NACA 0012 Airfoil
The NACA 0012 is the most widely used and studied airfoil in all of aerodynamics: a symmetric section, 12% thick, with no camber. Because it is symmetric it makes zero lift at zero angle of attack and behaves identically upright or inverted.
What the numbers mean
In the code 0012, the first two digits are 0 and 0, meaning zero camber — the airfoil is perfectly symmetric top-to-bottom. The last two digits, 12, give the maximum thickness as 12% of the chord. With no camber there is no ‘position of maximum camber,’ which is why the middle digit is zero.
Aerodynamic characteristics
Because it is symmetric, the NACA 0012 produces no lift at zero angle of attack and has essentially zero pitching moment about the quarter-chord. Its behaviour is smooth and well documented, which is exactly why it is the standard benchmark for validating wind-tunnel and CFD methods.
| Property | Typical value |
|---|---|
| Zero-lift angle α₀ | 0° |
| Lift-curve slope | ≈ 0.11 /deg |
| Max lift coefficient Cl,max | ≈ 1.5 |
| Stall angle | ≈ 15–16° |
| Minimum drag coefficient Cd,min | ≈ 0.006 |
| Pitching moment Cm,c/4 | ≈ 0.00 |
Conditions: Reynolds number Re = 1×10⁶, low speed (Mach ≤ 0.15), smooth surface. Values are representative figures from thin-airfoil theory and published XFoil / wind-tunnel data, rounded for reference.
Where it's used
- Tail surfaces — horizontal and vertical stabilizers, where a symmetric section that works equally in both directions is exactly what's needed.
- Control surfaces — rudders, elevators and ailerons.
- Rotor and turbine blades — helicopter rotors and some wind-turbine sections.
- Aerobatic aircraft — wings that must fly inverted as well as upright.
- CFD and wind-tunnel validation — the 0012 is the go-to reference case for checking that a simulation or experiment gives correct results.
Coordinates
Surface coordinates as fractions of the chord (x/c, y/c), generated from the standard NACA 4-digit equations with a closed trailing edge. For the full high-resolution coordinate file, use the DAT export inside the explorer.
| x/c (upper) | y/c (upper) | x/c (lower) | y/c (lower) |
|---|---|---|---|
| 0.000 | +0.0000 | 0.000 | +0.0000 |
| 0.024 | +0.0259 | 0.024 | -0.0259 |
| 0.095 | +0.0460 | 0.095 | -0.0460 |
| 0.206 | +0.0577 | 0.206 | -0.0577 |
| 0.345 | +0.0596 | 0.345 | -0.0596 |
| 0.500 | +0.0529 | 0.500 | -0.0529 |
| 0.655 | +0.0407 | 0.655 | -0.0407 |
| 0.794 | +0.0264 | 0.794 | -0.0264 |
| 0.905 | +0.0131 | 0.905 | -0.0131 |
| 0.976 | +0.0035 | 0.976 | -0.0035 |
| 1.000 | -0.0000 | 1.000 | +0.0000 |
Explore NACA 0012 in the interactive tool →
See the 3D wing, live Cl / Cd / L-D polar charts, and change the airfoil, angle of attack and Reynolds number in real time.
Open the explorer →Frequently asked questions
NACA 0012 is a symmetric airfoil from the NACA 4-digit family, with a maximum thickness of 12% of the chord and no camber. It is the most common reference airfoil in aerodynamics, used for tails, control surfaces and as a validation benchmark.
Yes, but only when it is at an angle of attack. Because it is symmetric, it makes zero lift at 0 degrees; as the angle of attack increases, lift rises roughly linearly until the airfoil stalls near 15 to 16 degrees.
At a Reynolds number of about 1 million, the maximum lift coefficient is roughly 1.5, reached near a stall angle of 15 to 16 degrees. The exact value depends on Reynolds number and surface condition.
Because it is simple, symmetric and extremely well documented, with decades of consistent wind-tunnel data. That makes it an ideal case for confirming that a new simulation or method produces physically correct results.