XFOIL         Version 6.94
  
 Calculated polar for: NACA 64-210                                     
  
 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   0.1694   0.00668   0.00215  -0.0398  0.7618  0.7624
   0.500   0.2241   0.00673   0.00225  -0.0397  0.7451  0.7824
   1.000   0.2800   0.00679   0.00230  -0.0398  0.7302  0.7979
   1.500   0.3325   0.00685   0.00229  -0.0390  0.6994  0.8204
   2.000   0.3836   0.00690   0.00220  -0.0377  0.6507  0.8403
   2.500   0.4352   0.00704   0.00219  -0.0368  0.5882  0.8593
   3.000   0.4836   0.00726   0.00234  -0.0352  0.5360  0.8825
   3.500   0.5239   0.00854   0.00273  -0.0327  0.3177  0.9040
   4.000   0.5487   0.01085   0.00375  -0.0278  0.0382  0.9382
   4.500   0.5966   0.01222   0.00541  -0.0263  0.0227  0.9701
   5.000   0.6508   0.01382   0.00721  -0.0268  0.0221  1.0000
   5.500   0.6958   0.01605   0.00961  -0.0254  0.0221  1.0000
   6.000   0.7490   0.01591   0.00955  -0.0260  0.0142  1.0000
   6.500   0.7886   0.01941   0.01329  -0.0239  0.0114  1.0000
   7.000   0.8370   0.02001   0.01393  -0.0233  0.0105  1.0000
   7.500   0.8822   0.02122   0.01522  -0.0222  0.0094  1.0000
   8.000   0.9245   0.02245   0.01653  -0.0208  0.0084  1.0000
   8.500   0.9526   0.02660   0.02102  -0.0176  0.0069  1.0000
   9.500   0.9650   0.03789   0.03344  -0.0059  0.0068  1.0000
  10.000   0.9510   0.04089   0.03673   0.0026  0.0070  1.0000
  10.500   0.9313   0.04464   0.04078   0.0096  0.0072  1.0000