XFOIL         Version 6.94
  
 Calculated polar for: GOE 282 (DAIMLER XIII) 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   0.5852   0.00938   0.00250  -0.1184  0.6591  0.1914
   0.500   0.6394   0.00951   0.00252  -0.1180  0.6170  0.2117
   1.000   0.6936   0.00965   0.00256  -0.1177  0.5815  0.2344
   1.500   0.7482   0.00969   0.00266  -0.1176  0.5546  0.2958
   3.500   0.9372   0.01140   0.00400  -0.1118  0.2341  1.0000
   4.000   0.9750   0.01357   0.00519  -0.1092  0.0710  1.0000
   4.500   1.0213   0.01475   0.00611  -0.1076  0.0073  1.0000
   5.000   1.0721   0.01533   0.00674  -0.1066  0.0083  1.0000
   5.500   1.1221   0.01597   0.00746  -0.1054  0.0102  1.0000
   6.000   1.1707   0.01671   0.00830  -0.1039  0.0117  1.0000
   6.500   1.2166   0.01767   0.00943  -0.1018  0.0124  1.0000
   7.000   1.2605   0.01873   0.01065  -0.0994  0.0135  1.0000
   7.500   1.3000   0.02008   0.01215  -0.0963  0.0146  1.0000
   8.000   1.3328   0.02173   0.01394  -0.0921  0.0152  1.0000
   8.500   1.3373   0.02466   0.01698  -0.0838  0.0144  1.0000
   9.000   1.3464   0.02747   0.01994  -0.0768  0.0132  1.0000
   9.500   1.3432   0.03189   0.02434  -0.0695  0.0117  1.0000
  10.000   1.3727   0.03620   0.02844  -0.0664  0.0081  1.0000
  10.500   1.3663   0.04021   0.03298  -0.0604  0.0052  1.0000
  11.000   1.3792   0.04458   0.03760  -0.0567  0.0037  1.0000
  11.500   1.4373   0.05073   0.04375  -0.0566  0.0027  1.0000
  12.000   1.3961   0.05469   0.04826  -0.0502  0.0025  1.0000
  12.500   1.3855   0.06050   0.05447  -0.0480  0.0021  1.0000
  13.000   1.3755   0.06748   0.06177  -0.0469  0.0018  1.0000
  13.500   1.3643   0.07525   0.06981  -0.0465  0.0016  1.0000