XFOIL         Version 6.94
  
 Calculated polar for: NACA 2410                                       
  
 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   0.2342   0.00761   0.00199  -0.0510  0.7240  0.6636
   0.500   0.2833   0.00741   0.00207  -0.0495  0.6800  0.7970
   1.000   0.3479   0.00750   0.00218  -0.0510  0.6336  0.9341
   1.500   0.4250   0.00773   0.00224  -0.0561  0.5909  0.9902
   2.000   0.4834   0.00803   0.00230  -0.0573  0.5495  1.0000
   2.500   0.5308   0.00841   0.00243  -0.0560  0.4996  1.0000
   3.000   0.5792   0.00875   0.00260  -0.0550  0.4592  1.0000
   3.500   0.6221   0.00952   0.00285  -0.0531  0.3621  1.0000
   4.000   0.6463   0.01255   0.00401  -0.0488  0.0605  1.0000
   4.500   0.6947   0.01324   0.00470  -0.0477  0.0454  1.0000
   5.000   0.7430   0.01398   0.00555  -0.0466  0.0442  1.0000
   5.500   0.7893   0.01497   0.00667  -0.0452  0.0440  1.0000
   6.000   0.8334   0.01623   0.00807  -0.0435  0.0442  1.0000
   6.500   0.8757   0.01785   0.00986  -0.0414  0.0446  1.0000
   7.000   0.9181   0.01975   0.01191  -0.0395  0.0448  1.0000
   7.500   0.9634   0.02080   0.01299  -0.0385  0.0414  1.0000
   8.000   1.0064   0.02254   0.01486  -0.0370  0.0397  1.0000
   8.500   1.0483   0.02448   0.01694  -0.0355  0.0380  1.0000
   9.000   1.0886   0.02634   0.01894  -0.0338  0.0355  1.0000
   9.500   1.1261   0.02780   0.02045  -0.0321  0.0319  1.0000
  10.500   1.1364   0.04106   0.03547  -0.0198  0.0174  1.0000
  11.000   1.1042   0.04951   0.04456  -0.0121  0.0181  1.0000
  11.500   1.0674   0.05780   0.05332  -0.0086  0.0184  1.0000
  12.000   1.0314   0.06750   0.06338  -0.0101  0.0185  1.0000
  12.500   0.9982   0.07867   0.07486  -0.0156  0.0184  1.0000
  13.000   0.9692   0.09119   0.08760  -0.0234  0.0182  1.0000