NACA63-715 a=0.2
NACA63715a=0.2
Curve fits of basic Javafoil data for NACA63-715 a=0.2, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.276e-04*α
3 + 5.854e-04*α
2 + 1.224e-01*α + 4.445e-01
Drag: Cd = f(α) = -5.170e-06*α
3 + 3.787e-04*α
2 + -1.502e-03*α + 7.201e-03
Moment: Cm = f(α) = 1.261e-06*α
3 + 1.045e-04*α
2 + -2.122e-03*α + -5.725e-02
Lift: Cl = f(α) = -2.430e-04*α
3 + 6.222e-04*α
2 + 1.250e-01*α + 4.426e-01
Drag: Cd = f(α) = -1.230e-05*α
3 + 3.764e-04*α
2 + -4.781e-04*α + 2.408e-03
Moment: Cm = f(α) = -3.834e-06*α
3 + 1.122e-04*α
2 + -1.299e-03*α + -6.047e-02
Lift: Cl = f(α) = -2.492e-04*α
3 + 6.412e-04*α
2 + 1.268e-01*α + 4.412e-01
Drag: Cd = f(α) = -1.576e-05*α
3 + 3.339e-04*α
2 + 1.774e-04*α + 5.347e-04
Moment: Cm = f(α) = -4.538e-06*α
3 + 9.445e-05*α
2 + -1.015e-03*α + -6.117e-02
Lift: Cl = f(α) = -2.554e-04*α
3 + 6.515e-04*α
2 + 1.283e-01*α + 4.387e-01
Drag: Cd = f(α) = -1.968e-05*α
3 + 3.107e-04*α
2 + 8.497e-04*α + -5.341e-04
Moment: Cm = f(α) = -6.382e-06*α
3 + 8.851e-05*α
2 + -7.040e-04*α + -6.171e-02