NACA66-715 a=0.8
NACA66715a=0.8
Curve fits of basic Javafoil data for NACA66-715 a=0.8, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.275e-04*α
3 + -5.672e-04*α
2 + 1.222e-01*α + 6.758e-01
Drag: Cd = f(α) = -1.618e-06*α
3 + 3.819e-04*α
2 + -1.561e-03*α + 9.602e-03
Moment: Cm = f(α) = 2.768e-05*α
3 + 2.905e-04*α
2 + -7.549e-03*α + -1.320e-01
Lift: Cl = f(α) = -2.313e-04*α
3 + -6.023e-04*α
2 + 1.228e-01*α + 6.801e-01
Drag: Cd = f(α) = -5.015e-06*α
3 + 3.667e-04*α
2 + -1.025e-03*α + 5.944e-03
Moment: Cm = f(α) = 1.634e-05*α
3 + 2.894e-04*α
2 + -5.896e-03*α + -1.380e-01
Lift: Cl = f(α) = -2.522e-04*α
3 + -6.820e-04*α
2 + 1.259e-01*α + 6.804e-01
Drag: Cd = f(α) = -1.886e-05*α
3 + 2.905e-04*α
2 + 8.261e-04*α + 1.941e-03
Moment: Cm = f(α) = -9.602e-06*α
3 + 1.933e-04*α
2 + -2.321e-03*α + -1.435e-01
Lift: Cl = f(α) = -2.043e-04*α
3 + -1.010e-03*α
2 + 1.244e-01*α + 6.871e-01
Drag: Cd = f(α) = 1.027e-06*α
3 + 8.810e-05*α
2 + 3.593e-04*α + 5.222e-03
Moment: Cm = f(α) = -1.580e-06*α
3 + 3.355e-05*α
2 + -1.881e-03*α + -1.418e-01