XFOIL         Version 6.94
  
 Calculated polar for: GOE 417 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.6404   0.01179   0.00525  -0.1258  0.8015  0.0262
   1.000   0.7296   0.00504  -0.00166  -0.1205  0.7072  0.0725
   1.500   0.7797   0.00466  -0.00243  -0.1186  0.6562  0.0595
   2.000   0.8273   0.00466  -0.00278  -0.1165  0.6020  0.0536
   2.500   0.8736   0.00464  -0.00301  -0.1146  0.5436  0.0518
   3.000   0.9191   0.00481  -0.00315  -0.1126  0.4818  0.0517
   3.500   0.9653   0.00507  -0.00315  -0.1108  0.4355  0.0594
   4.000   1.0146   0.00461  -0.00225  -0.1101  0.4035  1.0000
   4.500   1.0617   0.00499  -0.00203  -0.1086  0.3796  1.0000
   5.000   1.1024   0.00551  -0.00189  -0.1063  0.3083  1.0000
   5.500   1.1246   0.00758  -0.00135  -0.1018  0.0978  1.0000
   6.000   1.1574   0.00925  -0.00012  -0.0983  0.0049  1.0000
   6.500   1.2008   0.00994   0.00079  -0.0963  0.0046  1.0000
   7.000   1.2426   0.01073   0.00177  -0.0939  0.0049  1.0000
   7.500   1.2808   0.01181   0.00317  -0.0910  0.0054  1.0000
   8.000   1.3127   0.01326   0.00500  -0.0870  0.0061  1.0000
   8.500   1.3283   0.01546   0.00760  -0.0807  0.0067  1.0000
   9.000   1.3405   0.01775   0.01015  -0.0746  0.0077  1.0000
   9.500   1.3419   0.02091   0.01363  -0.0678  0.0090  1.0000
  10.000   1.3298   0.02609   0.01913  -0.0611  0.0107  1.0000
  10.500   1.3192   0.03304   0.02633  -0.0557  0.0120  1.0000