NACA67-514 a=0.6
NACA67514a=0.6
Curve fits of basic Javafoil data for NACA67-514 a=0.6, smooth airfoils, Mach 0:
Re:
500k |
1M |
3M |
6MRe = 500k | 1M | 3M | 6M
Lift: Cl = f(α) = -2.249e-04*α
3 + -8.547e-06*α
2 + 1.201e-01*α + 4.293e-01
Drag: Cd = f(α) = -4.882e-06*α
3 + 3.748e-04*α
2 + -1.459e-03*α + 7.760e-03
Moment: Cm = f(α) = 1.036e-05*α
3 + 1.257e-04*α
2 + -4.634e-03*α + -7.799e-02
Lift: Cl = f(α) = -2.445e-04*α
3 + -4.110e-05*α
2 + 1.229e-01*α + 4.276e-01
Drag: Cd = f(α) = -1.210e-05*α
3 + 3.511e-04*α
2 + -4.234e-04*α + 3.694e-03
Moment: Cm = f(α) = 8.760e-07*α
3 + 1.220e-04*α
2 + -3.281e-03*α + -8.219e-02
Lift: Cl = f(α) = -2.125e-04*α
3 + -5.987e-04*α
2 + 1.249e-01*α + 4.345e-01
Drag: Cd = f(α) = 2.891e-06*α
3 + 9.476e-05*α
2 + 5.501e-05*α + 4.925e-03
Moment: Cm = f(α) = -2.516e-07*α
3 + 1.568e-05*α
2 + -2.155e-03*α + -8.287e-02
Lift: Cl = f(α) = -2.189e-04*α
3 + -6.067e-04*α
2 + 1.254e-01*α + 4.350e-01
Drag: Cd = f(α) = 9.485e-07*α
3 + 7.666e-05*α
2 + 1.943e-04*α + 5.089e-03
Moment: Cm = f(α) = -4.146e-07*α
3 + 8.617e-06*α
2 + -2.162e-03*α + -8.280e-02