XFOIL         Version 6.94
  
 Calculated polar for: GOE 477 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   0.5355   0.01127   0.00455  -0.1101  0.7781  0.2699
   0.500   0.6273   0.00928   0.00428  -0.1177  0.7669  1.0000
   1.000   0.6770   0.00919   0.00406  -0.1161  0.7537  1.0000
   1.500   0.7270   0.00904   0.00381  -0.1145  0.7384  1.0000
   2.000   0.7756   0.00890   0.00362  -0.1127  0.7183  1.0000
   3.500   0.8494   0.01216   0.00426  -0.0941  0.2881  1.0000
   4.000   0.8754   0.01378   0.00509  -0.0894  0.1739  1.0000
   4.500   0.8960   0.01594   0.00639  -0.0837  0.0095  1.0000
   5.000   0.9386   0.01651   0.00710  -0.0814  0.0090  1.0000
   5.500   0.9792   0.01709   0.00781  -0.0786  0.0094  1.0000
   6.000   1.0178   0.01785   0.00868  -0.0755  0.0101  1.0000
   6.500   1.0552   0.01876   0.00982  -0.0723  0.0112  1.0000
   7.000   1.0910   0.01983   0.01112  -0.0689  0.0131  1.0000
   7.500   1.1254   0.02107   0.01260  -0.0654  0.0157  1.0000
   8.000   1.1488   0.02310   0.01494  -0.0607  0.0177  1.0000
   8.500   1.1705   0.02529   0.01747  -0.0559  0.0223  1.0000
   9.000   1.1716   0.02914   0.02154  -0.0497  0.0251  1.0000
   9.500   1.1283   0.02397   0.01699  -0.0398  0.0301  1.0000