XFOIL         Version 6.94
  
 Calculated polar for: davissm                                         
  
 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   0.7142   0.00991   0.00364  -0.1654  0.8356  0.0491
   0.500   0.7698   0.00958   0.00328  -0.1651  0.8203  0.0572
   1.000   0.8253   0.00921   0.00291  -0.1649  0.8032  0.0828
   2.000   0.9332   0.00768   0.00298  -0.1642  0.7580  1.0000
   2.500   0.9863   0.00789   0.00306  -0.1634  0.7277  1.0000
   3.000   1.0385   0.00816   0.00321  -0.1625  0.6911  1.0000
   3.500   1.0896   0.00855   0.00346  -0.1613  0.6481  1.0000
   4.000   1.1380   0.00914   0.00380  -0.1597  0.5848  1.0000
   4.500   1.1820   0.01012   0.00435  -0.1573  0.4888  1.0000
   5.000   1.2137   0.01241   0.00555  -0.1530  0.2846  1.0000
   5.500   1.2393   0.01562   0.00757  -0.1481  0.0697  1.0000
   6.000   1.2823   0.01683   0.00880  -0.1458  0.0493  1.0000
   6.500   1.3237   0.01807   0.01009  -0.1432  0.0344  1.0000
   7.000   1.3557   0.02002   0.01215  -0.1387  0.0262  1.0000
   7.500   1.3891   0.02159   0.01389  -0.1346  0.0229  1.0000
   8.000   1.4162   0.02349   0.01592  -0.1295  0.0197  1.0000
   8.500   1.4264   0.02724   0.01993  -0.1215  0.0177  1.0000
   9.000   1.4536   0.02978   0.02271  -0.1166  0.0168  1.0000
   9.500   1.4848   0.03293   0.02613  -0.1127  0.0154  1.0000
  10.000   1.5061   0.03520   0.02862  -0.1076  0.0137  1.0000
  10.500   1.5266   0.03903   0.03271  -0.1029  0.0127  1.0000
  11.000   1.5335   0.04925   0.04374  -0.0972  0.0119  1.0000
  11.500   1.5105   0.05676   0.05187  -0.0883  0.0118  1.0000
  12.000   1.4890   0.06266   0.05824  -0.0814  0.0115  1.0000
  12.500   1.4486   0.07243   0.06851  -0.0759  0.0116  1.0000