range occurs at the Mach number and CL where L/D is maximized.

Official sources

Reviewed by the CalculatorHub team, edited by James Graham, 14 June 2026. See our methodology.