NACA63-214 a=0.8
NACA63214a=0.8
Curve fits of basic Javafoil data for NACA63-214 a=0.8, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.361e-04*α
3 + 1.241e-04*α
2 + 1.238e-01*α + 1.843e-01
Drag: Cd = f(α) = -1.907e-05*α
3 + 2.994e-04*α
2 + 7.653e-04*α + 3.868e-03
Moment: Cm = f(α) = -8.681e-07*α
3 + 5.110e-05*α
2 + -1.163e-03*α + -3.810e-02
Lift: Cl = f(α) = -2.183e-04*α
3 + -2.553e-04*α
2 + 1.255e-01*α + 1.903e-01
Drag: Cd = f(α) = -6.144e-07*α
3 + 1.235e-04*α
2 + 1.559e-04*α + 5.993e-03
Moment: Cm = f(α) = 1.837e-07*α
3 + 1.816e-05*α
2 + -1.058e-03*α + -3.766e-02
Lift: Cl = f(α) = -2.199e-04*α
3 + -2.765e-04*α
2 + 1.266e-01*α + 1.912e-01
Drag: Cd = f(α) = -8.286e-07*α
3 + 9.423e-05*α
2 + 1.710e-04*α + 5.222e-03
Moment: Cm = f(α) = 1.944e-07*α
3 + 1.363e-05*α
2 + -1.057e-03*α + -3.766e-02
Lift: Cl = f(α) = -2.191e-04*α
3 + -2.718e-04*α
2 + 1.270e-01*α + 1.914e-01
Drag: Cd = f(α) = -1.355e-06*α
3 + 7.751e-05*α
2 + 2.156e-04*α + 5.112e-03
Moment: Cm = f(α) = 1.517e-07*α
3 + 1.129e-05*α
2 + -1.066e-03*α + -3.766e-02