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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.6794   0.01131   0.00371  -0.1227  0.6845  0.0658
   1.000   0.7304   0.01089   0.00323  -0.1216  0.6553  0.0957
   1.500   0.7944   0.00911   0.00338  -0.1238  0.6244  1.0000
   2.000   0.8425   0.00951   0.00347  -0.1222  0.5909  1.0000
   2.500   0.8896   0.00991   0.00362  -0.1204  0.5534  1.0000
   3.000   0.9355   0.01037   0.00380  -0.1185  0.5129  1.0000
   3.500   0.9815   0.01091   0.00409  -0.1167  0.4797  1.0000
   4.000   1.0201   0.01178   0.00455  -0.1136  0.4232  1.0000
   4.500   1.0615   0.01255   0.00502  -0.1112  0.3768  1.0000
   5.000   1.1048   0.01323   0.00549  -0.1092  0.3396  1.0000
   5.500   1.1490   0.01387   0.00599  -0.1074  0.3056  1.0000
   6.000   1.1806   0.01529   0.00687  -0.1037  0.2164  1.0000
   6.500   1.2052   0.01722   0.00828  -0.0991  0.1399  1.0000
   7.000   1.2097   0.02030   0.01067  -0.0914  0.0064  1.0000
   7.500   1.2444   0.02141   0.01192  -0.0884  0.0058  1.0000
   8.000   1.2771   0.02269   0.01334  -0.0852  0.0058  1.0000
   8.500   1.3066   0.02422   0.01508  -0.0818  0.0060  1.0000
   9.000   1.3314   0.02616   0.01724  -0.0782  0.0063  1.0000
   9.500   1.3487   0.02877   0.02013  -0.0741  0.0066  1.0000
  10.000   1.3655   0.03158   0.02318  -0.0705  0.0071  1.0000
  10.500   1.3718   0.03548   0.02735  -0.0667  0.0077  1.0000
  11.000   1.3663   0.04080   0.03292  -0.0631  0.0082  1.0000
  11.500   1.3517   0.04757   0.03993  -0.0603  0.0085  1.0000
  12.000   1.3504   0.05352   0.04614  -0.0587  0.0092  1.0000
  12.500   1.3381   0.06093   0.05377  -0.0572  0.0100  1.0000
  13.000   1.3339   0.06677   0.05962  -0.0544  0.0108  1.0000