XFOIL         Version 6.94
  
 Calculated polar for: NACA 1410                                       
  
 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   0.1228   0.00795   0.00125  -0.0262  0.5843  0.3932
   0.500   0.1719   0.00793   0.00129  -0.0255  0.5164  0.4900
   1.000   0.2244   0.00790   0.00132  -0.0253  0.4807  0.5505
   1.500   0.2716   0.00792   0.00146  -0.0241  0.4024  0.6613
   2.000   0.3172   0.00816   0.00169  -0.0225  0.2998  0.7957
   2.500   0.3573   0.01000   0.00240  -0.0204  0.0462  0.8987
   3.000   0.4257   0.01040   0.00281  -0.0234  0.0393  0.9567
   3.500   0.5172   0.01086   0.00337  -0.0317  0.0385  1.0000
   4.000   0.5643   0.01128   0.00388  -0.0303  0.0386  1.0000
   4.500   0.6106   0.01186   0.00455  -0.0288  0.0389  1.0000
   5.000   0.6553   0.01264   0.00542  -0.0269  0.0395  1.0000
   5.500   0.6990   0.01365   0.00653  -0.0248  0.0402  1.0000
   6.000   0.7426   0.01494   0.00796  -0.0228  0.0410  1.0000
   6.500   0.7888   0.01586   0.00890  -0.0216  0.0379  1.0000
   7.000   0.8343   0.01711   0.01022  -0.0203  0.0364  1.0000
   7.500   0.8796   0.01850   0.01170  -0.0191  0.0349  1.0000
   8.000   0.9247   0.01983   0.01308  -0.0180  0.0328  1.0000
   8.500   0.9682   0.02131   0.01462  -0.0168  0.0298  1.0000
   9.000   0.9990   0.02614   0.01987  -0.0140  0.0251  1.0000
  10.000   1.0365   0.03762   0.03288  -0.0044  0.0162  1.0000
  10.500   1.0045   0.04798   0.04402   0.0028  0.0177  1.0000
  11.000   0.9653   0.05600   0.05245   0.0065  0.0181  1.0000
  11.500   0.9298   0.06555   0.06232   0.0031  0.0181  1.0000
  12.000   0.9011   0.07784   0.07485  -0.0059  0.0177  1.0000
  12.500   0.8786   0.09160   0.08877  -0.0159  0.0172  1.0000