XFOIL         Version 6.94
  
 Calculated polar for: GOE 414 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
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.7038   0.00837   0.00297  -0.1238  0.5771  1.0000
   1.500   0.7498   0.00864   0.00306  -0.1220  0.5496  1.0000
   2.000   0.7950   0.00897   0.00318  -0.1200  0.5216  1.0000
   2.500   0.8408   0.00935   0.00337  -0.1183  0.4986  1.0000
   3.000   0.8774   0.01006   0.00370  -0.1148  0.4419  1.0000
   3.500   0.9181   0.01072   0.00408  -0.1122  0.4043  1.0000
   4.000   0.9620   0.01125   0.00442  -0.1103  0.3682  1.0000
   4.500   0.9933   0.01243   0.00504  -0.1062  0.2720  1.0000
   5.500   1.0153   0.01700   0.00835  -0.0919  0.0057  1.0000
   6.000   1.0540   0.01783   0.00925  -0.0894  0.0058  1.0000
   6.500   1.0913   0.01878   0.01029  -0.0868  0.0064  1.0000
   7.000   1.1260   0.01992   0.01155  -0.0840  0.0072  1.0000
   7.500   1.1573   0.02132   0.01309  -0.0808  0.0081  1.0000
   8.000   1.1834   0.02313   0.01509  -0.0772  0.0094  1.0000
   8.500   1.2051   0.02538   0.01752  -0.0733  0.0110  1.0000
   9.000   1.2145   0.02870   0.02103  -0.0687  0.0125  1.0000
   9.500   1.2116   0.03332   0.02588  -0.0639  0.0143  1.0000
  10.000   1.2096   0.03828   0.03106  -0.0598  0.0165  1.0000