Calculate lift coefficient (C_L), lift force, and aerodynamic parameters from angle of attack, airspeed, wing area, and air density. Includes airfoil data table.
The lift coefficient (C_L) is the dimensionless number that bridges aerodynamic theory and real-world flight. It quantifies how effectively a wing or airfoil converts dynamic pressure into lift, and it depends primarily on the angle of attack, airfoil shape, and Reynolds number. Understanding C_L is essential for aircraft design, wind-turbine blades, race car downforce, and any application where air flows over a shaped surface.
This Lift Coefficient Calculator uses thin-airfoil theory (C_L ≈ 2π sin α) for the two-dimensional case and applies a finite-wing correction based on aspect ratio. It computes lift force, dynamic pressure, and wing loading, and it visualizes how C_L varies with angle of attack. The airfoil data table provides real-world C_L max and stall angles for popular airfoil profiles.
Whether you are a student studying aerodynamics, a drone builder selecting a wing profile, or a pilot estimating stall speed, this calculator makes the core lift equation accessible and visual.
Computing lift from first principles involves multiple interrelated variables — density, velocity, wing area, and the lift coefficient itself. This calculator organizes the inputs logically, applies both 2D and finite-wing corrections, and visualizes the C_L curve so you can see stall-margin instantly. The airfoil reference table puts common C_L max values at your fingertips.
Thin-Airfoil Theory (2D): C_L = 2π sin(α) Lift Force: L = ½ρv²AC_L Finite-Wing Correction (elliptic loading): C_L_finite = C_L_2D / (1 + 2/AR) Dynamic Pressure: q = ½ρv² Where: α = angle of attack (rad) ρ = air density (kg/m³) v = velocity (m/s) A = wing area (m²) AR = aspect ratio (span/chord)
Result: C_L (2D) ≈ 0.657, Lift ≈ 36,300 N
At 6° angle of attack with a 1.5 m chord and 11 m span flying at 60 m/s in sea-level air, thin-airfoil theory gives C_L ≈ 0.657. The resulting lift force is about 36.3 kN — well within the operating envelope for a light aircraft.
A wing produces lift by creating a pressure difference between its upper and lower surfaces. The curved upper surface accelerates the flow, lowering pressure (Bernoulli's principle), while circulation around the airfoil (Kutta condition) establishes the net upward force. The lift coefficient encapsulates this complex interaction into a single number that scales with dynamic pressure and wing area.
High aspect-ratio wings (gliders, albatross wings) are aerodynamically efficient because they minimize tip-vortex effects. The induced drag coefficient is inversely proportional to AR: C_Di = C_L² / (π·e·AR), where e is the Oswald efficiency factor. This is why long, narrow wings are preferred for efficient cruise, while short, wide wings are used for maneuverability.
For real engineering design, thin-airfoil theory is a starting point. Panel methods (XFOIL), vortex-lattice methods, and computational fluid dynamics (CFD) provide progressively more accurate predictions. High-lift devices — flaps, slats, and vortex generators — can increase C_L max by 50–100% beyond the clean-wing value, enabling slower takeoff and landing speeds.
C_L is a dimensionless number that scales dynamic pressure and wing area to produce lift. It depends on the airfoil shape, angle of attack, and Reynolds number. Higher C_L means more lift per unit dynamic pressure.
A finite wing has tip vortices that induce downwash, reducing the effective angle of attack. The finite-wing correction reduces C_L compared to the 2D case, and the effect is greater for low aspect-ratio wings.
Beyond the stall angle (typically 12–18°), flow separates from the upper surface, C_L drops sharply, and drag spikes. The wing loses most of its lifting capability.
Thin-airfoil theory is accurate for small angles (<10°) and thin, symmetric airfoils. For cambered or thick airfoils at higher angles, panel methods or CFD give more accurate results.
Dynamic pressure q = ½ρv² represents the kinetic energy per unit volume of the airflow. It is the scaling factor between C_L and actual lift force.
Yes — air density decreases with altitude, reducing dynamic pressure for the same airspeed. To maintain the same lift at higher altitude, you need higher airspeed or higher C_L (higher angle of attack).