XFOIL         Version 6.94
  
 Calculated polar for: GOE 274 (DAIMLER V) 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.6476   0.01282   0.00548  -0.1191  0.5421  0.1029
   1.000   0.7040   0.01231   0.00438  -0.1173  0.4989  0.0495
   1.500   0.7590   0.01191   0.00387  -0.1171  0.4685  0.0456
   2.000   0.8145   0.01170   0.00362  -0.1172  0.4418  0.0471
   2.500   0.8698   0.01170   0.00356  -0.1172  0.4182  0.0519
   3.000   0.9246   0.01182   0.00359  -0.1172  0.3970  0.0591
   4.000   1.0155   0.01173   0.00430  -0.1141  0.2300  1.0000
   4.500   1.0626   0.01296   0.00499  -0.1130  0.1585  1.0000
   5.000   1.1080   0.01435   0.00586  -0.1116  0.0816  1.0000
   5.500   1.1517   0.01593   0.00712  -0.1097  0.0049  1.0000
   6.000   1.2017   0.01661   0.00785  -0.1087  0.0048  1.0000
   6.500   1.2502   0.01739   0.00874  -0.1075  0.0050  1.0000
   7.000   1.2972   0.01828   0.00976  -0.1060  0.0055  1.0000
   7.500   1.3422   0.01929   0.01094  -0.1040  0.0064  1.0000
   8.000   1.3836   0.02058   0.01242  -0.1015  0.0072  1.0000
   8.500   1.4190   0.02228   0.01435  -0.0979  0.0086  1.0000
   9.000   1.4481   0.02412   0.01642  -0.0932  0.0102  1.0000
   9.500   1.4602   0.02655   0.01910  -0.0858  0.0121  1.0000
  10.000   1.4605   0.02990   0.02269  -0.0775  0.0144  1.0000
  10.500   1.4564   0.03434   0.02728  -0.0695  0.0181  1.0000