XFOIL         Version 6.94
  
 Calculated polar for: GOE 358 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   0.7276   0.01082   0.00408  -0.1574  0.6908  0.3957
   1.000   0.7813   0.01106   0.00417  -0.1565  0.6548  0.4107
   1.500   0.8341   0.01129   0.00424  -0.1556  0.6118  0.4220
   2.000   0.8861   0.01157   0.00428  -0.1546  0.5648  0.4309
   2.500   0.9325   0.01226   0.00447  -0.1526  0.4810  0.4405
   3.000   0.9777   0.01318   0.00489  -0.1507  0.3859  0.4500
   3.500   1.0256   0.01392   0.00533  -0.1493  0.3209  0.4605
   4.000   1.0700   0.01512   0.00593  -0.1475  0.2043  0.4727
   4.500   1.1042   0.01759   0.00755  -0.1442  0.0489  0.4853
   5.000   1.1498   0.01860   0.00853  -0.1424  0.0087  0.5000
   5.500   1.1980   0.01923   0.00937  -0.1408  0.0091  0.5190
   6.000   1.2452   0.01993   0.01032  -0.1389  0.0102  0.5457
   6.500   1.2908   0.02064   0.01144  -0.1368  0.0116  0.6007
   8.500   1.4104   0.02681   0.01878  -0.1181  0.0150  1.0000
   9.000   1.4175   0.02991   0.02201  -0.1112  0.0155  1.0000
   9.500   1.4311   0.03287   0.02518  -0.1058  0.0166  1.0000
  10.000   1.4342   0.03713   0.02954  -0.1000  0.0176  1.0000
  10.500   1.4510   0.04069   0.03332  -0.0955  0.0193  1.0000
  11.000   1.4828   0.04421   0.03706  -0.0908  0.0222  1.0000
  12.000   1.5873   0.05562   0.04921  -0.0858  0.0243  1.0000
  12.500   1.5906   0.06336   0.05741  -0.0814  0.0237  1.0000
  13.000   1.5635   0.07048   0.06506  -0.0757  0.0237  1.0000
  13.500   1.5359   0.07888   0.07391  -0.0723  0.0237  1.0000
  14.000   1.5049   0.08825   0.08369  -0.0710  0.0238  1.0000