XFOIL         Version 6.94
  
 Calculated polar for: GOE 330 (PFALZ 59) 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   0.6256   0.00965   0.00247  -0.1224  0.6633  0.0377
   0.500   0.6775   0.00985   0.00260  -0.1217  0.6357  0.0538
   1.000   0.7296   0.00988   0.00259  -0.1208  0.6109  0.0746
   1.500   0.7805   0.00976   0.00255  -0.1200  0.5813  0.1433
   2.500   0.8878   0.00881   0.00301  -0.1199  0.4940  1.0000
   3.000   0.9341   0.00935   0.00326  -0.1182  0.4584  1.0000
   3.500   0.9814   0.00987   0.00359  -0.1167  0.4350  1.0000
   4.000   1.0212   0.01078   0.00405  -0.1140  0.3647  1.0000
   4.500   1.0677   0.01134   0.00442  -0.1126  0.3309  1.0000
   5.000   1.1079   0.01232   0.00496  -0.1101  0.2617  1.0000
   5.500   1.1357   0.01426   0.00617  -0.1059  0.1509  1.0000
   6.000   1.1535   0.01687   0.00807  -0.1000  0.0050  1.0000
   6.500   1.1938   0.01760   0.00886  -0.0975  0.0046  1.0000
   7.000   1.2321   0.01844   0.00979  -0.0948  0.0047  1.0000
   7.500   1.2678   0.01947   0.01095  -0.0918  0.0050  1.0000
   8.000   1.3021   0.02060   0.01221  -0.0887  0.0056  1.0000
   8.500   1.3322   0.02205   0.01383  -0.0851  0.0063  1.0000
   9.000   1.3575   0.02389   0.01584  -0.0812  0.0070  1.0000
   9.500   1.3769   0.02624   0.01842  -0.0770  0.0081  1.0000
  10.000   1.3903   0.02921   0.02160  -0.0726  0.0092  1.0000
  10.500   1.3868   0.03384   0.02649  -0.0676  0.0105  1.0000
  11.000   1.3827   0.03896   0.03189  -0.0636  0.0121  1.0000
  11.500   1.3597   0.04622   0.03929  -0.0594  0.0132  1.0000