XFOIL         Version 6.94
  
 Calculated polar for: GOE 81 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
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.7536   0.00402  -0.00228  -0.1237  0.6141  0.0848
   1.500   0.8073   0.00364  -0.00272  -0.1238  0.5929  0.1211
   2.000   0.8664   0.00287  -0.00410  -0.1222  0.5698  0.0454
   2.500   0.9190   0.00266  -0.00438  -0.1216  0.5384  0.0449
   3.000   0.9698   0.00262  -0.00457  -0.1208  0.4925  0.0473
   3.500   1.0188   0.00273  -0.00464  -0.1198  0.4548  0.0561
   4.000   1.0611   0.00244  -0.00369  -0.1179  0.3902  1.0000
   4.500   1.1037   0.00285  -0.00383  -0.1160  0.3061  1.0000
   5.000   1.1338   0.00393  -0.00368  -0.1124  0.1560  1.0000
   5.500   1.1716   0.00478  -0.00318  -0.1098  0.1039  1.0000
   6.000   1.2051   0.00589  -0.00244  -0.1067  0.0375  1.0000
   6.500   1.2406   0.00684  -0.00158  -0.1037  0.0041  1.0000
   7.000   1.2800   0.00749  -0.00086  -0.1012  0.0042  1.0000
   7.500   1.3152   0.00826   0.00005  -0.0982  0.0044  1.0000
   8.000   1.3416   0.00918   0.00113  -0.0937  0.0047  1.0000
   8.500   1.3661   0.01022   0.00234  -0.0893  0.0052  1.0000
   9.000   1.3844   0.01169   0.00400  -0.0845  0.0058  1.0000
   9.500   1.3956   0.01383   0.00638  -0.0795  0.0063  1.0000
  10.000   1.4100   0.01617   0.00892  -0.0755  0.0070  1.0000
  10.500   1.4089   0.02026   0.01325  -0.0707  0.0078  1.0000
  11.000   1.4156   0.02426   0.01746  -0.0673  0.0087  1.0000
  11.500   1.4022   0.03114   0.02458  -0.0636  0.0096  1.0000
  12.000   1.3972   0.03781   0.03149  -0.0609  0.0106  1.0000
  12.500   1.3762   0.04716   0.04111  -0.0585  0.0115  1.0000
  13.000   1.3458   0.05826   0.05241  -0.0567  0.0120  1.0000