XFOIL         Version 6.94
  
 Calculated polar for: BOEING-VERTOL VR-8 AIRFOIL WITH                 
  
 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
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.1691   0.00751   0.00176  -0.0038  0.5766  1.0000
   1.500   0.2288   0.00771   0.00176  -0.0049  0.5357  1.0000
   2.000   0.2885   0.00791   0.00182  -0.0061  0.5081  1.0000
   2.500   0.3482   0.00820   0.00194  -0.0072  0.4586  1.0000
   3.000   0.4080   0.00919   0.00221  -0.0091  0.2860  1.0000
   3.500   0.4676   0.01123   0.00321  -0.0118  0.0458  1.0000
   4.000   0.5262   0.01192   0.00380  -0.0131  0.0266  1.0000
   4.500   0.5836   0.01367   0.00559  -0.0147  0.0207  1.0000
   5.500   0.6941   0.01654   0.00859  -0.0165  0.0144  1.0000
   6.000   0.7487   0.01771   0.00986  -0.0171  0.0129  1.0000
   7.000   0.8615   0.01879   0.01159  -0.0179  0.0048  1.0000
   7.500   0.9100   0.02190   0.01506  -0.0174  0.0048  1.0000
   8.000   0.9494   0.02970   0.02383  -0.0153  0.0053  1.0000
   8.500   0.9664   0.04254   0.03817  -0.0126  0.0063  1.0000
   9.000   0.9572   0.05503   0.05158  -0.0117  0.0073  1.0000
  10.000   0.9024   0.07640   0.07366  -0.0229  0.0076  1.0000
  10.500   0.8689   0.09962   0.09707  -0.0407  0.0074  1.0000
  11.000   0.8547   0.11495   0.11235  -0.0493  0.0076  1.0000