Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 100,000
Max Cl/Cd: 28.44 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66206-il-100000.txt
Download as CSV file: xf-naca66206-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-206                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5432   0.09681   0.09191  -0.0216   1.0000   0.0593
  -8.250  -0.5508   0.09359   0.08876  -0.0266   1.0000   0.0596
  -8.000  -0.5546   0.09050   0.08562  -0.0313   1.0000   0.0599
  -7.750  -0.5396   0.08423   0.07949  -0.0240   1.0000   0.0639
  -7.500  -0.5369   0.08057   0.07585  -0.0256   1.0000   0.0665
  -7.250  -0.5355   0.07674   0.07200  -0.0285   1.0000   0.0696
  -7.000  -0.5344   0.07357   0.06868  -0.0336   1.0000   0.0725
  -6.500  -0.5231   0.06508   0.06024  -0.0328   1.0000   0.0786
  -6.250  -0.5165   0.06334   0.05804  -0.0356   1.0000   0.0863
  -6.000  -0.5094   0.05783   0.05276  -0.0340   1.0000   0.0899
  -5.750  -0.5002   0.05587   0.05035  -0.0348   1.0000   0.1006
  -5.500  -0.4905   0.05147   0.04622  -0.0329   1.0000   0.1068
  -5.250  -0.4804   0.04838   0.04300  -0.0322   1.0000   0.1189
  -5.000  -0.4694   0.04558   0.04008  -0.0312   1.0000   0.1343
  -4.750  -0.4588   0.04282   0.03719  -0.0303   1.0000   0.1568
  -4.500  -0.4482   0.04020   0.03449  -0.0287   1.0000   0.1849
  -4.250  -0.4388   0.03784   0.03219  -0.0266   1.0000   0.2250
  -3.250  -0.3138   0.02514   0.01662  -0.0245   1.0000   0.0826
  -3.000  -0.2863   0.02279   0.01376  -0.0228   1.0000   0.0689
  -2.750  -0.2599   0.02087   0.01140  -0.0214   1.0000   0.0631
  -2.500  -0.2348   0.01942   0.00972  -0.0202   1.0000   0.0618
  -2.250  -0.2110   0.01800   0.00831  -0.0192   1.0000   0.0665
  -2.000  -0.1875   0.01694   0.00720  -0.0182   1.0000   0.0750
  -1.750  -0.1646   0.01585   0.00610  -0.0172   1.0000   0.0814
  -1.500  -0.1414   0.01504   0.00523  -0.0163   1.0000   0.0949
  -1.250  -0.1098   0.01166   0.00496  -0.0136   1.0000   1.0000
  -1.000  -0.0925   0.01168   0.00465  -0.0121   1.0000   1.0000
  -0.750  -0.0733   0.01174   0.00446  -0.0111   1.0000   1.0000
  -0.500  -0.0529   0.01182   0.00436  -0.0103   1.0000   1.0000
  -0.250  -0.0317   0.01193   0.00431  -0.0096   1.0000   1.0000
   0.000  -0.0102   0.01206   0.00427  -0.0090   1.0000   1.0000
   0.250   0.0117   0.01222   0.00432  -0.0085   1.0000   1.0000
   0.500   0.0337   0.01240   0.00442  -0.0080   1.0000   1.0000
   0.750   0.0557   0.01259   0.00456  -0.0076   1.0000   1.0000
   1.000   0.0778   0.01281   0.00475  -0.0072   1.0000   1.0000
   1.250   0.0999   0.01305   0.00499  -0.0068   1.0000   1.0000
   1.500   0.1219   0.01332   0.00528  -0.0065   1.0000   1.0000
   1.750   0.1438   0.01360   0.00559  -0.0061   1.0000   1.0000
   2.000   0.1655   0.01392   0.00596  -0.0058   1.0000   1.0000
   2.250   0.1871   0.01426   0.00637  -0.0054   1.0000   1.0000
   2.500   0.2084   0.01463   0.00684  -0.0051   1.0000   1.0000
   2.750   0.2296   0.01504   0.00737  -0.0048   1.0000   1.0000
   3.000   0.2673   0.01570   0.00835  -0.0079   0.9934   1.0000
   3.250   0.4692   0.01650   0.00642  -0.0274   0.0806   1.0000
   3.500   0.4899   0.01778   0.00766  -0.0260   0.0677   1.0000
   3.750   0.5128   0.01982   0.00951  -0.0249   0.0634   1.0000
   4.000   0.5410   0.02158   0.01150  -0.0241   0.0622   1.0000
   4.250   0.5704   0.02399   0.01408  -0.0235   0.0627   1.0000
   4.500   0.5979   0.02584   0.01634  -0.0225   0.0603   1.0000
   4.750   0.6246   0.02837   0.01927  -0.0214   0.0602   1.0000
   5.000   0.6534   0.03168   0.02321  -0.0196   0.0728   1.0000
   6.750   0.7864   0.05953   0.05421  -0.0111   0.1189   1.0000
   7.000   0.8005   0.06584   0.06027  -0.0108   0.1031   1.0000
   7.250   0.7959   0.06698   0.06208  -0.0102   0.0935   1.0000
   7.500   0.8061   0.07309   0.06805  -0.0099   0.0891   1.0000
   7.750   0.7954   0.07553   0.07093  -0.0109   0.0829   1.0000
   8.000   0.8055   0.07960   0.07497  -0.0101   0.0784   1.0000
   8.250   0.8107   0.08607   0.08136  -0.0102   0.0761   1.0000
   8.500   0.7937   0.08916   0.08466  -0.0122   0.0753   1.0000
   8.750   0.7766   0.09300   0.08855  -0.0146   0.0745   1.0000
   9.000   0.7663   0.09790   0.09344  -0.0189   0.0741   1.0000
   9.250   0.7576   0.10317   0.09865  -0.0243   0.0725   1.0000
   9.500   0.7542   0.10791   0.10334  -0.0276   0.0699   1.0000
   9.750   0.7553   0.11225   0.10767  -0.0294   0.0676   1.0000
  10.000   0.6609   0.11367   0.10942  -0.0240   0.0755   1.0000
  10.250   0.6499   0.11825   0.11396  -0.0280   0.0743   1.0000
<< Back to NACA 66-206 (naca66206-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66-206 (naca66206-il)