XFOIL         Version 6.94
  
 Calculated polar for: RAF 32 MOD 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   0.5017   0.00714   0.00226  -0.1199  0.7891  0.7024
   0.500   0.5489   0.00667   0.00232  -0.1175  0.7731  0.8779
   1.000   0.6314   0.00666   0.00227  -0.1233  0.7579  1.0000
   1.500   0.6831   0.00680   0.00232  -0.1223  0.7400  1.0000
   2.000   0.7333   0.00695   0.00237  -0.1209  0.7146  1.0000
   2.500   0.7766   0.00719   0.00233  -0.1179  0.6570  1.0000
   3.000   0.8226   0.00755   0.00252  -0.1157  0.6110  1.0000
   3.500   0.8556   0.00847   0.00281  -0.1111  0.4902  1.0000
   4.000   0.8758   0.01055   0.00369  -0.1047  0.2707  1.0000
   4.500   0.9055   0.01233   0.00470  -0.1004  0.1222  1.0000
   5.000   0.9410   0.01365   0.00566  -0.0970  0.0413  1.0000
   5.500   0.9797   0.01469   0.00648  -0.0940  0.0090  1.0000
   6.000   1.0241   0.01528   0.00711  -0.0919  0.0061  1.0000
   6.500   1.0637   0.01602   0.00794  -0.0889  0.0058  1.0000
   7.000   1.0968   0.01711   0.00923  -0.0848  0.0043  1.0000
   7.500   1.1281   0.01834   0.01065  -0.0805  0.0041  1.0000
   8.000   1.1562   0.01984   0.01243  -0.0758  0.0041  1.0000
   8.500   1.1766   0.02192   0.01478  -0.0703  0.0043  1.0000
   9.000   1.1942   0.02434   0.01743  -0.0649  0.0046  1.0000
   9.500   1.2031   0.02809   0.02150  -0.0587  0.0051  1.0000
  10.000   1.1679   0.02426   0.01825  -0.0492  0.0057  1.0000
  10.500   1.1911   0.02776   0.02202  -0.0457  0.0071  1.0000