XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0612 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   0.4010   0.01225   0.00772  -0.0966  0.9103  0.7320
   0.500   0.4588   0.01192   0.00744  -0.0960  0.8842  0.7368
   1.000   0.5190   0.01156   0.00705  -0.0960  0.8412  0.7420
   1.500   0.5731   0.01157   0.00679  -0.0943  0.7384  0.7459
   2.000   0.6092   0.01368   0.00711  -0.0904  0.3814  0.7512
   2.500   0.6490   0.01566   0.00794  -0.0879  0.1599  0.7551
   3.000   0.7043   0.01665   0.00857  -0.0883  0.1120  0.7606
   3.500   0.7549   0.01736   0.00925  -0.0872  0.0943  0.7638
   4.000   0.8098   0.01814   0.00997  -0.0872  0.0829  0.7687
   5.500   0.9629   0.02125   0.01311  -0.0846  0.0554  0.7820
   6.000   1.0106   0.02295   0.01491  -0.0831  0.0482  0.7863
   6.500   1.0628   0.02411   0.01619  -0.0825  0.0436  0.7911
   7.000   1.1096   0.02591   0.01807  -0.0810  0.0409  0.7952
   7.500   1.1583   0.02934   0.02178  -0.0801  0.0393  0.8007
   8.000   1.2022   0.03160   0.02441  -0.0779  0.0383  0.8050
   8.500   1.2448   0.03456   0.02779  -0.0761  0.0369  0.8104
   9.000   1.2775   0.03777   0.03155  -0.0726  0.0351  0.8145
   9.500   1.3063   0.04115   0.03533  -0.0692  0.0333  0.8202
  10.000   1.3243   0.04446   0.03900  -0.0643  0.0319  0.8248
  10.500   1.3257   0.04948   0.04447  -0.0583  0.0311  0.8307
  11.000   1.2929   0.05589   0.05142  -0.0489  0.0308  0.8357
  11.500   1.2457   0.06483   0.06094  -0.0430  0.0308  0.8406