XFOIL         Version 6.94
  
 Calculated polar for: AG17                                            
  
 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   0.2515   0.00493   0.00126  -0.0540  0.8567  1.0000
   0.500   0.3040   0.00517   0.00119  -0.0526  0.7955  1.0000
   1.000   0.3574   0.00549   0.00118  -0.0515  0.7324  1.0000
   1.500   0.4114   0.00585   0.00124  -0.0507  0.6661  1.0000
   2.000   0.4656   0.00626   0.00136  -0.0501  0.5958  1.0000
   2.500   0.5198   0.00673   0.00153  -0.0496  0.5218  1.0000
   3.000   0.5741   0.00726   0.00180  -0.0492  0.4461  1.0000
   3.500   0.6281   0.00788   0.00213  -0.0488  0.3714  1.0000
   4.000   0.6820   0.00853   0.00255  -0.0485  0.3023  1.0000
   4.500   0.7357   0.00922   0.00304  -0.0482  0.2384  1.0000
   5.500   0.8418   0.01091   0.00436  -0.0474  0.1259  1.0000
   6.000   0.8939   0.01192   0.00523  -0.0470  0.0797  1.0000
   6.500   0.9451   0.01314   0.00638  -0.0463  0.0452  1.0000
   7.000   0.9950   0.01468   0.00794  -0.0454  0.0236  1.0000
   7.500   1.0430   0.01661   0.01001  -0.0442  0.0133  1.0000
   8.000   1.0849   0.01993   0.01373  -0.0421  0.0101  1.0000
   8.500   1.1258   0.02334   0.01755  -0.0400  0.0089  1.0000
   9.000   1.1645   0.02648   0.02099  -0.0383  0.0075  1.0000
   9.500   1.1804   0.03451   0.02987  -0.0348  0.0069  1.0000
  10.000   1.1952   0.04039   0.03645  -0.0320  0.0066  1.0000
  10.500   1.1821   0.04792   0.04462  -0.0287  0.0065  1.0000
  11.000   1.1483   0.05742   0.05457  -0.0307  0.0066  1.0000
  11.500   1.1126   0.07308   0.07060  -0.0434  0.0067  1.0000