NACA M2

NACA M2 AIRFOIL

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Curve fits of basic Javafoil data for NACA M2, smooth airfoils, Mach 0:

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

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

Lift: Cl = f(α) = -2.476e-04*α³ + -1.010e-05*α² + 1.027e-01*α + 9.541e-04
Drag: Cd = f(α) = 3.108e-08*α³ + 5.935e-04*α² + -3.141e-06*α + 1.201e-02
Moment: Cm = f(α) = 4.481e-06*α³ + -1.259e-07*α² + -1.318e-03*α + 2.707e-05

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

Lift: Cl = f(α) = -3.349e-04*α³ + -5.369e-05*α² + 1.128e-01*α + 2.733e-03
Drag: Cd = f(α) = 1.291e-06*α³ + 5.926e-04*α² + -8.189e-05*α + 5.172e-03
Moment: Cm = f(α) = 8.353e-06*α³ + 4.033e-06*α² + -1.809e-03*α + -9.533e-05

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

Lift: Cl = f(α) = -1.917e-04*α³ + 8.950e-05*α² + 1.165e-01*α + -2.599e-04
Drag: Cd = f(α) = -1.591e-07*α³ + 5.386e-05*α² + 2.357e-06*α + 9.366e-03
Moment: Cm = f(α) = 1.148e-06*α³ + 3.006e-07*α² + -2.011e-03*α + -2.551e-05

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

Lift: Cl = f(α) = -1.973e-04*α³ + 3.783e-05*α² + 1.168e-01*α + 9.606e-04
Drag: Cd = f(α) = -1.383e-09*α³ + 4.142e-05*α² + -5.473e-06*α + 9.153e-03
Moment: Cm = f(α) = 1.118e-06*α³ + -2.133e-13*α² + -2.015e-03*α + 4.835e-24