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

EPPLER 668 AIRFOIL (e668-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 668 AIRFOIL (e668-il)
Reynolds number: 100,000
Max Cl/Cd: 51.6 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e668-il-100000.txt
Download as CSV file: xf-e668-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 668 AIRFOIL                              
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.5467   0.10185   0.09817  -0.0219   1.0000   0.1122
  -6.750  -0.5820   0.09678   0.09307  -0.0285   1.0000   0.1135
  -6.500  -0.6050   0.09121   0.08732  -0.0328   1.0000   0.1147
  -6.250  -0.5802   0.08994   0.08631  -0.0241   1.0000   0.1183
  -6.000  -0.5812   0.08692   0.08328  -0.0236   1.0000   0.1217
  -5.750  -0.5974   0.08012   0.07616  -0.0314   1.0000   0.1304
  -5.500  -0.5885   0.07749   0.07367  -0.0281   1.0000   0.1330
  -5.250  -0.5818   0.07457   0.07072  -0.0277   1.0000   0.1380
  -5.000  -0.5761   0.06979   0.06575  -0.0304   1.0000   0.1483
  -4.750  -0.5319   0.04788   0.04175  -0.0418   1.0000   0.0593
  -4.500  -0.5025   0.04201   0.03467  -0.0429   1.0000   0.0517
  -4.250  -0.4784   0.03903   0.03122  -0.0430   1.0000   0.0513
  -4.000  -0.4537   0.03686   0.02853  -0.0431   1.0000   0.0528
  -3.750  -0.4325   0.03548   0.02719  -0.0431   1.0000   0.0569
  -3.500  -0.4074   0.03397   0.02527  -0.0427   1.0000   0.0602
  -3.250  -0.3835   0.03261   0.02375  -0.0424   1.0000   0.0650
  -3.000  -0.3602   0.03207   0.02299  -0.0420   1.0000   0.0733
  -2.750  -0.3369   0.03123   0.02222  -0.0418   1.0000   0.0847
  -2.500  -0.3130   0.03057   0.02160  -0.0417   1.0000   0.1009
  -2.250  -0.2885   0.03002   0.02123  -0.0418   1.0000   0.1283
  -2.000  -0.2621   0.02936   0.02103  -0.0425   1.0000   0.1909
  -1.750  -0.2315   0.02820   0.02166  -0.0442   0.9992   0.4810
  -1.500  -0.2193   0.02905   0.02360  -0.0390   0.9945   0.7588
  -1.250  -0.2065   0.03001   0.02451  -0.0352   0.9878   0.8394
  -1.000  -0.1389   0.03170   0.02593  -0.0372   0.9234   0.9318
  -0.750  -0.0220   0.03296   0.02675  -0.0524   0.9002   0.9970
  -0.500  -0.0066   0.03304   0.02665  -0.0516   0.8900   1.0000
  -0.250   0.0412   0.03389   0.02724  -0.0558   0.8818   1.0000
   0.000   0.0474   0.03372   0.02695  -0.0533   0.8707   1.0000
   0.250   0.0779   0.03417   0.02723  -0.0547   0.8607   1.0000
   0.500   0.1144   0.03468   0.02758  -0.0569   0.8524   1.0000
   0.750   0.1370   0.03506   0.02784  -0.0570   0.8419   1.0000
   1.000   0.1840   0.03569   0.02830  -0.0607   0.8353   1.0000
   1.250   0.2016   0.03599   0.02853  -0.0601   0.8238   1.0000
   1.500   0.2357   0.03650   0.02894  -0.0617   0.8154   1.0000
   1.750   0.2668   0.03691   0.02926  -0.0630   0.8067   1.0000
   2.000   0.2924   0.03736   0.02965  -0.0634   0.7966   1.0000
   2.250   0.3324   0.03772   0.02993  -0.0657   0.7899   1.0000
   2.500   0.3534   0.03818   0.03035  -0.0655   0.7790   1.0000
   2.750   0.3971   0.03842   0.03054  -0.0681   0.7736   1.0000
   3.000   0.4159   0.03887   0.03098  -0.0676   0.7621   1.0000
   3.250   0.4627   0.03892   0.03100  -0.0704   0.7577   1.0000
   3.500   0.4801   0.03937   0.03145  -0.0697   0.7458   1.0000
   3.750   0.5063   0.03970   0.03177  -0.0700   0.7362   1.0000
   4.000   0.5466   0.03957   0.03167  -0.0718   0.7301   1.0000
   4.250   0.5685   0.03994   0.03205  -0.0715   0.7191   1.0000
   4.500   0.6139   0.03950   0.03163  -0.0737   0.7147   1.0000
   4.750   0.6343   0.03983   0.03200  -0.0731   0.7028   1.0000
