XFOIL         Version 6.94
  
 Calculated polar for: GOE 278 (DAIMLER IX) 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   0.4780   0.00941   0.00287  -0.0901  0.7751  0.0291
   0.500   0.5300   0.00932   0.00256  -0.0890  0.7154  0.0343
   1.000   0.5801   0.00947   0.00242  -0.0878  0.6463  0.0441
   1.500   0.6310   0.00955   0.00225  -0.0867  0.5901  0.0474
   2.000   0.6819   0.00970   0.00218  -0.0858  0.5219  0.0517
   2.500   0.7318   0.01009   0.00227  -0.0847  0.4436  0.0764
   4.500   0.9258   0.01143   0.00384  -0.0810  0.1600  1.0000
   5.000   0.9668   0.01334   0.00508  -0.0790  0.0347  1.0000
   5.500   1.0165   0.01404   0.00571  -0.0782  0.0245  1.0000
   6.000   1.0667   0.01464   0.00634  -0.0774  0.0193  1.0000
   6.500   1.1142   0.01554   0.00719  -0.0762  0.0038  1.0000
   7.000   1.1623   0.01633   0.00811  -0.0750  0.0038  1.0000
   7.500   1.2088   0.01725   0.00914  -0.0736  0.0041  1.0000
   8.000   1.2532   0.01832   0.01041  -0.0718  0.0046  1.0000
   8.500   1.2970   0.01937   0.01162  -0.0699  0.0053  1.0000
   9.000   1.3368   0.02072   0.01318  -0.0674  0.0060  1.0000
   9.500   1.3748   0.02210   0.01478  -0.0646  0.0070  1.0000
  10.000   1.3983   0.02436   0.01730  -0.0598  0.0073  1.0000
  10.500   1.4034   0.02741   0.02058  -0.0527  0.0076  1.0000
  11.000   1.4132   0.03040   0.02383  -0.0473  0.0080  1.0000
  11.500   1.4115   0.03483   0.02861  -0.0418  0.0086  1.0000
  12.000   1.4045   0.04062   0.03472  -0.0369  0.0093  1.0000
  12.500   1.3983   0.05209   0.04713  -0.0294  0.0127  1.0000