XFOIL         Version 6.94
  
 Calculated polar for: NACA 64-108 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.0854   0.00590   0.00176  -0.0204  0.8414  0.8446
   0.500   0.1378   0.00591   0.00178  -0.0194  0.8189  0.8682
   1.000   0.1895   0.00594   0.00183  -0.0182  0.7965  0.8932
   1.500   0.2388   0.00595   0.00186  -0.0163  0.7671  0.9184
   2.000   0.2861   0.00600   0.00174  -0.0138  0.7015  0.9464
   2.500   0.3456   0.00617   0.00182  -0.0143  0.6349  0.9724
   3.000   0.4048   0.00896   0.00238  -0.0173  0.1041  0.9970
   3.500   0.4538   0.00981   0.00289  -0.0168  0.0365  1.0000
   4.000   0.5057   0.01038   0.00333  -0.0166  0.0185  1.0000
   4.500   0.5586   0.01095   0.00391  -0.0164  0.0129  1.0000
   5.000   0.6114   0.01156   0.00460  -0.0162  0.0105  1.0000
   5.500   0.6638   0.01224   0.00539  -0.0158  0.0064  1.0000
   6.000   0.7145   0.01326   0.00655  -0.0152  0.0037  1.0000
   6.500   0.7632   0.01475   0.00829  -0.0140  0.0034  1.0000
   7.000   0.8078   0.01736   0.01136  -0.0122  0.0035  1.0000
   7.500   0.8424   0.02411   0.01914  -0.0088  0.0038  1.0000
   8.000   0.8107   0.04945   0.04649   0.0004  0.0056  1.0000