The NACA four digit wing sections were derived to have the same thickness distribution as the G�ttingen 398 and the Clark Y airfoils when their camber was removed. These airfoils were noted as being particularly efficient by the designers of the NACA family.

The 4-digit thickness distribution is given by the following equation:

y = (t/c) * ( A sqrt(x) + B x + C x^{2} + D x^{3} + E x^{4} )

where ** denotes exponentiation, (t/c) is the maximum thickness to chord ratio of the airfoil, x is the position as fraction of chord, and y is the half-thickness as fraction of chord.

The coefficients are:

A = 1.4845

B = -0.630

C = -1.758

D = 1.4215

E = -0.5075

You may note that these numbers are 5 times the values usually shown in references. The usual numbers are for computing a 20 percent airfoil which you are then expected to scale to your desired thickness.

Of course, for computing efficiency, y is computed from

y=(t/c)*(A sqrt(x) + x*(B + x*(C + x*(D + x*E)))))

Be sure to note that y is the HALF-THICKNESS. The maximum thickness ocurrs at approximately 30 percent of chord.

Also note that the trailing edge thickness is NOT zero.