XFOIL         Version 6.94
  
 Calculated polar for: GOE 474 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3338   0.00787   0.00181  -0.0715  0.8354  0.2298
   1.000   0.5018   0.00594   0.00179  -0.0835  0.7858  1.0000
   1.500   0.5481   0.00605   0.00175  -0.0812  0.7560  1.0000
   2.000   0.5945   0.00619   0.00178  -0.0790  0.7242  1.0000
   2.500   0.6304   0.00672   0.00174  -0.0743  0.6047  1.0000
   3.000   0.6489   0.00880   0.00229  -0.0671  0.2931  1.0000
   3.500   0.6833   0.01026   0.00302  -0.0633  0.1297  1.0000
   4.000   0.7253   0.01114   0.00356  -0.0607  0.0663  1.0000
   4.500   0.7658   0.01224   0.00436  -0.0577  0.0067  1.0000
   5.000   0.8128   0.01276   0.00499  -0.0557  0.0073  1.0000
   5.500   0.8593   0.01339   0.00575  -0.0534  0.0086  1.0000
   6.000   0.9043   0.01426   0.00681  -0.0508  0.0095  1.0000
   6.500   0.9447   0.01566   0.00841  -0.0475  0.0095  1.0000
   7.000   0.9791   0.01763   0.01052  -0.0433  0.0096  1.0000
   7.500   1.0112   0.02016   0.01317  -0.0387  0.0097  1.0000
   8.000   1.0503   0.02391   0.01704  -0.0356  0.0100  1.0000
   9.500   1.1306   0.03978   0.03480  -0.0215  0.0122  1.0000
  10.000   1.1195   0.04343   0.03907  -0.0129  0.0107  1.0000
  10.500   1.0845   0.04926   0.04545  -0.0032  0.0109  1.0000
  11.000   1.0535   0.05534   0.05190   0.0016  0.0108  1.0000
  11.500   1.0175   0.06346   0.06036   0.0021  0.0107  1.0000
  12.000   0.9898   0.07266   0.06975  -0.0013  0.0105  1.0000
  12.500   0.9735   0.08243   0.07958  -0.0057  0.0102  1.0000