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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4167   0.00832   0.00432  -0.1017  0.9233  0.7402
   0.500   0.4728   0.00819   0.00417  -0.1006  0.8749  0.7435
   1.000   0.5225   0.00871   0.00401  -0.0981  0.6976  0.7483
   1.500   0.5596   0.01116   0.00470  -0.0947  0.3031  0.7523
   2.000   0.6083   0.01263   0.00530  -0.0938  0.1204  0.7563
   2.500   0.6631   0.01340   0.00582  -0.0937  0.0809  0.7610
   3.000   0.7166   0.01400   0.00638  -0.0931  0.0636  0.7646
   3.500   0.7721   0.01469   0.00706  -0.0931  0.0528  0.7699
   5.000   0.9257   0.01794   0.01046  -0.0899  0.0293  0.7822
   5.500   0.9767   0.01914   0.01187  -0.0887  0.0264  0.7866
   6.000   1.0262   0.02109   0.01403  -0.0873  0.0247  0.7914
   6.500   1.0735   0.02375   0.01701  -0.0854  0.0241  0.7955
   7.000   1.1204   0.02663   0.02024  -0.0838  0.0233  0.8007
   7.500   1.1627   0.02922   0.02317  -0.0815  0.0223  0.8050
   8.000   1.1981   0.03372   0.02823  -0.0785  0.0220  0.8104
   8.500   1.2215   0.03909   0.03423  -0.0739  0.0218  0.8145
   9.000   1.2328   0.04531   0.04110  -0.0687  0.0213  0.8200
   9.500   1.2196   0.05291   0.04933  -0.0615  0.0209  0.8239
  10.000   1.1756   0.06100   0.05794  -0.0524  0.0208  0.8295
  10.500   1.1112   0.07301   0.07048  -0.0497  0.0210  0.8334