XFOIL         Version 6.94
  
 Calculated polar for: USA 41 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   0.4906   0.00943   0.00229  -0.1031  0.7936  0.0585
   0.500   0.5419   0.00912   0.00209  -0.1016  0.7516  0.1356
   1.000   0.5932   0.00927   0.00212  -0.1002  0.7058  0.2037
   1.500   0.6436   0.00935   0.00213  -0.0988  0.6600  0.2526
   2.500   0.7506   0.00888   0.00240  -0.0980  0.5161  1.0000
   3.000   0.8019   0.00927   0.00269  -0.0969  0.4871  1.0000
   3.500   0.8462   0.01010   0.00292  -0.0947  0.3821  1.0000
   4.000   0.8934   0.01087   0.00334  -0.0931  0.3112  1.0000
   4.500   0.9230   0.01388   0.00486  -0.0895  0.0364  1.0000
   5.500   1.0195   0.01548   0.00647  -0.0863  0.0076  1.0000
   6.000   1.0670   0.01636   0.00762  -0.0845  0.0077  1.0000
   6.500   1.1119   0.01752   0.00913  -0.0822  0.0083  1.0000
   7.000   1.1503   0.01936   0.01140  -0.0788  0.0089  1.0000
   7.500   1.1775   0.02213   0.01448  -0.0739  0.0098  1.0000
   8.000   1.1987   0.02598   0.01856  -0.0683  0.0106  1.0000
   8.500   1.2346   0.02936   0.02213  -0.0647  0.0120  1.0000
   9.000   1.2943   0.03929   0.03257  -0.0637  0.0194  1.0000
   9.500   1.3267   0.05123   0.04614  -0.0557  0.0434  1.0000
  10.000   1.3262   0.05746   0.05292  -0.0499  0.0405  1.0000
  10.500   1.3213   0.06456   0.06035  -0.0452  0.0388  1.0000