XFOIL         Version 6.94
  
 Calculated polar for: AG18                                            
  
 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   0.2648   0.00484   0.00123  -0.0559  0.8785  1.0000
   0.500   0.3165   0.00507   0.00115  -0.0542  0.8177  1.0000
   1.000   0.3694   0.00537   0.00114  -0.0530  0.7533  1.0000
   1.500   0.4228   0.00574   0.00119  -0.0520  0.6835  1.0000
   2.000   0.4764   0.00618   0.00130  -0.0512  0.6044  1.0000
   2.500   0.5300   0.00671   0.00148  -0.0506  0.5158  1.0000
   3.000   0.5836   0.00734   0.00176  -0.0501  0.4268  1.0000
   3.500   0.6373   0.00799   0.00211  -0.0497  0.3474  1.0000
   4.000   0.6910   0.00867   0.00253  -0.0494  0.2785  1.0000
   4.500   0.7446   0.00938   0.00305  -0.0490  0.2173  1.0000
   5.000   0.7977   0.01015   0.00365  -0.0486  0.1616  1.0000
   5.500   0.8503   0.01106   0.00441  -0.0482  0.1135  1.0000
   6.000   0.9024   0.01205   0.00529  -0.0476  0.0754  1.0000
   6.500   0.9538   0.01317   0.00642  -0.0470  0.0470  1.0000
   7.000   1.0036   0.01472   0.00799  -0.0460  0.0258  1.0000
   7.500   1.0521   0.01649   0.00992  -0.0449  0.0146  1.0000
   8.500   1.1363   0.02296   0.01718  -0.0408  0.0088  1.0000
   9.000   1.1707   0.02726   0.02188  -0.0388  0.0071  1.0000
   9.500   1.1956   0.03331   0.02869  -0.0360  0.0067  1.0000
  10.000   1.2088   0.03991   0.03605  -0.0330  0.0063  1.0000
  10.500   1.1939   0.04810   0.04493  -0.0299  0.0062  1.0000
  11.000   1.1587   0.05806   0.05535  -0.0326  0.0062  1.0000
  11.500   1.1238   0.07488   0.07254  -0.0470  0.0064  1.0000
  12.000   1.0808   0.09648   0.09434  -0.0621  0.0067  1.0000