XFOIL         Version 6.94
  
 Calculated polar for: NACA 64-210                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1647   0.00839   0.00351  -0.0389  0.7955  0.7975
   0.500   0.2171   0.00828   0.00348  -0.0382  0.7796  0.8185
   1.000   0.2691   0.00834   0.00351  -0.0372  0.7625  0.8383
   1.500   0.3224   0.00847   0.00365  -0.0368  0.7450  0.8576
   2.000   0.3709   0.00843   0.00371  -0.0349  0.7273  0.8801
   2.500   0.4194   0.00840   0.00357  -0.0328  0.6915  0.8996
   3.000   0.4614   0.00816   0.00308  -0.0288  0.6161  0.9295
   3.500   0.5104   0.00837   0.00312  -0.0273  0.5540  0.9542
   4.000   0.5625   0.01089   0.00387  -0.0290  0.1595  0.9854
   4.500   0.6026   0.01325   0.00569  -0.0275  0.0308  1.0000
   5.000   0.6432   0.01513   0.00776  -0.0253  0.0277  1.0000
   5.500   0.6872   0.01727   0.01003  -0.0236  0.0274  1.0000
   6.000   0.7339   0.02009   0.01337  -0.0222  0.0207  1.0000
   6.500   0.7813   0.02106   0.01439  -0.0215  0.0160  1.0000
   7.000   0.8256   0.02237   0.01571  -0.0206  0.0140  1.0000
   7.500   0.8622   0.02765   0.02140  -0.0185  0.0120  1.0000
   8.000   0.8905   0.03258   0.02688  -0.0151  0.0112  1.0000
   8.500   0.9097   0.03723   0.03202  -0.0108  0.0108  1.0000
   9.000   0.9006   0.04389   0.03927  -0.0038  0.0109  1.0000