NACA65A320

NACA 65A-320

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Curve fits of basic Javafoil data for NACA65A320, smooth airfoils, Mach 0:

Re: 500k | 1M | 3M | 6M

Re = 500k | 1M | 3M | 6M

Lift: Cl = f(α) = -1.547e-04*α³ + -3.488e-04*α² + 1.291e-01*α + 3.144e-01
Drag: Cd = f(α) = -9.379e-07*α³ + 1.389e-04*α² + 3.309e-04*α + 9.144e-03
Moment: Cm = f(α) = 6.106e-07*α³ + 2.579e-05*α² + -2.658e-03*α + -6.268e-02

Re = 1M | 500k | 3M | 6M

Lift: Cl = f(α) = -1.570e-04*α³ + -3.470e-04*α² + 1.300e-01*α + 3.152e-01
Drag: Cd = f(α) = -1.426e-06*α³ + 1.221e-04*α² + 3.735e-04*α + 7.621e-03
Moment: Cm = f(α) = 7.771e-07*α³ + 2.377e-05*α² + -2.678e-03*α + -6.290e-02

Re = 3M | 500k | 1M | 6M

Lift: Cl = f(α) = -1.559e-04*α³ + -3.557e-04*α² + 1.307e-01*α + 3.159e-01
Drag: Cd = f(α) = -7.723e-07*α³ + 9.403e-05*α² + 2.850e-04*α + 6.390e-03
Moment: Cm = f(α) = 7.254e-07*α³ + 2.087e-05*α² + -2.692e-03*α + -6.309e-02

Re = 6M | 500k | 1M | 3M

Lift: Cl = f(α) = -1.551e-04*α³ + -3.609e-04*α² + 1.311e-01*α + 3.162e-01
Drag: Cd = f(α) = -5.521e-07*α³ + 7.953e-05*α² + 2.472e-04*α + 5.896e-03
Moment: Cm = f(α) = 7.691e-07*α³ + 1.942e-05*α² + -2.701e-03*α + -6.319e-02