XFOIL         Version 6.94
  
 Calculated polar for: GOE 405 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   0.7399   0.01169   0.00433  -0.1412  0.6897  0.2480
   1.000   0.7906   0.01157   0.00423  -0.1402  0.6750  0.2713
   1.500   0.8401   0.01145   0.00413  -0.1390  0.6589  0.3040
   2.000   0.9188   0.01008   0.00428  -0.1445  0.6402  1.0000
   2.500   0.9629   0.01029   0.00431  -0.1420  0.6201  1.0000
   3.000   1.0057   0.01053   0.00439  -0.1394  0.5976  1.0000
   3.500   1.0472   0.01083   0.00454  -0.1365  0.5742  1.0000
   4.000   1.0532   0.01147   0.00466  -0.1262  0.4867  1.0000
   4.500   1.0768   0.01216   0.00510  -0.1200  0.4464  1.0000
   5.000   1.0907   0.01314   0.00573  -0.1122  0.3942  1.0000
   5.500   1.1161   0.01400   0.00640  -0.1069  0.3545  1.0000
   6.000   1.1342   0.01524   0.00734  -0.1007  0.2989  1.0000
   6.500   1.1394   0.01730   0.00889  -0.0930  0.2179  1.0000
   7.500   1.1299   0.02361   0.01417  -0.0773  0.0067  1.0000
   8.000   1.1577   0.02514   0.01584  -0.0741  0.0066  1.0000
   9.000   1.2059   0.02897   0.02003  -0.0676  0.0074  1.0000
   9.500   1.2254   0.03142   0.02269  -0.0644  0.0081  1.0000
  10.000   1.2363   0.03472   0.02624  -0.0608  0.0087  1.0000
  10.500   1.2482   0.03817   0.02992  -0.0578  0.0096  1.0000
  11.000   1.2453   0.04327   0.03528  -0.0546  0.0105  1.0000
  11.500   1.2329   0.04972   0.04193  -0.0520  0.0109  1.0000
  12.000   1.2327   0.05538   0.04786  -0.0501  0.0124  1.0000
  12.500   1.2186   0.06268   0.05530  -0.0484  0.0133  1.0000
  13.000   1.2259   0.06725   0.05997  -0.0455  0.0159  1.0000