XFOIL         Version 6.94
  
 Calculated polar for: AG03 (flat aft bottom)                          
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2252   0.00608   0.00183  -0.0411  0.9118  1.0000
   0.500   0.2757   0.00618   0.00168  -0.0389  0.8549  1.0000
   1.000   0.3257   0.00639   0.00159  -0.0368  0.7925  1.0000
   1.500   0.3767   0.00671   0.00159  -0.0352  0.7265  1.0000
   2.000   0.4286   0.00710   0.00167  -0.0340  0.6589  1.0000
   2.500   0.4808   0.00755   0.00182  -0.0329  0.5867  1.0000
   3.000   0.5332   0.00806   0.00204  -0.0321  0.5115  1.0000
   3.500   0.5855   0.00867   0.00237  -0.0314  0.4341  1.0000
   4.000   0.6378   0.00935   0.00277  -0.0308  0.3602  1.0000
   5.000   0.7422   0.01086   0.00390  -0.0298  0.2342  1.0000
   5.500   0.7939   0.01174   0.00466  -0.0293  0.1795  1.0000
   6.000   0.8449   0.01279   0.00557  -0.0288  0.1292  1.0000
   6.500   0.8948   0.01406   0.00677  -0.0280  0.0912  1.0000
   7.000   0.9435   0.01551   0.00823  -0.0271  0.0672  1.0000
   9.000   1.1200   0.02432   0.01781  -0.0215  0.0204  1.0000
  10.000   1.1828   0.03268   0.02721  -0.0166  0.0133  1.0000
  10.500   1.2033   0.03732   0.03240  -0.0139  0.0115  1.0000
  11.000   1.2022   0.04377   0.03934  -0.0107  0.0107  1.0000