NACA Airfoil Calculator
Enter any NACA 4-digit airfoil, an angle of attack and a Reynolds number to get the lift coefficient, drag coefficient, lift-to-drag ratio and pitching moment, plus a live lift curve and the real lift force for your wing. Same engine as the interactive explorer.
An airfoil calculator estimates how much lift and drag a wing section makes. Lift coefficient is Cl = a₀·(α − α₀) and lift force is L = Cl·½ρV²·S. Type a NACA code, angle and Reynolds number below to get Cl, Cd, L/D, pitching moment and lift force instantly.
1 Airfoil & conditions
2 Results
Lift curve — Cl vs angle of attack
This tool estimates the aerodynamic coefficients of a NACA 4-digit airfoil and converts them into real forces. Lift and pitching moment come from thin-airfoil theory; drag, maximum lift and the stall angle use an empirical model calibrated to XFoil polars at a Reynolds number of 1×10⁶ and Ncrit = 9. It is the same calculation the interactive 3D explorer uses, so the numbers match.
Lift coefficient (Cl)
Lift coefficient is a dimensionless measure of how much lift a section produces at a given angle:
- a₀Lift-curve slope, about 2π(1 + 0.77·t/c)·0.95 per radian — how fast lift builds with angle.
- αAngle of attack, in degrees.
- α₀Zero-lift angle, found by integrating the camber line. Zero for symmetric sections, negative for cambered ones.
Near the stall angle the lift is rolled off and then capped at an XFoil-calibrated Cl,max, so the model does not over-predict peak lift the way pure thin-airfoil theory would.
Drag coefficient (Cd) and L/D
Drag is a minimum-drag floor plus a quadratic "drag bucket" that grows as lift moves away from the design point, scaled for Reynolds number:
Cd,min and k calibrated to XFoil Re = 1×10⁶, Ncrit = 9
The lift-to-drag ratio L/D = Cl / Cd is the efficiency number: higher means more lift for less drag.
Pitching moment (Cm,c/4)
The quarter-chord pitching moment comes from the camber line's Fourier coefficients. Symmetric airfoils (like 0012) have Cm ≈ 0; cambered airfoils have a nose-down (negative) moment.
From coefficient to real lift force
Once you have Cl, the actual lift force on a wing is:
ρ = air density · V = airspeed · S = wing area
The term ½ρV² is the dynamic pressure. The calculator reads air density from the International Standard Atmosphere at your chosen altitude, so you only enter speed, altitude and wing area. Drag force uses the same formula with Cd.
If you are new to airfoil data, here is what each output actually tells you, in plain terms:
- ClLift coefficient — how hard the section lifts at this angle. Bigger means more lift for the same speed and size.
- CdDrag coefficient — the aerodynamic "cost." Lower is better, and it climbs fast near stall.
- L/DLift-to-drag ratio — efficiency. A glider might reach 40–60; a thick, cambered wing far less. Peak L/D is the sweet spot for cruise.
- CmPitching moment — how much the section tries to twist nose-down. Matters for trim and structural loads.
- α₀Zero-lift angle — the angle where lift is exactly zero. Cambered wings still lift at 0°, so α₀ is negative.
- α stallStall angle — push past it and the flow separates, lift drops and drag spikes.
- Decode the airfoil: NACA 2412 is 2% camber at 40% chord, 12% thick. Its zero-lift angle works out to
α₀ ≈ −2.1°. - Lift coefficient:
Cl = a₀·(α − α₀)with the thick-section slope givesCl ≈ 0.81at 5°. - Drag from the XFoil-calibrated bucket:
Cd ≈ 0.0086, so efficiency isL/D ≈ 94. - Now a wing: fly at V = 50 m/s at sea level (ρ = 1.225 kg/m³) with wing area S = 16 m². Dynamic pressure
q = ½ρV² ≈ 1531 Pa. - Lift force:
L = Cl · q · S = 0.805 × 1531 × 16 ≈ 19.7 kN— roughly the weight of a 2-tonne aircraft.
Reynolds number Re = ρVc/μ sets how "sticky" the air feels to the wing. It mainly changes drag and the stall angle. Low Re means thicker boundary layers, more drag and an earlier, gentler stall. Pick a value near your real operating point:
| Application | Typical Reynolds number |
|---|---|
| Indoor / micro RC, small drones | 50,000 – 200,000 |
| RC aircraft, large UAVs | 200,000 – 500,000 |
| Light aircraft wing section | ~1,000,000 |
| General aviation / gliders | 1 – 3 million |
| Airliners, large aircraft | 10 – 50 million |
Aircraft & glider design
Pick a wing section and size it: estimate cruise Cl, check L/D at the design point, and size the wing area for a target weight.
RC aircraft & drones
Choose an airfoil for low-Reynolds flight, compare stall behaviour, and sanity-check lift at your model's speed and wing area.
Wind turbines & propellers
Blade sections live in the same NACA family; L/D and stall angle drive how much torque a blade makes before it separates.
Teaching & CFD validation
A fast, transparent first estimate for coursework, and a quick reality-check for the trends you expect before running XFoil or CFD.
Cl is not lift. Lift coefficient is dimensionless. The actual force also depends on speed², air density and wing area. Doubling speed roughly quadruples lift at the same Cl.
Stall is not a single number. The stall angle shifts with Reynolds number and surface roughness, and real stall is gradual. Treat the stall angle here as an estimate, not a hard cliff.
Thin-airfoil theory has limits. It is accurate for thin, lightly cambered sections at small angles. For very thick or highly cambered airfoils, and above stall, expect larger errors.
Highest Cl is rarely the goal. More camber raises lift but also drag. For efficient cruise you want peak L/D, not peak Cl.
- Type a NACA 4-digit code (like 2412 or 0012) into the first box.
- Set the angle of attack you care about, between −10° and 22°.
- Enter a Reynolds number near your real operating point — use the table above if unsure.
- Read off Cl, Cd, L/D and Cm, and watch where your point sits on the lift curve.
- For real forces, add airspeed, altitude and wing area to get lift and drag in newtons.
3D Airfoil Explorer
The full interactive simulator. Orbit a 3D wing, drag the angle of attack and watch the polars update live.
Open the simulator →Airfoil database
Reference pages for individual NACA profiles with geometry, coordinates, aerodynamic data and charts.
Browse airfoils →Compare airfoils
NACA 0012 vs 2412 vs 4412 side by side. See exactly what changing camber does to lift, drag and stall.
Compare profiles →NACA 4-digit explained
What each digit in a code like 2412 actually means, and how it changes the way a wing flies.
Read the guide →How do you calculate the lift coefficient of a NACA airfoil?
Cl = a₀ × (α − α₀), where a₀ is the lift-curve slope (about 2π per radian with a small thickness correction), α is the angle of attack and α₀ is the zero-lift angle set by the airfoil's camber. This calculator finds α₀ by integrating the NACA camber line, applies the slope, then rolls lift off near stall and caps it at an XFoil-calibrated Cl,max.How is lift force calculated from the lift coefficient?
L = Cl × ½ρV² × S: lift coefficient times dynamic pressure times wing area. ρ is air density, V is airspeed and S is wing area. The calculator uses the International Standard Atmosphere to get density from altitude, so you only enter speed, altitude and wing area.What Reynolds number should I use?
Re = ρVc/μ depends on density, speed and chord. Use roughly 100,000–500,000 for RC models and small drones, about 1×10⁶ for light-aircraft wing sections, and several million for full-size aircraft. Reynolds number mainly affects drag and the stall angle, not the lift-curve slope.