 NACA 0012

NACA 0012 is a NACA four-digit airfoil with zero specified airfoil camber and nominal maximum airfoil thickness-to-chord ratio $0.12$. Its upper and lower thickness envelopes are reflected across the airfoil chord, making it a symmetric airfoil. The designation describes section geometry; it is not a measured lift or drag specification.^[ref]

To construct it, set the airfoil mean camber line to zero directly and use the standard thickness envelope in trailing edge with $\eta =0.12$. No mean-line location parameter is needed. For the finite-edge convention, the normalized half-thickness at $x/c=0.5$ is $y_t/c\approx 0.052940252$. An original section with $c=0.30\mathrm{m}$ therefore has surface heights $\pm 15.8821\mathrm{mm}$ at that station and full local thickness $31.7642\mathrm{mm}$. The nominal maximum thickness is about $36\mathrm{mm}$; twelve percent is not the thickness at every station.

Its nose-radius formula is $r_{\mathrm{LE}}/c\simeq 1.1019(0.12{)}^2$. The same $0.30\mathrm{m}$ chord gives $r_{\mathrm{LE}}\simeq 4.7602\mathrm{mm}$. The finite-edge polynomial also leaves a small nonzero rear gap; closing that endpoint changes the coordinate specification. These are consequences of the stated construction, not independent manufacturing measurements.

A useful physical test records pressure and section lift while varying angle of attack at controlled Reynolds number. Thin airfoil theory predicts a zero zero-lift angle of attack for the symmetric ideal model, but it does not supply a universal aerodynamic stall angle. The original NACA experimental programme systematically compared related shapes in a variable-density wind tunnel, providing a reason to identify both the section and the flow conditions when citing results.^[ref]