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NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 100,000
Max Cl/Cd: 35.94 at α=2.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66206-il-100000-n5.txt
Download as CSV file: xf-naca66206-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4522   0.08939   0.08472  -0.0246   1.0000   0.0352
  -8.500  -0.4555   0.08534   0.08071  -0.0272   1.0000   0.0352
  -8.250  -0.4592   0.08127   0.07668  -0.0297   1.0000   0.0353
  -8.000  -0.4650   0.07739   0.07282  -0.0320   1.0000   0.0353
  -7.750  -0.4688   0.07346   0.06887  -0.0335   1.0000   0.0353
  -7.500  -0.4704   0.06937   0.06473  -0.0347   1.0000   0.0354
  -7.250  -0.4703   0.06522   0.06052  -0.0355   1.0000   0.0354
  -7.000  -0.4719   0.05886   0.05422  -0.0358   1.0000   0.0359
  -6.750  -0.4700   0.05378   0.04916  -0.0354   1.0000   0.0366
  -6.500  -0.4674   0.04935   0.04472  -0.0349   1.0000   0.0372
  -6.250  -0.4642   0.04531   0.04062  -0.0345   1.0000   0.0377
  -6.000  -0.4599   0.04137   0.03659  -0.0340   1.0000   0.0383
  -5.750  -0.4542   0.03758   0.03266  -0.0333   1.0000   0.0387
  -5.250  -0.4501   0.04407   0.03800  -0.0324   1.0000   0.0206
  -5.000  -0.4370   0.04051   0.03420  -0.0314   1.0000   0.0200
  -4.750  -0.4216   0.03724   0.03062  -0.0302   1.0000   0.0195
  -4.500  -0.4042   0.03411   0.02712  -0.0289   1.0000   0.0192
  -4.250  -0.3851   0.03121   0.02382  -0.0276   1.0000   0.0190
  -4.000  -0.3642   0.02852   0.02070  -0.0263   1.0000   0.0191
  -3.750  -0.3417   0.02672   0.01843  -0.0249   1.0000   0.0211
  -3.500  -0.3196   0.02414   0.01541  -0.0240   1.0000   0.0240
  -3.250  -0.2961   0.02205   0.01298  -0.0230   1.0000   0.0249
  -3.000  -0.2723   0.02031   0.01090  -0.0219   1.0000   0.0262
  -2.750  -0.2487   0.01889   0.00929  -0.0209   1.0000   0.0290
  -2.500  -0.2257   0.01830   0.00846  -0.0198   1.0000   0.0353
  -2.250  -0.2031   0.01674   0.00690  -0.0189   1.0000   0.0377
  -2.000  -0.1803   0.01590   0.00597  -0.0181   1.0000   0.0412
  -1.750  -0.1503   0.01530   0.00516  -0.0188   0.9972   0.0480
  -1.500  -0.1191   0.01464   0.00456  -0.0198   0.9940   0.0783
  -1.250  -0.1150   0.01184   0.00480  -0.0139   0.9916   0.8625
  -1.000  -0.0640   0.01172   0.00449  -0.0176   0.9924   1.0000
  -0.750  -0.0314   0.01183   0.00436  -0.0192   0.9880   1.0000
  -0.500   0.0002   0.01192   0.00428  -0.0205   0.9833   1.0000
  -0.250   0.0328   0.01205   0.00424  -0.0221   0.9791   1.0000
   0.000   0.0638   0.01217   0.00425  -0.0233   0.9743   1.0000
   0.250   0.0959   0.01231   0.00432  -0.0247   0.9698   1.0000
   0.500   0.1285   0.01246   0.00444  -0.0262   0.9657   1.0000
   0.750   0.1585   0.01260   0.00457  -0.0272   0.9601   1.0000
   1.000   0.1940   0.01276   0.00476  -0.0292   0.9558   1.0000
   1.250   0.2242   0.01290   0.00495  -0.0301   0.9490   1.0000
   1.500   0.2597   0.01304   0.00518  -0.0320   0.9438   1.0000
   1.750   0.2915   0.01318   0.00544  -0.0331   0.9372   1.0000
   2.000   0.3274   0.01329   0.00580  -0.0350   0.9311   1.0000
   2.250   0.3776   0.01296   0.00574  -0.0385   0.9106   1.0000
   2.500   0.4263   0.01186   0.00487  -0.0392   0.8445   1.0000
   2.750   0.4418   0.01233   0.00376  -0.0328   0.4410   1.0000
   3.000   0.4486   0.01522   0.00462  -0.0300   0.0653   1.0000
   3.250   0.4711   0.01607   0.00562  -0.0290   0.0510   1.0000
   3.500   0.4937   0.01693   0.00667  -0.0279   0.0451   1.0000
   3.750   0.5145   0.01810   0.00791  -0.0267   0.0386   1.0000
   4.000   0.5374   0.01915   0.00908  -0.0258   0.0324   1.0000
   4.250   0.5607   0.02073   0.01072  -0.0248   0.0295   1.0000
   4.500   0.5861   0.02322   0.01331  -0.0241   0.0275   1.0000
   4.750   0.6128   0.02532   0.01574  -0.0233   0.0253   1.0000
   5.000   0.6388   0.02708   0.01795  -0.0224   0.0212   1.0000
   5.250   0.6635   0.02990   0.02137  -0.0211   0.0202   1.0000
   5.500   0.6860   0.03308   0.02507  -0.0197   0.0196   1.0000
   5.750   0.7063   0.03656   0.02907  -0.0181   0.0195   1.0000
   6.000   0.7244   0.04021   0.03320  -0.0165   0.0193   1.0000
   6.250   0.7399   0.04288   0.03613  -0.0157   0.0168   1.0000
   6.750   0.7550   0.05209   0.04595  -0.0136   0.0145   1.0000
   7.000   0.7648   0.05600   0.05020  -0.0124   0.0143   1.0000
   7.250   0.7703   0.06075   0.05521  -0.0117   0.0147   1.0000
   7.500   0.7801   0.06398   0.05869  -0.0108   0.0150   1.0000
   7.750   0.7890   0.06760   0.06257  -0.0101   0.0157   1.0000
   8.000   0.7872   0.07359   0.06888  -0.0106   0.0177   1.0000
   8.250   0.7815   0.07868   0.07409  -0.0116   0.0186   1.0000
   8.500   0.7715   0.08341   0.07887  -0.0129   0.0195   1.0000
   8.750   0.7628   0.08842   0.08390  -0.0158   0.0200   1.0000
   9.000   0.7561   0.09427   0.08973  -0.0206   0.0208   1.0000
   9.250   0.7524   0.09988   0.09529  -0.0245   0.0216   1.0000
   9.500   0.7504   0.10512   0.10050  -0.0275   0.0229   1.0000
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