XFOIL         Version 6.94
  
 Calculated polar for: GOE 123 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   0.6312   0.01256   0.00597  -0.1179  0.8032  0.0517
   1.000   0.6863   0.01127   0.00438  -0.1169  0.7858  0.0502
   2.000   0.7944   0.01017   0.00330  -0.1154  0.7403  0.0597
   2.500   0.8476   0.00988   0.00302  -0.1145  0.7047  0.0882
   3.000   0.8975   0.00879   0.00318  -0.1136  0.6653  0.7511
   3.500   0.9473   0.00873   0.00316  -0.1118  0.6169  1.0000
   4.000   0.9870   0.01012   0.00356  -0.1087  0.4517  1.0000
   4.500   1.0288   0.01163   0.00423  -0.1068  0.3199  1.0000
   5.000   1.0721   0.01309   0.00503  -0.1052  0.2101  1.0000
   5.500   1.1092   0.01536   0.00641  -0.1029  0.0715  1.0000
   6.000   1.1532   0.01673   0.00756  -0.1012  0.0238  1.0000
   6.500   1.1990   0.01781   0.00864  -0.0997  0.0063  1.0000
   7.000   1.2447   0.01882   0.00981  -0.0980  0.0062  1.0000
   7.500   1.2880   0.02003   0.01124  -0.0960  0.0067  1.0000
   8.000   1.3264   0.02168   0.01314  -0.0932  0.0072  1.0000
   8.500   1.3617   0.02343   0.01517  -0.0899  0.0081  1.0000
   9.000   1.3791   0.02625   0.01824  -0.0841  0.0089  1.0000
   9.500   1.3895   0.02939   0.02163  -0.0777  0.0100  1.0000
  10.000   1.3946   0.03415   0.02654  -0.0712  0.0113  1.0000
  10.500   1.3540   0.03190   0.02486  -0.0571  0.0140  1.0000