XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0410 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   0.2182   0.01077   0.00621  -0.0647  0.9426  0.7558
   0.500   0.2765   0.01063   0.00610  -0.0635  0.9077  0.7654
   1.000   0.3308   0.01047   0.00590  -0.0619  0.8629  0.7697
   1.500   0.3876   0.01043   0.00574  -0.0613  0.8114  0.7753
   2.000   0.4398   0.01065   0.00564  -0.0595  0.7045  0.7799
   2.500   0.4850   0.01201   0.00580  -0.0572  0.4344  0.7841
   3.000   0.5347   0.01377   0.00634  -0.0571  0.1873  0.7887
   3.500   0.5863   0.01489   0.00701  -0.0565  0.1149  0.7923
   4.000   0.6418   0.01578   0.00774  -0.0567  0.0917  0.7972
   4.500   0.6956   0.01659   0.00851  -0.0563  0.0798  0.8010
   5.000   0.7483   0.01762   0.00950  -0.0558  0.0710  0.8057
   5.500   0.8023   0.01878   0.01066  -0.0556  0.0643  0.8103
   6.000   0.8540   0.02001   0.01191  -0.0548  0.0588  0.8144
   6.500   0.9077   0.02150   0.01338  -0.0547  0.0542  0.8200
   7.000   0.9586   0.02276   0.01482  -0.0536  0.0499  0.8246
   7.500   1.0109   0.02464   0.01691  -0.0532  0.0464  0.8304
   8.000   1.0595   0.02668   0.01906  -0.0521  0.0437  0.8354
   8.500   1.1064   0.02914   0.02201  -0.0507  0.0411  0.8417
   9.000   1.1525   0.03188   0.02480  -0.0498  0.0393  0.8482
   9.500   1.1827   0.03619   0.03008  -0.0461  0.0372  0.8543
  10.000   1.2242   0.03790   0.03182  -0.0448  0.0352  0.8617
  11.000   1.2179   0.05279   0.04844  -0.0338  0.0342  0.8777
  11.500   1.1596   0.06173   0.05814  -0.0262  0.0338  0.8882
  12.500   0.9847   0.10668   0.10406  -0.0522  0.0348  0.9027
  13.500   0.8737   0.16567   0.16294  -0.0882  0.0296  0.8822
  14.000   0.8539   0.18417   0.18156  -0.0957  0.0410  0.9071