XFOIL         Version 6.94
  
 Calculated polar for: GOE 408 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   0.4032   0.00978   0.00264  -0.0759  0.7204  0.0504
   0.500   0.4488   0.00204  -0.00501  -0.0732  0.6825  0.0668
   1.000   0.4935   0.00196  -0.00514  -0.0708  0.6582  0.0710
   1.500   0.5367   0.00192  -0.00524  -0.0680  0.6302  0.0757
   2.000   0.5788   0.00189  -0.00527  -0.0650  0.5990  0.0853
   3.000   0.8362   0.00137  -0.00455  -0.1006  0.4468  1.0000
   3.500   0.8683   0.00162  -0.00481  -0.0960  0.3301  1.0000
   4.000   0.8935   0.00219  -0.00511  -0.0904  0.1633  1.0000
   4.500   0.9190   0.00294  -0.00499  -0.0849  0.0423  1.0000
   5.000   0.9548   0.00343  -0.00463  -0.0811  0.0059  1.0000
   5.500   0.9949   0.00377  -0.00418  -0.0780  0.0062  1.0000
   6.000   1.0340   0.00414  -0.00369  -0.0747  0.0073  1.0000
   6.500   1.0712   0.00461  -0.00305  -0.0710  0.0086  1.0000
   7.000   1.1020   0.00535  -0.00211  -0.0660  0.0089  1.0000
   7.500   1.1213   0.00643  -0.00084  -0.0589  0.0091  1.0000
   8.000   1.1284   0.00785   0.00077  -0.0499  0.0093  1.0000
   8.500   1.1276   0.00983   0.00289  -0.0403  0.0094  1.0000
   9.000   1.1293   0.01236   0.00556  -0.0320  0.0094  1.0000
   9.500   1.1445   0.01468   0.00816  -0.0258  0.0105  1.0000
  10.000   1.1777   0.02023   0.01381  -0.0225  0.0116  1.0000
  10.500   1.3096   0.03781   0.03143  -0.0318  0.0120  1.0000
  11.000   1.3035   0.04295   0.03716  -0.0235  0.0114  1.0000
  11.500   1.2816   0.04694   0.04171  -0.0145  0.0107  1.0000
  12.000   1.2623   0.05168   0.04669  -0.0094  0.0090  1.0000
  12.500   1.2336   0.05837   0.05386  -0.0060  0.0086  1.0000
  13.000   1.2039   0.06701   0.06291  -0.0054  0.0087  1.0000