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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2321   0.00801   0.00398  -0.0682  0.9380  0.7456
   0.500   0.2903   0.00792   0.00382  -0.0679  0.8956  0.7509
   1.000   0.3427   0.00804   0.00376  -0.0661  0.8266  0.7549
   1.500   0.3965   0.00840   0.00369  -0.0651  0.7119  0.7592
   2.000   0.4426   0.01006   0.00397  -0.0633  0.3903  0.7621
   2.500   0.4933   0.01154   0.00451  -0.0630  0.1638  0.7659
   3.000   0.5491   0.01242   0.00498  -0.0633  0.0919  0.7703
   3.500   0.6036   0.01307   0.00551  -0.0631  0.0695  0.7734
   4.000   0.6593   0.01368   0.00610  -0.0631  0.0582  0.7774
   4.500   0.7136   0.01452   0.00697  -0.0628  0.0500  0.7815
   5.000   0.7689   0.01494   0.00748  -0.0627  0.0432  0.7853
   7.000   0.9733   0.02099   0.01416  -0.0588  0.0243  0.8036
   7.500   1.0209   0.02386   0.01738  -0.0573  0.0238  0.8089
   8.000   1.0668   0.02614   0.01997  -0.0558  0.0227  0.8138
   8.500   1.1081   0.02960   0.02386  -0.0539  0.0221  0.8201
   9.000   1.1394   0.03432   0.02918  -0.0508  0.0218  0.8253
   9.500   1.1591   0.04020   0.03575  -0.0469  0.0214  0.8315
  10.000   1.1563   0.04797   0.04423  -0.0417  0.0210  0.8380
  10.500   1.1205   0.05681   0.05370  -0.0348  0.0208  0.8449