XFOIL         Version 6.94
  
 Calculated polar for: GOE 269 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   0.5112   0.01075   0.00380  -0.0987  0.6769  0.0540
   0.500   0.5679   0.01061   0.00356  -0.0985  0.6576  0.0635
   1.000   0.6245   0.01044   0.00334  -0.0985  0.6372  0.0754
   1.500   0.6811   0.01020   0.00307  -0.0984  0.6143  0.0852
   2.000   0.7373   0.01003   0.00287  -0.0983  0.5838  0.0935
   2.500   0.7933   0.00998   0.00276  -0.0982  0.5446  0.1022
   3.000   0.8456   0.01049   0.00278  -0.0976  0.4413  0.1100
   3.500   0.8995   0.01092   0.00303  -0.0973  0.3898  0.1272
   4.000   0.9458   0.01016   0.00347  -0.0960  0.3033  1.0000
   4.500   0.9971   0.01115   0.00394  -0.0955  0.2174  1.0000
   5.000   1.0469   0.01235   0.00469  -0.0948  0.1401  1.0000
   5.500   1.0889   0.01480   0.00621  -0.0931  0.0062  1.0000
   6.000   1.1413   0.01534   0.00686  -0.0925  0.0046  1.0000
   6.500   1.1923   0.01606   0.00764  -0.0917  0.0045  1.0000
   7.000   1.2421   0.01690   0.00861  -0.0906  0.0045  1.0000
   7.500   1.2905   0.01786   0.00976  -0.0894  0.0047  1.0000
   8.000   1.3364   0.01904   0.01115  -0.0877  0.0049  1.0000
   8.500   1.3791   0.02049   0.01289  -0.0857  0.0053  1.0000
   9.000   1.4177   0.02223   0.01487  -0.0830  0.0057  1.0000
   9.500   1.4501   0.02433   0.01722  -0.0795  0.0064  1.0000
  10.000   1.4706   0.02703   0.02022  -0.0745  0.0070  1.0000
  10.500   1.4729   0.03041   0.02389  -0.0674  0.0076  1.0000
  11.000   1.4691   0.03507   0.02886  -0.0617  0.0080  1.0000