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NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 500,000
Max Cl/Cd: 59.37 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66206-il-500000-n5.txt
Download as CSV file: xf-naca66206-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5252   0.10135   0.09898  -0.0124   1.0000   0.0047
  -9.000  -0.5221   0.09756   0.09521  -0.0141   1.0000   0.0047
  -8.750  -0.5201   0.09354   0.09121  -0.0159   1.0000   0.0047
  -8.500  -0.5187   0.08954   0.08724  -0.0178   1.0000   0.0047
  -8.250  -0.5170   0.08590   0.08362  -0.0198   1.0000   0.0047
  -8.000  -0.5176   0.08203   0.07978  -0.0220   1.0000   0.0047
  -7.750  -0.5214   0.07811   0.07590  -0.0250   1.0000   0.0047
  -7.500  -0.5194   0.07378   0.07155  -0.0285   1.0000   0.0047
  -7.250  -0.5227   0.06883   0.06658  -0.0307   1.0000   0.0044
  -7.000  -0.5230   0.06412   0.06184  -0.0326   1.0000   0.0041
  -6.750  -0.5184   0.06007   0.05773  -0.0337   1.0000   0.0039
  -6.500  -0.5136   0.05597   0.05355  -0.0341   1.0000   0.0037
  -6.250  -0.4995   0.05131   0.04876  -0.0361   0.9978   0.0035
  -6.000  -0.4746   0.04572   0.04295  -0.0397   0.9924   0.0033
  -5.750  -0.4475   0.04020   0.03716  -0.0426   0.9880   0.0031
  -5.500  -0.4201   0.03482   0.03145  -0.0444   0.9833   0.0030
  -5.250  -0.3912   0.02967   0.02591  -0.0457   0.9792   0.0029
  -5.000  -0.3627   0.02490   0.02069  -0.0462   0.9747   0.0030
  -4.750  -0.3337   0.02034   0.01562  -0.0461   0.9702   0.0034
  -4.250  -0.2758   0.01456   0.00902  -0.0461   0.9618   0.0053
  -4.000  -0.2467   0.01317   0.00741  -0.0464   0.9577   0.0059
  -3.750  -0.2179   0.01185   0.00591  -0.0466   0.9539   0.0065
  -3.500  -0.1911   0.01094   0.00489  -0.0463   0.9483   0.0084
  -3.250  -0.1627   0.01053   0.00435  -0.0465   0.9437   0.0103
  -3.000  -0.1369   0.00958   0.00331  -0.0463   0.9383   0.0135
  -2.750  -0.1100   0.00917   0.00282  -0.0462   0.9329   0.0157
  -2.500  -0.0824   0.00887   0.00246  -0.0462   0.9286   0.0184
  -2.250  -0.0558   0.00854   0.00205  -0.0460   0.9228   0.0201
  -2.000  -0.0286   0.00834   0.00174  -0.0459   0.9178   0.0221
  -1.750  -0.0016   0.00807   0.00144  -0.0458   0.9128   0.0330
  -1.500   0.0252   0.00780   0.00128  -0.0458   0.9077   0.0745
  -1.250   0.0466   0.00605   0.00105  -0.0458   0.9028   0.5418
  -1.000   0.0666   0.00533   0.00118  -0.0441   0.8969   0.7884
  -0.750   0.0905   0.00527   0.00123  -0.0429   0.8920   0.8392
  -0.500   0.1146   0.00526   0.00128  -0.0419   0.8876   0.8688
  -0.250   0.1398   0.00526   0.00130  -0.0411   0.8824   0.8852
   0.000   0.1666   0.00527   0.00129  -0.0409   0.8777   0.8921
   0.250   0.1934   0.00527   0.00129  -0.0407   0.8716   0.8976
   0.500   0.2199   0.00528   0.00127  -0.0403   0.8616   0.9031
   0.750   0.2448   0.00527   0.00123  -0.0395   0.8419   0.9085
   1.000   0.2705   0.00530   0.00121  -0.0389   0.8235   0.9147
   1.250   0.2961   0.00532   0.00122  -0.0383   0.8052   0.9200
   1.500   0.3206   0.00540   0.00120  -0.0374   0.7697   0.9265
   1.750   0.3346   0.00625   0.00120  -0.0343   0.5491   0.9346
   2.000   0.3480   0.00784   0.00158  -0.0323   0.2352   0.9443
   2.250   0.3691   0.00873   0.00187  -0.0314   0.0669   0.9532
   2.500   0.3958   0.00905   0.00209  -0.0314   0.0353   0.9613
   2.750   0.4244   0.00933   0.00239  -0.0317   0.0260   0.9697
   3.000   0.4550   0.00957   0.00268  -0.0325   0.0212   0.9783
   3.250   0.4862   0.00995   0.00315  -0.0334   0.0177   0.9888
   3.500   0.5125   0.01076   0.00407  -0.0334   0.0145   1.0000
   3.750   0.5376   0.01124   0.00462  -0.0330   0.0136   1.0000
   4.000   0.5635   0.01158   0.00501  -0.0328   0.0116   1.0000
   4.250   0.5882   0.01224   0.00576  -0.0323   0.0098   1.0000
   4.500   0.6130   0.01285   0.00638  -0.0320   0.0067   1.0000
   4.750   0.6378   0.01360   0.00728  -0.0315   0.0057   1.0000
   5.000   0.6622   0.01472   0.00856  -0.0308   0.0044   1.0000
   5.250   0.6865   0.01623   0.01027  -0.0301   0.0038   1.0000
   5.500   0.7107   0.01830   0.01264  -0.0292   0.0035   1.0000
   5.750   0.7330   0.02206   0.01689  -0.0276   0.0034   1.0000
   6.000   0.7513   0.02779   0.02324  -0.0249   0.0035   1.0000
   6.250   0.7667   0.03374   0.02972  -0.0221   0.0037   1.0000
   6.500   0.7803   0.03916   0.03556  -0.0199   0.0040   1.0000
   6.750   0.7920   0.04432   0.04106  -0.0180   0.0042   1.0000
   7.000   0.8016   0.04939   0.04641  -0.0165   0.0045   1.0000
   7.250   0.8099   0.05417   0.05142  -0.0155   0.0045   1.0000
   7.500   0.8159   0.05897   0.05642  -0.0147   0.0044   1.0000
   7.750   0.8192   0.06376   0.06137  -0.0144   0.0042   1.0000
   8.000   0.8256   0.06677   0.06449  -0.0145   0.0035   1.0000
   8.250   0.8297   0.06930   0.06709  -0.0147   0.0031   1.0000
   8.500   0.8226   0.07344   0.07133  -0.0154   0.0028   1.0000
   8.750   0.8134   0.07785   0.07580  -0.0155   0.0030   1.0000
   9.000   0.7959   0.08290   0.08089  -0.0182   0.0028   1.0000
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