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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2627   0.00864   0.00439  -0.0759  0.8776  0.7180
   0.500   0.3176   0.00869   0.00435  -0.0749  0.8393  0.7227
   1.000   0.3725   0.00883   0.00437  -0.0739  0.7885  0.7267
   1.500   0.4287   0.00904   0.00432  -0.0734  0.7131  0.7306
   2.000   0.4763   0.01025   0.00452  -0.0716  0.4748  0.7340
   2.500   0.5271   0.01183   0.00502  -0.0714  0.2412  0.7387
   3.000   0.5803   0.01283   0.00548  -0.0713  0.1398  0.7420
   3.500   0.6350   0.01349   0.00597  -0.0710  0.1040  0.7456
   4.000   0.6916   0.01408   0.00642  -0.0712  0.0849  0.7502
   4.500   0.7454   0.01470   0.00706  -0.0707  0.0742  0.7533
   5.500   0.8530   0.01628   0.00861  -0.0700  0.0592  0.7613
   6.000   0.9064   0.01689   0.00929  -0.0694  0.0530  0.7651
   7.000   1.0081   0.01904   0.01155  -0.0677  0.0396  0.7731
   7.500   1.0591   0.02006   0.01272  -0.0668  0.0368  0.7776
   8.000   1.1079   0.02135   0.01413  -0.0655  0.0346  0.7820
   8.500   1.1550   0.02271   0.01563  -0.0641  0.0329  0.7868
   9.000   1.1976   0.02507   0.01822  -0.0621  0.0309  0.7916
   9.500   1.2368   0.02799   0.02150  -0.0597  0.0296  0.7963
  10.000   1.2789   0.02962   0.02339  -0.0577  0.0284  0.8014
  10.500   1.3139   0.03203   0.02614  -0.0548  0.0267  0.8065
  11.000   1.3400   0.03533   0.02985  -0.0510  0.0254  0.8120
  11.500   1.3520   0.03958   0.03458  -0.0458  0.0247  0.8175
  12.000   1.3373   0.04472   0.04025  -0.0379  0.0244  0.8231
  12.500   1.3063   0.05182   0.04788  -0.0317  0.0242  0.8295
  13.000   1.2592   0.06189   0.05848  -0.0300  0.0241  0.8344