XFOIL         Version 6.94
  
 Calculated polar for: NACA 64A410                                     
  
 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   0.3408   0.00784   0.00246  -0.0804  0.7937  0.7484
   0.500   0.3928   0.00767   0.00236  -0.0795  0.7712  0.7869
   1.000   0.4459   0.00758   0.00228  -0.0787  0.7469  0.8237
   1.500   0.5016   0.00748   0.00230  -0.0785  0.7216  0.8932
   2.000   0.5794   0.00758   0.00222  -0.0834  0.6710  1.0000
   2.500   0.6235   0.00808   0.00226  -0.0811  0.5905  1.0000
   3.000   0.6414   0.01055   0.00298  -0.0747  0.2558  1.0000
   3.500   0.6832   0.01182   0.00365  -0.0730  0.1386  1.0000
   4.000   0.7323   0.01242   0.00414  -0.0722  0.1195  1.0000
   4.500   0.7824   0.01292   0.00473  -0.0716  0.0951  1.0000
   5.000   0.8200   0.01487   0.00657  -0.0685  0.0275  1.0000
   5.500   0.8565   0.01696   0.00896  -0.0651  0.0245  1.0000
   6.000   0.8930   0.01941   0.01154  -0.0619  0.0242  1.0000
   6.500   0.9366   0.02180   0.01408  -0.0599  0.0247  1.0000
   8.500   1.0909   0.03515   0.02890  -0.0510  0.0100  1.0000
   9.000   1.1148   0.03989   0.03412  -0.0471  0.0097  1.0000
   9.500   1.1062   0.04885   0.04387  -0.0398  0.0100  1.0000