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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.6542   0.01277   0.00528  -0.1194  0.6618  0.0693
   1.000   0.7036   0.01274   0.00458  -0.1174  0.5639  0.0577
   1.500   0.7561   0.01271   0.00426  -0.1166  0.5162  0.0555
   2.000   0.8103   0.01272   0.00415  -0.1164  0.4847  0.0595
   2.500   0.8650   0.01277   0.00411  -0.1163  0.4578  0.0643
   3.000   0.9098   0.01120   0.00434  -0.1142  0.4344  1.0000
   4.500   1.0601   0.01316   0.00515  -0.1116  0.2705  1.0000
   5.000   1.1056   0.01443   0.00584  -0.1102  0.1885  1.0000
   5.500   1.1524   0.01552   0.00664  -0.1090  0.1481  1.0000
   6.000   1.1930   0.01731   0.00790  -0.1069  0.0666  1.0000
   6.500   1.2333   0.01905   0.00944  -0.1044  0.0060  1.0000
   7.000   1.2793   0.01998   0.01051  -0.1028  0.0060  1.0000
   7.500   1.3229   0.02108   0.01181  -0.1007  0.0062  1.0000
   8.000   1.3626   0.02244   0.01341  -0.0980  0.0066  1.0000
   8.500   1.3963   0.02416   0.01541  -0.0943  0.0070  1.0000
   9.000   1.4262   0.02588   0.01739  -0.0901  0.0077  1.0000
   9.500   1.4376   0.02845   0.02024  -0.0831  0.0083  1.0000
  10.500   1.4448   0.03533   0.02763  -0.0701  0.0098  1.0000
  11.000   1.4335   0.04071   0.03322  -0.0643  0.0106  1.0000
  11.500   1.4345   0.04570   0.03843  -0.0605  0.0118  1.0000