XFOIL         Version 6.94
  
 Calculated polar for: NACA 64(1)-212                                  
  
 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   0.1690   0.00915   0.00349  -0.0394  0.7526  0.7525
   0.500   0.2251   0.00920   0.00353  -0.0394  0.7389  0.7680
   1.000   0.2810   0.00926   0.00360  -0.0393  0.7258  0.7839
   1.500   0.3364   0.00935   0.00374  -0.0391  0.7132  0.8000
   2.000   0.3913   0.00942   0.00390  -0.0388  0.6980  0.8178
   2.500   0.4454   0.00951   0.00410  -0.0382  0.6837  0.8353
   3.000   0.4971   0.00945   0.00409  -0.0369  0.6582  0.8542
   3.500   0.5463   0.00934   0.00400  -0.0349  0.6183  0.8745
   4.000   0.5951   0.00938   0.00413  -0.0330  0.5867  0.8963
   4.500   0.6399   0.00944   0.00427  -0.0303  0.5428  0.9206
   5.000   0.6773   0.00992   0.00439  -0.0264  0.4058  0.9529
   5.500   0.7156   0.01369   0.00631  -0.0265  0.0696  1.0000
   6.000   0.7615   0.01519   0.00771  -0.0260  0.0472  1.0000
   6.500   0.8037   0.01696   0.00953  -0.0246  0.0384  1.0000
   7.000   0.8439   0.01893   0.01148  -0.0229  0.0337  1.0000
   7.500   0.8860   0.02126   0.01391  -0.0212  0.0310  1.0000
   8.000   0.9329   0.02357   0.01639  -0.0202  0.0292  1.0000
   8.500   0.9815   0.02650   0.01954  -0.0195  0.0281  1.0000
   9.000   1.0276   0.03013   0.02351  -0.0186  0.0274  1.0000
   9.500   1.0658   0.03380   0.02737  -0.0175  0.0255  1.0000
  10.500   1.0981   0.04502   0.03979  -0.0106  0.0243  1.0000
  11.000   1.0907   0.04996   0.04525  -0.0052  0.0240  1.0000
  11.500   1.0644   0.05653   0.05226  -0.0006  0.0242  1.0000
  12.000   1.0304   0.06440   0.06052   0.0010  0.0243  1.0000