✈️Flight Physics Simulator - Airfoil Lift, Stall Speed, Glide Ratio

How to use Flight Physics Simulator (Lift, Stall & Aircraft Performance)

1. 1On the Airfoil tab, enter a NACA code (e.g. 2412) and drag the angle of attack to see the streamlines, pressure distribution and C_L–α curve
2. 2Adjust aspect ratio, Oswald efficiency and zero-lift drag to see how the wing lift slope and drag polar change
3. 3On the Aircraft tab, choose an aircraft, set altitude and true airspeed, and read stall speed, whether there is enough lift, power required and max L/D
4. 4On the Turn tab, drag the bank angle for load factor, turn radius and stall speed in the turn; on the Climb tab, move the throttle for climb angle and rate

FAQ

Why does a wing stall, and how is that related to speed?

A stall is about angle of attack: past a critical angle the flow can no longer follow the upper surface and separates, so the lift coefficient peaks at C_Lmax and then drops. In level flight lift must equal weight, L = ½ρV²SC_L, so the slower you fly the larger the C_L you need and the closer you get to C_Lmax. The stall speed V_s = √(2W/(ρSC_Lmax)) is just the speed where 1 g flight reaches that angle. It rises with weight and with thin air (high, hot airfields), falls when flaps raise C_Lmax, and grows by √n in a turn with load factor n.

How accurate is thin-airfoil theory, C_L = 2π(α − α₀)?

For thin, mildly cambered airfoils below the stall it agrees well with wind-tunnel data; measured lift slopes are typically 90–100% of 2π. The zero-lift angle α₀ comes from the camber line, so a cambered airfoil such as NACA 2412 already lifts at 0° (α₀ ≈ −2°). The theory cannot predict stall, so the tool switches to a simplified post-stall model above C_Lmax. On a finite wing, tip vortices reduce the effective angle and the slope becomes a₀/(1 + a₀/(πeAR)).

Is the streamline animation the real airflow?

It is idealised potential flow: inviscid, incompressible air. The Joukowski transform maps the flow around a cylinder onto an airfoil, and the Kutta condition makes the flow leave the sharp trailing edge smoothly, which fixes the circulation and hence the lift. It correctly shows faster flow and lower pressure over the top and the suction peak near the leading edge, but it has no boundary layer and so never stalls. The Joukowski airfoil is matched to the NACA thickness and camber, so it is close to, but not exactly, the NACA shape.

How are best-glide speed and glide distance worked out?

In a glide the descent angle γ satisfies tanγ = D/L, so a higher lift-to-drag ratio glides farther. With a parabolic drag polar, (L/D)max = ½√(πeAR/C_D0), reached where induced drag equals parasite drag, at V_md = √(2W/(ρSC_L*)). In still air you can cover about h·(L/D)max from height h: roughly 9:1 for a Cessna 172 and over 40:1 for a high-performance glider. Weight does not change the best glide ratio, only the speed that gives it; wind and rising or sinking air change the distance over the ground.

Why is a 60° bank turn exactly 2 g?

In a coordinated turn the lift tilts by the bank angle φ and its vertical part must hold up the weight: L cosφ = W, so the load factor n = L/W = 1/cosφ, which is 2 at 60°. The horizontal part L sinφ supplies the centripetal force, giving a turn radius R = V²/(g tanφ) and turn rate ω = g tanφ/V. Slower and steeper means a tighter turn, but the extra lift raises the stall speed to V_s√n, which is why steep turns at low speed are dangerous.