XFOIL         Version 6.94
  
 Calculated polar for: AH 80-140                                       
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.3922   0.01260   0.00465  -0.0463  0.5121  0.0785
   1.000   0.4462   0.01001   0.00468  -0.0462  0.5046  0.8949
   1.500   0.4877   0.01257   0.00482  -0.0424  0.5034  0.1944
   2.000   0.5298   0.01212   0.00485  -0.0398  0.4971  0.3591
   2.500   0.5902   0.01345   0.00541  -0.0402  0.4969  0.0749
   3.000   0.7382   0.01087   0.00544  -0.0604  0.4897  0.9495
   3.500   0.7879   0.00982   0.00399  -0.0581  0.4678  0.9429
   4.000   0.7340   0.01182   0.00458  -0.0345  0.4762  0.4041
   4.500   0.7873   0.01172   0.00376  -0.0329  0.4563  0.1366
   5.500   0.9864   0.01054   0.00485  -0.0530  0.4520  0.9371
   6.000   1.1130   0.01082   0.00532  -0.0682  0.4444  1.0000
   6.500   1.1539   0.01045   0.00483  -0.0646  0.4145  1.0000
   7.000   1.1563   0.01067   0.00509  -0.0536  0.4160  0.9520
   7.500   1.2406   0.01108   0.00565  -0.0598  0.4077  1.0000
   8.000   1.2820   0.01143   0.00610  -0.0573  0.4006  1.0000
   8.500   1.0686   0.02382   0.01568  -0.0240  0.0044  1.0000
   9.000   0.9715   0.02762   0.01754   0.0012  0.0187  0.0705
   9.500   1.3122   0.01581   0.00935  -0.0347  0.2098  1.0000
  10.000   1.3251   0.01726   0.01072  -0.0288  0.1934  1.0000
  10.500   1.2981   0.02229   0.01561  -0.0233  0.1453  1.0000
  11.000   1.1046   0.03613   0.02706  -0.0010  0.0552  0.0687
  11.500   1.1754   0.03463   0.02618  -0.0019  0.0999  0.0738
  12.000   1.1792   0.04786   0.04046  -0.0194  0.0096  1.0000
  12.500   1.2286   0.03982   0.03188  -0.0015  0.1130  0.1017
  13.000   1.1614   0.05071   0.04221  -0.0013  0.0473  0.0829