XFOIL         Version 6.94
  
 Calculated polar for: AH21 9% version (Andrew Hollom)                 
  
 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   0.1526   0.00934   0.00488  -0.0464  0.9624  1.0000
   0.500   0.2381   0.00935   0.00485  -0.0528  0.9542  1.0000
   1.000   0.3083   0.00906   0.00460  -0.0555  0.9379  1.0000
   1.500   0.4133   0.00832   0.00393  -0.0650  0.9280  1.0000
   2.000   0.4780   0.00767   0.00336  -0.0658  0.9023  1.0000
   2.500   0.5448   0.00711   0.00284  -0.0669  0.8531  1.0000
   3.000   0.6121   0.00716   0.00251  -0.0681  0.7173  1.0000
   3.500   0.6344   0.00907   0.00292  -0.0609  0.4082  1.0000
   4.000   0.6670   0.01104   0.00374  -0.0571  0.1703  1.0000
   4.500   0.7082   0.01276   0.00524  -0.0545  0.0685  1.0000
   5.000   0.7558   0.01360   0.00583  -0.0535  0.0343  1.0000
   5.500   0.8021   0.01481   0.00704  -0.0518  0.0198  1.0000
   6.000   0.8480   0.01616   0.00847  -0.0499  0.0125  1.0000
   7.000   0.9316   0.02155   0.01444  -0.0448  0.0083  1.0000
   7.500   0.9711   0.02598   0.01948  -0.0417  0.0076  1.0000
   8.000   0.9955   0.03322   0.02769  -0.0366  0.0075  1.0000
   8.500   0.9965   0.04276   0.03822  -0.0297  0.0077  1.0000
   9.000   0.9754   0.05307   0.04928  -0.0228  0.0080  1.0000
   9.500   0.9388   0.06168   0.05831  -0.0166  0.0081  1.0000
  10.000   0.8957   0.07227   0.06920  -0.0175  0.0082  1.0000
  10.500   0.8571   0.09310   0.09019  -0.0333  0.0085  1.0000
  11.000   0.8358   0.10809   0.10513  -0.0420  0.0087  1.0000
  11.500   0.8192   0.12198   0.11897  -0.0490  0.0090  1.0000