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N-11 (n11-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: N-11 (n11-il)
Reynolds number: 50,000
Max Cl/Cd: 27.1 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n11-il-50000.txt
Download as CSV file: xf-n11-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-11                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3292   0.10967   0.10291  -0.0281   1.0000   0.2139
  -9.500  -0.3257   0.10614   0.09945  -0.0271   1.0000   0.2182
  -9.250  -0.3221   0.10358   0.09695  -0.0255   1.0000   0.2271
  -9.000  -0.3540   0.10435   0.09795  -0.0235   1.0000   0.2305
  -8.750  -0.3306   0.09927   0.09285  -0.0218   1.0000   0.2408
  -8.500  -0.3596   0.09950   0.09327  -0.0188   1.0000   0.2456
  -8.250  -0.3541   0.09612   0.08995  -0.0165   1.0000   0.2529
  -8.000  -0.3724   0.09533   0.08930  -0.0137   1.0000   0.2604
  -7.750  -0.3899   0.09378   0.08790  -0.0120   1.0000   0.2654
  -7.500  -0.3893   0.09148   0.08565  -0.0092   1.0000   0.2759
  -7.000  -0.4142   0.08830   0.08266  -0.0071   1.0000   0.2946
  -6.750  -0.4066   0.08495   0.07934  -0.0032   1.0000   0.3033
  -6.500  -0.4142   0.08284   0.07731  -0.0022   1.0000   0.3156
  -6.250  -0.4188   0.08071   0.07524  -0.0012   1.0000   0.3303
  -6.000  -0.4197   0.07840   0.07298   0.0006   1.0000   0.3465
  -5.750  -0.4180   0.07603   0.07067   0.0035   1.0000   0.3645
  -5.500  -0.4213   0.07432   0.06900   0.0044   1.0000   0.3928
  -5.250  -0.4170   0.07176   0.06652   0.0100   1.0000   0.4150
  -5.000  -0.4171   0.06974   0.06457   0.0142   1.0000   0.4472
  -4.500  -0.2695   0.05079   0.04341  -0.0428   1.0000   0.1993
  -4.250  -0.2416   0.04675   0.03909  -0.0449   1.0000   0.1791
  -4.000  -0.2056   0.04326   0.03474  -0.0487   1.0000   0.1667
  -3.750  -0.1800   0.04125   0.03240  -0.0498   1.0000   0.1661
  -3.500  -0.1541   0.03946   0.03028  -0.0508   1.0000   0.1649
  -3.250  -0.1277   0.03794   0.02837  -0.0516   1.0000   0.1643
  -3.000  -0.1015   0.03676   0.02679  -0.0523   1.0000   0.1654
  -2.750  -0.0765   0.03585   0.02560  -0.0528   1.0000   0.1694
  -2.500  -0.0544   0.03538   0.02507  -0.0529   1.0000   0.1779
  -2.250  -0.0132   0.03487   0.02432  -0.0561   0.9949   0.1908
  -2.000   0.0448   0.03436   0.02368  -0.0617   0.9839   0.2227
  -1.750   0.1040   0.03327   0.02300  -0.0673   0.9718   0.3365
  -1.500   0.1453   0.03061   0.02247  -0.0684   0.9583   1.0000
  -1.250   0.1994   0.03159   0.02271  -0.0737   0.9419   1.0000
  -1.000   0.2407   0.03233   0.02306  -0.0768   0.9252   1.0000
  -0.750   0.2804   0.03304   0.02348  -0.0795   0.9094   1.0000
  -0.500   0.3187   0.03371   0.02391  -0.0819   0.8939   1.0000
  -0.250   0.3560   0.03434   0.02434  -0.0841   0.8782   1.0000
   0.000   0.3927   0.03493   0.02478  -0.0860   0.8624   1.0000
   0.250   0.4289   0.03549   0.02520  -0.0877   0.8465   1.0000
   0.500   0.4648   0.03598   0.02559  -0.0892   0.8301   1.0000
   0.750   0.5004   0.03644   0.02596  -0.0905   0.8137   1.0000
   1.000   0.5358   0.03683   0.02629  -0.0916   0.7972   1.0000
   1.250   0.5714   0.03713   0.02654  -0.0926   0.7805   1.0000
   1.500   0.6066   0.03736   0.02675  -0.0933   0.7641   1.0000
   1.750   0.6420   0.03749   0.02686  -0.0939   0.7479   1.0000
   2.000   0.6780   0.03750   0.02687  -0.0944   0.7320   1.0000
   2.250   0.7124   0.03752   0.02689  -0.0946   0.7164   1.0000
   2.500   0.7450   0.03754   0.02693  -0.0945   0.7010   1.0000
   2.750   0.7764   0.03759   0.02700  -0.0942   0.6857   1.0000
   3.000   0.8062   0.03774   0.02717  -0.0938   0.6708   1.0000
   3.250   0.8355   0.03787   0.02733  -0.0932   0.6561   1.0000
   3.500   0.8643   0.03804   0.02754  -0.0926   0.6416   1.0000
   3.750   0.8938   0.03809   0.02763  -0.0919   0.6273   1.0000
   4.000   0.9242   0.03807   0.02763  -0.0913   0.6133   1.0000
   4.250   0.9569   0.03788   0.02750  -0.0908   0.5998   1.0000
   4.500   0.9992   0.03704   0.02667  -0.0912   0.5880   1.0000
   4.750   1.0221   0.03771   0.02740  -0.0902   0.5740   1.0000
   5.000   1.0389   0.03888   0.02866  -0.0887   0.5600   1.0000
   5.250   1.0530   0.04032   0.03021  -0.0872   0.5465   1.0000
   5.500   1.0687   0.04169   0.03168  -0.0858   0.5336   1.0000
   5.750   1.0931   0.04241   0.03249  -0.0850   0.5217   1.0000
   6.000   1.1345   0.04193   0.03208  -0.0855   0.5111   1.0000
   6.250   1.1311   0.04486   0.03519  -0.0828   0.4984   1.0000
   6.500   1.1277   0.04791   0.03837  -0.0803   0.4864   1.0000
   6.750   1.1475   0.04911   0.03969  -0.0793   0.4755   1.0000
   7.000   1.1925   0.04838   0.03908  -0.0798   0.4651   1.0000
   7.250   1.0890   0.06036   0.05111  -0.0731   0.4545   1.0000
   7.500   1.0010   0.07484   0.06545  -0.0738   0.4448   1.0000
   7.750   0.9871   0.08059   0.07126  -0.0741   0.4369   1.0000
   8.000   1.2451   0.05444   0.04565  -0.0731   0.4177   1.0000
   8.250   1.2789   0.05370   0.04503  -0.0722   0.4027   1.0000
   8.500   1.1269   0.07203   0.06331  -0.0665   0.4032   1.0000
   8.750   1.1696   0.06920   0.06067  -0.0641   0.3888   1.0000
   9.000   1.0160   0.09535   0.08646  -0.0708   0.3880   1.0000
   9.250   0.9594   0.10673   0.09776  -0.0739   0.3851   1.0000
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