XFOIL         Version 6.94
  
 Calculated polar for: FX 66-S-196 AIRFOIL                             
  
 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   0.4470   0.01646   0.00960  -0.0959  0.6057  0.6451
   0.500   0.5037   0.01657   0.00965  -0.0968  0.6029  0.6476
   1.000   0.5614   0.01679   0.00979  -0.0979  0.5996  0.6494
   1.500   0.6144   0.01717   0.01018  -0.0984  0.5966  0.6512
   2.000   0.6659   0.01749   0.01057  -0.0986  0.5935  0.6529
   2.500   0.7173   0.01787   0.01102  -0.0988  0.5901  0.6547
   3.000   0.7694   0.01826   0.01145  -0.0991  0.5866  0.6576
   3.500   0.8242   0.01849   0.01169  -0.0997  0.5830  0.6607
   4.000   0.8810   0.01870   0.01192  -0.1007  0.5799  0.6630
   4.500   0.9349   0.01952   0.01279  -0.1014  0.5768  0.6655
   5.000   0.9729   0.02030   0.01383  -0.0994  0.5734  0.6690
   5.500   1.0119   0.02106   0.01478  -0.0976  0.5683  0.6725
   6.000   1.0586   0.02155   0.01540  -0.0970  0.5638  0.6766
   6.500   1.1161   0.02157   0.01551  -0.0978  0.5599  0.6813
   7.000   1.1770   0.02180   0.01583  -0.0994  0.5566  0.6868
   7.500   1.1939   0.02326   0.01761  -0.0942  0.5507  0.6934
   8.000   1.2250   0.02375   0.01830  -0.0909  0.5435  0.7021
   8.500   1.3437   0.02044   0.01492  -0.1000  0.5336  0.7132
   9.000   1.3626   0.01994   0.01470  -0.0936  0.5210  0.7273
  10.000   1.4087   0.01828   0.01337  -0.0814  0.4895  0.7761
  10.500   1.3948   0.01905   0.01459  -0.0709  0.4731  0.8549