XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 417 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   0.2474   0.01896   0.01362  -0.0920  0.8805  0.6705
   1.500   0.4623   0.01709   0.01182  -0.1017  0.8608  0.6863
   2.000   0.5154   0.01654   0.01136  -0.1010  0.8464  0.6908
   3.500   0.7313   0.01336   0.00853  -0.1090  0.8119  0.7065
   4.000   0.8108   0.01217   0.00745  -0.1128  0.7799  0.7108
   4.500   0.8888   0.01185   0.00662  -0.1162  0.6533  0.7163
   5.000   0.9150   0.01288   0.00712  -0.1104  0.5449  0.7220
   5.500   0.9268   0.01433   0.00795  -0.1023  0.4218  0.7269
   6.000   0.9457   0.01579   0.00891  -0.0961  0.3082  0.7315
   6.500   0.9687   0.01757   0.01008  -0.0910  0.1899  0.7371
   7.000   0.9969   0.01950   0.01142  -0.0873  0.0878  0.7434
   7.500   1.0210   0.02193   0.01336  -0.0829  0.0165  0.7483
   8.000   1.0534   0.02348   0.01509  -0.0795  0.0128  0.7534
   9.000   1.1124   0.02718   0.01919  -0.0724  0.0111  0.7660
   9.500   1.1341   0.02961   0.02178  -0.0682  0.0107  0.7718
  10.000   1.1540   0.03251   0.02486  -0.0640  0.0105  0.7776
  10.500   1.1806   0.03574   0.02825  -0.0611  0.0104  0.7841
  11.000   1.2165   0.03915   0.03188  -0.0595  0.0104  0.7903
  11.500   1.2469   0.04264   0.03578  -0.0570  0.0106  0.7976
  12.000   1.2653   0.04785   0.04173  -0.0535  0.0111  0.8052
  12.500   1.2644   0.05512   0.04974  -0.0493  0.0119  0.8114
  13.000   1.2513   0.06248   0.05766  -0.0456  0.0123  0.8180
  13.500   1.2341   0.07065   0.06624  -0.0436  0.0126  0.8251
  14.500   1.1150   0.10040   0.09742  -0.0501  0.0147  0.8356
  15.000   1.0752   0.11603   0.11340  -0.0593  0.0150  0.8432