XFOIL         Version 6.94
  
 Calculated polar for: GOE 571 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   0.5089   0.02386   0.01506  -0.0588  0.4524  0.0499
   0.500   0.5599   0.02305   0.01394  -0.0579  0.4480  0.0488
   1.000   0.6134   0.02289   0.01366  -0.0580  0.4435  0.0491
   1.500   0.6315   0.02310   0.01393  -0.0520  0.4404  0.0498
   2.000   0.6535   0.02346   0.01431  -0.0470  0.4365  0.0519
   2.500   0.6846   0.02369   0.01461  -0.0439  0.4323  0.0538
   3.000   0.7257   0.02398   0.01485  -0.0424  0.4285  0.0567
   3.500   0.7794   0.02433   0.01516  -0.0432  0.4249  0.0627
   4.000   0.8275   0.02524   0.01623  -0.0437  0.4210  0.1016
   6.000   1.2131   0.02974   0.02273  -0.0903  0.4050  1.0000
   6.500   1.2588   0.03080   0.02364  -0.0903  0.4009  1.0000
   7.000   1.2408   0.03416   0.02719  -0.0830  0.3977  1.0000
   7.500   1.1783   0.04053   0.03387  -0.0729  0.3924  1.0000
   8.000   1.1483   0.04622   0.03971  -0.0674  0.3867  1.0000
   8.500   1.1907   0.04690   0.04028  -0.0673  0.3836  1.0000
   9.000   1.2542   0.04613   0.03933  -0.0687  0.3809  1.0000