XFOIL         Version 6.94
  
 Calculated polar for: GOE 210 (DAIMLER) 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
 ------- -------- --------- --------- -------- ------- -------
   1.500   0.6574   0.00788   0.00254  -0.0966  0.7258  1.0000
   2.000   0.7111   0.00804   0.00258  -0.0959  0.7017  1.0000
   2.500   0.7641   0.00819   0.00261  -0.0952  0.6695  1.0000
   3.000   0.8161   0.00841   0.00266  -0.0943  0.6268  1.0000
   3.500   0.8584   0.00938   0.00274  -0.0916  0.4611  1.0000
   4.000   0.8949   0.01159   0.00364  -0.0890  0.2402  1.0000
   4.500   0.9401   0.01295   0.00440  -0.0877  0.1539  1.0000
   5.000   0.9829   0.01468   0.00573  -0.0858  0.0489  1.0000
   5.500   1.0297   0.01588   0.00681  -0.0844  0.0084  1.0000
   6.000   1.0791   0.01671   0.00782  -0.0830  0.0092  1.0000
   6.500   1.1279   0.01764   0.00898  -0.0815  0.0111  1.0000
   7.000   1.1749   0.01879   0.01040  -0.0795  0.0128  1.0000
   7.500   1.2150   0.02075   0.01264  -0.0765  0.0134  1.0000
   8.000   1.2453   0.02350   0.01553  -0.0723  0.0138  1.0000
   8.500   1.2768   0.02611   0.01835  -0.0680  0.0149  1.0000
   9.000   1.3123   0.03057   0.02275  -0.0644  0.0163  1.0000
   9.500   1.3647   0.03558   0.02817  -0.0624  0.0192  1.0000
  10.500   1.3977   0.05002   0.04449  -0.0509  0.0228  1.0000
  11.000   1.3813   0.05683   0.05191  -0.0436  0.0227  1.0000
  11.500   1.3522   0.06344   0.05896  -0.0371  0.0226  1.0000
  12.000   1.3191   0.07133   0.06724  -0.0340  0.0225  1.0000
  12.500   1.2887   0.08074   0.07694  -0.0340  0.0223  1.0000