NACA65-116 a=0.0
NACA65116a=0.0
Curve fits of basic Javafoil data for NACA65-116 a=0.0, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.315e-04*α
3 + 5.646e-04*α
2 + 1.233e-01*α + 5.279e-02
Drag: Cd = f(α) = -1.660e-05*α
3 + 3.143e-04*α
2 + 1.973e-04*α + 5.154e-03
Moment: Cm = f(α) = 3.069e-06*α
3 + -1.801e-05*α
2 + -1.853e-03*α + -7.577e-03
Lift: Cl = f(α) = -1.976e-04*α
3 + -9.623e-05*α
2 + 1.252e-01*α + 6.269e-02
Drag: Cd = f(α) = 7.492e-07*α
3 + 1.168e-04*α
2 + -2.342e-05*α + 6.723e-03
Moment: Cm = f(α) = 8.928e-07*α
3 + 3.152e-06*α
2 + -1.838e-03*α + -7.938e-03
Lift: Cl = f(α) = -2.066e-04*α
3 + -4.365e-05*α
2 + 1.267e-01*α + 6.102e-02
Drag: Cd = f(α) = 4.976e-07*α
3 + 9.003e-05*α
2 + -5.147e-06*α + 5.636e-03
Moment: Cm = f(α) = 8.215e-07*α
3 + 7.623e-07*α
2 + -1.841e-03*α + -7.903e-03
Lift: Cl = f(α) = -2.064e-04*α
3 + -3.726e-05*α
2 + 1.271e-01*α + 6.088e-02
Drag: Cd = f(α) = 2.948e-07*α
3 + 7.529e-05*α
2 + 1.103e-05*α + 5.299e-03
Moment: Cm = f(α) = 8.070e-07*α
3 + 1.064e-06*α
2 + -1.857e-03*α + -7.923e-03