XFOIL         Version 6.94
  
 Calculated polar for: GOE 359 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   0.6045   0.00984   0.00236  -0.1159  0.6088  0.0208
   0.500   0.6587   0.01002   0.00254  -0.1155  0.5854  0.0258
   1.000   0.7125   0.01069   0.00316  -0.1152  0.5633  0.0410
   1.500   0.7666   0.01018   0.00258  -0.1148  0.5422  0.0450
   2.000   0.8207   0.01008   0.00237  -0.1144  0.5213  0.0620
   2.500   0.8713   0.00851   0.00268  -0.1138  0.5006  1.0000
   3.000   0.9245   0.00882   0.00283  -0.1133  0.4819  1.0000
   4.000   0.9889   0.01376   0.00503  -0.1064  0.0072  1.0000
   4.500   1.0401   0.01425   0.00562  -0.1056  0.0053  1.0000
   5.000   1.0905   0.01481   0.00627  -0.1046  0.0062  1.0000
   5.500   1.1403   0.01542   0.00692  -0.1034  0.0081  1.0000
   6.000   1.1891   0.01612   0.00778  -0.1019  0.0103  1.0000
   6.500   1.2390   0.01666   0.00839  -0.1006  0.0148  1.0000
   7.000   1.2880   0.01731   0.00926  -0.0989  0.0202  1.0000
   7.500   1.3342   0.01821   0.01034  -0.0965  0.0234  1.0000
   8.000   1.3681   0.02010   0.01253  -0.0924  0.0229  1.0000
   8.500   1.3846   0.02284   0.01557  -0.0861  0.0218  1.0000
   9.000   1.3870   0.02602   0.01895  -0.0784  0.0197  1.0000
   9.500   1.3779   0.03063   0.02372  -0.0715  0.0168  1.0000