XFOIL         Version 6.94
  
 Calculated polar for: GIII BL207 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   0.0983   0.00606   0.00206  -0.0134  0.8281  0.9304
   0.500   0.1568   0.00618   0.00211  -0.0133  0.8098  0.9537
   1.000   0.2203   0.00627   0.00212  -0.0149  0.7913  0.9682
   1.500   0.2929   0.00632   0.00211  -0.0186  0.7718  0.9766
   2.000   0.3627   0.00638   0.00213  -0.0219  0.7496  0.9874
   3.000   0.4961   0.00655   0.00207  -0.0271  0.6411  1.0000
   3.500   0.5341   0.00789   0.00226  -0.0243  0.3802  1.0000
   4.000   0.5745   0.00915   0.00276  -0.0223  0.2110  1.0000
   4.500   0.6121   0.01077   0.00351  -0.0199  0.0515  1.0000
   5.000   0.6576   0.01146   0.00407  -0.0182  0.0268  1.0000
   5.500   0.7046   0.01206   0.00468  -0.0168  0.0186  1.0000
   6.000   0.7519   0.01276   0.00538  -0.0155  0.0109  1.0000
   6.500   0.7999   0.01347   0.00618  -0.0144  0.0078  1.0000
   7.000   0.8464   0.01448   0.00728  -0.0130  0.0032  1.0000
   7.500   0.8923   0.01563   0.00862  -0.0115  0.0030  1.0000
   8.000   0.9367   0.01707   0.01032  -0.0098  0.0029  1.0000
   8.500   0.9783   0.01898   0.01257  -0.0078  0.0030  1.0000
   9.000   1.0146   0.02171   0.01572  -0.0051  0.0031  1.0000
   9.500   1.0425   0.02588   0.02047  -0.0017  0.0032  1.0000
  10.000   1.0557   0.03197   0.02737   0.0027  0.0033  1.0000
  10.500   1.0556   0.03863   0.03483   0.0076  0.0036  1.0000