XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0610 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   0.3407   0.01223   0.00775  -0.0848  0.9496  0.7638
   0.500   0.4212   0.01143   0.00707  -0.0885  0.9306  0.7684
   1.000   0.4932   0.01074   0.00641  -0.0908  0.8882  0.7743
   1.500   0.5465   0.01073   0.00595  -0.0882  0.7247  0.7779
   2.000   0.5758   0.01369   0.00664  -0.0832  0.2564  0.7828
   2.500   0.6270   0.01528   0.00738  -0.0831  0.1099  0.7877
   3.000   0.6772   0.01622   0.00820  -0.0819  0.0845  0.7918
   3.500   0.7326   0.01708   0.00905  -0.0820  0.0693  0.7973
   4.000   0.7847   0.01806   0.01003  -0.0812  0.0582  0.8015
   4.500   0.8355   0.01929   0.01128  -0.0802  0.0499  0.8064
   5.000   0.8897   0.02076   0.01279  -0.0800  0.0434  0.8115
   5.500   0.9394   0.02265   0.01477  -0.0787  0.0394  0.8154
   6.000   0.9920   0.02425   0.01653  -0.0781  0.0359  0.8210
   6.500   1.0420   0.02717   0.01975  -0.0771  0.0339  0.8255
   7.000   1.0874   0.03090   0.02412  -0.0749  0.0325  0.8304
   7.500   1.1296   0.03488   0.02863  -0.0730  0.0311  0.8359
   8.000   1.1663   0.03840   0.03240  -0.0705  0.0300  0.8410
   8.500   1.1312   0.05701   0.05299  -0.0597  0.0355  0.8458