NACA66-516 a=0.2
NACA66516a=0.2
Curve fits of basic Javafoil data for NACA66-516 a=0.2, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.171e-04*α
3 + 5.328e-04*α
2 + 1.199e-01*α + 3.296e-01
Drag: Cd = f(α) = -3.131e-06*α
3 + 3.447e-04*α
2 + -1.892e-03*α + 1.133e-02
Moment: Cm = f(α) = 7.729e-06*α
3 + 2.430e-05*α
2 + -3.615e-03*α + -4.268e-02
Lift: Cl = f(α) = -2.335e-04*α
3 + 5.675e-04*α
2 + 1.222e-01*α + 3.274e-01
Drag: Cd = f(α) = -8.785e-06*α
3 + 3.492e-04*α
2 + -1.108e-03*α + 6.654e-03
Moment: Cm = f(α) = 2.879e-06*α
3 + 4.667e-05*α
2 + -3.002e-03*α + -4.601e-02
Lift: Cl = f(α) = -2.662e-04*α
3 + 5.979e-04*α
2 + 1.275e-01*α + 3.168e-01
Drag: Cd = f(α) = -2.056e-05*α
3 + 3.245e-04*α
2 + 6.357e-04*α + 1.589e-03
Moment: Cm = f(α) = -2.839e-06*α
3 + 4.861e-05*α
2 + -2.158e-03*α + -4.822e-02
Lift: Cl = f(α) = -2.706e-04*α
3 + 5.936e-04*α
2 + 1.286e-01*α + 3.157e-01
Drag: Cd = f(α) = -2.206e-05*α
3 + 2.989e-04*α
2 + 9.240e-04*α + 1.050e-03
Moment: Cm = f(α) = -3.048e-06*α
3 + 4.382e-05*α
2 + -2.076e-03*α + -4.830e-02