   5.000   0.6826   0.03904   0.03128  -0.0754   0.6994   1.0000
   5.250   0.7021   0.03931   0.03160  -0.0746   0.6870   1.0000
   5.500   0.7521   0.03816   0.03055  -0.0768   0.6841   1.0000
   5.750   0.7721   0.03830   0.03075  -0.0759   0.6713   1.0000
   6.000   0.7957   0.03828   0.03081  -0.0753   0.6596   1.0000
   6.250   0.8454   0.03664   0.02930  -0.0770   0.6560   1.0000
   6.500   0.8697   0.03641   0.02916  -0.0763   0.6440   1.0000
   6.750   0.9211   0.03423   0.02714  -0.0778   0.6409   1.0000
   7.000   0.9758   0.03164   0.02470  -0.0794   0.6386   1.0000
   7.250   1.0043   0.03077   0.02398  -0.0787   0.6261   1.0000
   7.750   1.0925   0.02713   0.02062  -0.0801   0.6038   1.0000
   8.000   1.1363   0.02555   0.01917  -0.0809   0.5864   1.0000
   8.250   1.1857   0.02402   0.01768  -0.0826   0.5637   1.0000
   8.500   1.2113   0.02385   0.01751  -0.0817   0.5355   1.0000
   8.750   1.2343   0.02392   0.01752  -0.0805   0.5048   1.0000
   9.000   1.2512   0.02435   0.01785  -0.0787   0.4735   1.0000
   9.250   1.2628   0.02505   0.01846  -0.0762   0.4422   1.0000
   9.500   1.2710   0.02596   0.01925  -0.0735   0.4115   1.0000
   9.750   1.2766   0.02703   0.02021  -0.0707   0.3813   1.0000
  10.000   1.2806   0.02826   0.02131  -0.0678   0.3518   1.0000
  10.250   1.2833   0.02962   0.02253  -0.0649   0.3232   1.0000
  10.500   1.2857   0.03110   0.02384  -0.0622   0.2959   1.0000
  10.750   1.2860   0.03275   0.02543  -0.0595   0.2690   1.0000
  11.000   1.2873   0.03449   0.02706  -0.0570   0.2435   1.0000
  11.250   1.2887   0.03632   0.02871  -0.0546   0.2205   1.0000
  11.500   1.2899   0.03827   0.03063  -0.0524   0.1979   1.0000
  11.750   1.2919   0.04026   0.03250  -0.0504   0.1782   1.0000
  12.000   1.2954   0.04232   0.03442  -0.0486   0.1600   1.0000
  12.250   1.2989   0.04442   0.03652  -0.0469   0.1437   1.0000
  12.500   1.3031   0.04657   0.03866  -0.0454   0.1292   1.0000
  12.750   1.3095   0.04873   0.04078  -0.0440   0.1164   1.0000
  13.000   1.3182   0.05085   0.04282  -0.0429   0.1048   1.0000
  13.250   1.3236   0.05308   0.04515  -0.0417   0.0955   1.0000
  13.500   1.3334   0.05541   0.04760  -0.0407   0.0869   1.0000
  13.750   1.3483   0.05760   0.04974  -0.0400   0.0790   1.0000
  14.000   1.3537   0.06008   0.05244  -0.0389   0.0734   1.0000
  14.250   1.3695   0.06262   0.05501  -0.0384   0.0675   1.0000
  14.500   1.3703   0.06550   0.05817  -0.0373   0.0635   1.0000
  14.750   1.3901   0.06821   0.06080  -0.0372   0.0585   1.0000
  15.000   1.3802   0.07183   0.06484  -0.0359   0.0564   1.0000
  15.250   1.3762   0.07529   0.06855  -0.0351   0.0538   1.0000
  15.500   1.3951   0.07833   0.07146  -0.0352   0.0497   1.0000
  15.750   1.3755   0.08277   0.07632  -0.0345   0.0490   1.0000
  16.000   1.3563   0.08771   0.08164  -0.0343   0.0482   1.0000
  16.250   1.3370   0.09308   0.08735  -0.0349   0.0474   1.0000
  16.500   1.3174   0.09890   0.09348  -0.0360   0.0468   1.0000
  16.750   1.2964   0.10524   0.10012  -0.0380   0.0464   1.0000
  17.000   1.2727   0.11240   0.10755  -0.0408   0.0465   1.0000
  17.250   1.2470   0.12040   0.11582  -0.0446   0.0468   1.0000
  17.500   1.2199   0.12931   0.12496  -0.0496   0.0474   1.0000
  17.750   1.1930   0.13889   0.13473  -0.0554   0.0483   1.0000
  18.000   1.1681   0.14884   0.14481  -0.0617   0.0490   1.0000
  18.250   1.1462   0.15893   0.15498  -0.0681   0.0497   1.0000
  18.500   1.0444   0.20469   0.20058  -0.0943   0.0697   1.0000
<< Back to EPPLER 668 AIRFOIL (e668-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 668 AIRFOIL (e668-il)