XFOIL         Version 6.94
  
 Calculated polar for: GOE 529 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   0.5744   0.00998   0.00263  -0.1088  0.6124  0.0813
   0.500   0.6281   0.00989   0.00262  -0.1083  0.6008  0.1298
   1.000   0.6796   0.00945   0.00273  -0.1077  0.5872  0.3215
   2.000   0.8033   0.00825   0.00279  -0.1104  0.5553  1.0000
   2.500   0.8544   0.00841   0.00285  -0.1094  0.5393  1.0000
   3.000   0.9049   0.00859   0.00292  -0.1083  0.5200  1.0000
   3.500   0.9517   0.00887   0.00296  -0.1065  0.4809  1.0000
   4.000   0.9882   0.00982   0.00335  -0.1030  0.3864  1.0000
   4.500   1.0265   0.01086   0.00391  -0.1001  0.3037  1.0000
   5.000   1.0595   0.01229   0.00478  -0.0965  0.2113  1.0000
   5.500   1.0822   0.01433   0.00616  -0.0912  0.0972  1.0000
   6.000   1.1052   0.01611   0.00758  -0.0859  0.0051  1.0000
   6.500   1.1423   0.01682   0.00836  -0.0827  0.0053  1.0000
   7.000   1.1743   0.01759   0.00924  -0.0785  0.0059  1.0000
   7.500   1.2031   0.01855   0.01032  -0.0741  0.0066  1.0000
   8.000   1.2311   0.01962   0.01152  -0.0698  0.0076  1.0000
   8.500   1.2533   0.02109   0.01317  -0.0650  0.0087  1.0000
   9.000   1.2747   0.02277   0.01502  -0.0605  0.0100  1.0000
   9.500   1.2801   0.02566   0.01812  -0.0551  0.0107  1.0000
  10.000   1.2911   0.02857   0.02122  -0.0511  0.0115  1.0000
  10.500   1.2866   0.03315   0.02600  -0.0471  0.0126  1.0000
  11.000   1.2864   0.03790   0.03094  -0.0444  0.0136  1.0000
  11.500   1.2777   0.04383   0.03706  -0.0418  0.0152  1.0000
  12.000   1.2843   0.04829   0.04163  -0.0387  0.0181  1.0000