XFOIL         Version 6.94
  
 Calculated polar for: GOE 451 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   0.4158   0.00997   0.00236  -0.0780  0.7194  0.0972
   1.000   0.4690   0.00896   0.00257  -0.0779  0.7033  0.4541
   1.500   0.5145   0.00780   0.00227  -0.0743  0.6253  1.0000
   2.000   0.5735   0.00802   0.00259  -0.0747  0.6469  1.0000
   3.000   0.6808   0.00844   0.00290  -0.0731  0.5947  1.0000
   3.500   0.7328   0.00870   0.00299  -0.0719  0.5469  1.0000
   4.000   0.7855   0.00909   0.00326  -0.0709  0.5025  1.0000
   4.500   0.8348   0.00974   0.00364  -0.0694  0.4267  1.0000
   5.000   0.8842   0.01054   0.00415  -0.0682  0.3495  1.0000
   6.000   0.9588   0.01509   0.00685  -0.0633  0.0463  1.0000
   6.500   1.0089   0.01649   0.00850  -0.0624  0.0487  1.0000
   7.000   1.0518   0.01842   0.01046  -0.0606  0.0425  1.0000
   7.500   1.0916   0.02053   0.01264  -0.0581  0.0381  1.0000
   8.000   1.1325   0.02287   0.01502  -0.0561  0.0334  1.0000
   8.500   1.1736   0.02532   0.01761  -0.0538  0.0300  1.0000
   9.000   1.2164   0.02823   0.02064  -0.0520  0.0278  1.0000
   9.500   1.2582   0.03138   0.02412  -0.0500  0.0261  1.0000
  10.000   1.3000   0.03546   0.02837  -0.0487  0.0245  1.0000
  10.500   1.3242   0.04008   0.03378  -0.0448  0.0231  1.0000
  11.000   1.3533   0.04364   0.03747  -0.0426  0.0214  1.0000
  11.500   1.3568   0.05091   0.04536  -0.0382  0.0210  1.0000
  12.000   1.3359   0.05747   0.05254  -0.0322  0.0209  1.0000
  12.500   1.2827   0.06407   0.05985  -0.0286  0.0196  1.0000