Compute aerodynamic lift from air density, true airspeed, wing reference area, and lift coefficient using the fundamental lift equation. The calculator also reports the dynamic pressure, lift per unit wing area, and the rearranged lift coefficient so you can see how each term drives the force on the wing.
Lift is the aerodynamic force that holds an aircraft in the air, generated by the pressure difference between the upper and lower surfaces of a wing as air flows past it. The lift equation distills that complex flow into four measurable quantities — air density, airspeed, wing area, and a single dimensionless lift coefficient — making it the workhorse relation of aircraft performance and conceptual design.
L = ½ · ρ · V² · S · C_L.
Here L is the lift force (N), ρ the air density (kg/m³), V the true airspeed relative to the air (m/s), S the wing reference (planform) area (m²), and C_L the dimensionless lift coefficient. Because lift scales with V², doubling speed quadruples lift at a fixed coefficient — which is why takeoff and landing happen at high angle of attack (large C_L) to make up for the low speed.
The group q = ½ρV² is the dynamic pressure, the kinetic energy per unit volume of the oncoming flow. Lift can be written compactly as L = q · S · C_L, and the same q appears in drag, moment, and pressure-coefficient relations. Dynamic pressure (often shown on aircraft as the "q" limit) is what physically loads the structure, so it bounds how fast an airframe can fly at a given altitude.
C_L is a non-dimensional measure of how effectively a wing converts dynamic pressure into lift at a given flow condition. It bundles airfoil shape, angle of attack, camber, and (through Reynolds and Mach effects) the flow regime into one number. Typical cruise values are 0.2–0.5; high-lift configurations with flaps and slats can reach 2.5–3.0 before the wing stalls.
For a given configuration, C_L rises almost linearly with angle of attack until the flow can no longer follow the upper surface. At the stall angle (commonly around 15° for a clean wing) the flow separates, C_L reaches its maximum C_Lmax and then drops sharply. Because stall is tied to C_Lmax — not directly to speed — the stall speed grows with weight and falls with air density, which is why pilots fly margins above 1g stall speed.
Use consistent SI units: density in kg/m³, airspeed in m/s, and area in m². The result is then in newtons (N). At sea level on a standard day, density is about 1.225 kg/m³; it drops with altitude, which is why aircraft must fly faster (higher true airspeed) to hold the same lift up high.
The physics depends on the actual density and the actual speed of the air over the wing, which is true airspeed (TAS). Indicated airspeed (IAS) is essentially a measure of dynamic pressure q, so it conveniently stays constant for a given C_L at fixed weight regardless of altitude — but the lift equation in this form needs ρ and V (TAS) separately.
Use the reference (planform) area S — the projected area of the wing including the part that notionally passes through the fuselage, as defined for that aircraft. The lift coefficient is always quoted with respect to this same reference area, so as long as you are consistent the product S·C_L gives the correct lift.
In steady level flight lift equals weight, so set L = W and solve C_L = W / (q·S) = 2W / (ρV²S). This is the required coefficient; comparing it to the wing's C_Lmax tells you your stall margin. The calculator's rearranged C_L card does exactly this once you enter the lift you want to support.
It remains valid as the definition of C_L, but the value of C_L itself changes with Mach number once compressibility matters (above roughly Mach 0.3). Near and above the drag-divergence Mach number, shock waves alter the pressure distribution and C_L, so transonic and supersonic design pairs this equation with compressible-flow and Mach analysis.