XFOIL         Version 6.94
  
 Calculated polar for: GOE 240 (KOLLER) 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   0.5407   0.00955   0.00271  -0.1055  0.7358  0.0340
   0.500   0.5954   0.00918   0.00231  -0.1052  0.7145  0.0439
   1.000   0.6500   0.00894   0.00206  -0.1047  0.6859  0.0505
   1.500   0.7045   0.00882   0.00202  -0.1042  0.6516  0.0942
   3.500   0.9012   0.00942   0.00282  -0.0999  0.3518  1.0000
   4.000   0.9391   0.01162   0.00389  -0.0976  0.1529  1.0000
   4.500   0.9823   0.01320   0.00487  -0.0959  0.0513  1.0000
   5.000   1.0299   0.01422   0.00573  -0.0945  0.0062  1.0000
   5.500   1.0802   0.01483   0.00642  -0.0935  0.0066  1.0000
   6.000   1.1297   0.01554   0.00726  -0.0923  0.0078  1.0000
   6.500   1.1782   0.01634   0.00819  -0.0908  0.0096  1.0000
   7.000   1.2237   0.01745   0.00950  -0.0889  0.0102  1.0000
   7.500   1.2646   0.01899   0.01123  -0.0861  0.0110  1.0000
   8.000   1.2944   0.02135   0.01374  -0.0818  0.0118  1.0000
   8.500   1.3248   0.02335   0.01593  -0.0775  0.0130  1.0000
   9.000   1.3345   0.02677   0.01940  -0.0703  0.0144  1.0000
   9.500   1.3595   0.02943   0.02222  -0.0655  0.0148  1.0000
  10.000   1.3900   0.03336   0.02634  -0.0619  0.0139  1.0000
  10.500   1.4122   0.03746   0.03070  -0.0580  0.0112  1.0000
  11.000   1.3287   0.03315   0.02702  -0.0420  0.0112  1.0000
  11.500   1.4132   0.04859   0.04246  -0.0488  0.0056  1.0000
  12.000   1.3887   0.05371   0.04813  -0.0446  0.0049  1.0000