The naming convention of NACA 6-series of airfoils is a combination of the location of maximum thickness, the type of mean line used to generate the chamber line, and the thickness. Of the maximum camber at a coefficient of lift (Cl) value of 0.3. A NACA 23015 has a design lift coefficient of 0.2, the location of the maximum chamber is at 15 percent behind the leading edge, and the airfoil is 15 percent thick. The values for the constants r, k 1 and k 2/k 1 are tabulated for various positions There are also different equations for standard and reflex camber lines. The equation for the camber line is split into two sections like the 4 digit series but the division between the two sections is not at the point of maximum camber. The maximum thickness as percentage.In the examble XX=12 so the maximum thickness is 0.12 or 12% chord. In the examble P=3 so maximum camber is at 0.15 or 15% chordÄ = normal camber line, 1 = reflex camber line The position of maximum camber divided by 20.
Classic NACA Theory Of Wing Sections drag polar display format. The chord can be varied and the trailing edge either made sharp or blunt. Includes all NACA standard airfoil geometry coordinate generators.
#Naca airfoil generator generator
It indicates the designed coefficient of lift (Cl) multiplied by 3/20. NACA 4 Series Airfoil Generator The calculator below can be used to plot and extract airfoil coordinates for any NACA 4-series airfoil. NACA 5 digit airfoils in the database NACA 22112 NACA 23012 NACA 23015 NACA 23018 NACA 23021 NACA 23024 NACA 23112 NACA 24112 NACA 25112 Design coefficient of lift