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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.6753   0.00900   0.00265  -0.1583  0.6793  0.3536
   0.500   0.7302   0.00925   0.00275  -0.1577  0.6458  0.3685
   1.000   0.7841   0.00951   0.00288  -0.1569  0.6039  0.3810
   1.500   0.8366   0.00990   0.00307  -0.1560  0.5553  0.3939
   2.000   0.8856   0.01061   0.00338  -0.1545  0.4738  0.4058
   2.500   0.9354   0.01127   0.00368  -0.1533  0.4061  0.4145
   3.000   0.9837   0.01212   0.00408  -0.1520  0.3283  0.4233
   3.500   1.0344   0.01273   0.00446  -0.1511  0.2787  0.4336
   4.000   1.0776   0.01421   0.00533  -0.1491  0.1534  0.4437
   4.500   1.1143   0.01646   0.00683  -0.1460  0.0086  0.4544
   5.000   1.1654   0.01689   0.00739  -0.1450  0.0065  0.4677
   5.500   1.2147   0.01749   0.00815  -0.1436  0.0063  0.4831
   6.000   1.2619   0.01827   0.00921  -0.1418  0.0049  0.5014
   6.500   1.3066   0.01923   0.01048  -0.1395  0.0050  0.5259
   7.000   1.3465   0.02050   0.01223  -0.1365  0.0055  0.5784
   7.500   1.3813   0.02119   0.01363  -0.1324  0.0065  1.0000
   8.000   1.4193   0.02244   0.01502  -0.1288  0.0087  1.0000
   8.500   1.4362   0.02467   0.01747  -0.1219  0.0105  1.0000
   9.000   1.4645   0.02620   0.01914  -0.1171  0.0131  1.0000
   9.500   1.4762   0.02905   0.02227  -0.1103  0.0158  1.0000
  10.000   1.4741   0.03339   0.02685  -0.1033  0.0165  1.0000
  10.500   1.4774   0.03792   0.03165  -0.0977  0.0164  1.0000
  11.000   1.4860   0.04236   0.03632  -0.0932  0.0156  1.0000
  11.500   1.4925   0.04722   0.04142  -0.0893  0.0140  1.0000
  12.000   1.4920   0.05309   0.04744  -0.0859  0.0119  1.0000