XFOIL         Version 6.94
  
 Calculated polar for: NACA 65-410                                     
  
 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   0.3192   0.00922   0.00367  -0.0779  0.7777  0.7457
   0.500   0.3747   0.00914   0.00344  -0.0774  0.7561  0.7535
   1.000   0.4280   0.00908   0.00330  -0.0764  0.7329  0.7622
   1.500   0.4792   0.00969   0.00431  -0.0746  0.7410  0.8247
   2.000   0.5420   0.00936   0.00370  -0.0767  0.7158  0.7954
   2.500   0.5834   0.00954   0.00414  -0.0721  0.6950  0.8537
   3.000   0.6461   0.00950   0.00405  -0.0738  0.6755  0.8434
   3.500   0.6892   0.00949   0.00406  -0.0702  0.6329  0.8753
   4.000   0.7367   0.00963   0.00438  -0.0679  0.6122  0.9026
   4.500   0.7930   0.00987   0.00438  -0.0683  0.5327  0.8933
   5.000   0.8349   0.01176   0.00482  -0.0677  0.3094  0.8595
   5.500   0.8717   0.01434   0.00631  -0.0661  0.1199  0.8240
   6.500   0.9258   0.01716   0.00898  -0.0555  0.0376  0.9508
   7.000   0.9931   0.01897   0.01063  -0.0597  0.0439  0.8367
   9.000   1.1236   0.02433   0.01645  -0.0465  0.0315  1.0000
  11.000   1.2896   0.03679   0.02982  -0.0403  0.0296  1.0000
  11.500   1.3126   0.04031   0.03356  -0.0372  0.0273  1.0000
  12.000   1.3155   0.04354   0.03726  -0.0321  0.0259  1.0000
  13.000   1.0986   0.03921   0.03409  -0.0044  0.0258  1.0000
  14.000   1.2062   0.07828   0.07429  -0.0231  0.0269  1.0000
  14.500   1.1644   0.08993   0.08630  -0.0269  0.0271  1.0000