XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0412 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   0.2364   0.01191   0.00735  -0.0693  0.9225  0.7407
   0.500   0.2897   0.01169   0.00713  -0.0672  0.8913  0.7451
   1.000   0.3481   0.01142   0.00679  -0.0670  0.8540  0.7511
   1.500   0.4012   0.01137   0.00665  -0.0654  0.8040  0.7551
   2.000   0.4597   0.01137   0.00636  -0.0655  0.7176  0.7605
   2.500   0.5009   0.01273   0.00654  -0.0621  0.4449  0.7636
   3.000   0.5483   0.01452   0.00712  -0.0615  0.2016  0.7681
   3.500   0.6013   0.01557   0.00771  -0.0615  0.1306  0.7719
   4.000   0.6542   0.01637   0.00839  -0.0609  0.1052  0.7763
   4.500   0.7103   0.01715   0.00907  -0.0614  0.0890  0.7809
   5.000   0.7618   0.01799   0.00990  -0.0605  0.0788  0.7845
   5.500   0.8159   0.01909   0.01093  -0.0606  0.0699  0.7894
   6.000   0.8680   0.01987   0.01182  -0.0598  0.0627  0.7931
   7.500   1.0202   0.02385   0.01614  -0.0575  0.0439  0.8063
   8.000   1.0688   0.02594   0.01831  -0.0566  0.0411  0.8114
   8.500   1.1129   0.02920   0.02194  -0.0549  0.0394  0.8164
   9.000   1.1575   0.03158   0.02471  -0.0534  0.0380  0.8216
   9.500   1.1954   0.03448   0.02813  -0.0509  0.0360  0.8268
  10.000   1.2284   0.03752   0.03159  -0.0482  0.0339  0.8321
  10.500   1.2551   0.04052   0.03491  -0.0449  0.0323  0.8378
  11.000   1.2648   0.04533   0.04014  -0.0402  0.0312  0.8435
  11.500   1.2367   0.05259   0.04801  -0.0324  0.0307  0.8501
  12.000   1.1936   0.06067   0.05667  -0.0267  0.0307  0.8555
  12.500   1.1410   0.07205   0.06859  -0.0280  0.0309  0.8613