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EPPLER 379 AIRFOIL (e379-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 379 AIRFOIL (e379-il)
Reynolds number: 500,000
Max Cl/Cd: 127.97 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e379-il-500000.txt
Download as CSV file: xf-e379-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 379 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.750   0.2705   0.04624   0.04235  -0.0750   0.5656   0.0032
  -2.500   0.3026   0.04400   0.04005  -0.0783   0.5611   0.0035
  -2.250   0.3353   0.04188   0.03783  -0.0814   0.5570   0.0036
  -2.000   0.3686   0.03981   0.03569  -0.0844   0.5527   0.0036
  -1.750   0.4022   0.03780   0.03361  -0.0873   0.5482   0.0036
  -1.500   0.4357   0.03590   0.03161  -0.0900   0.5441   0.0036
  -1.250   0.4695   0.03407   0.02968  -0.0924   0.5401   0.0036
  -1.000   0.5032   0.03230   0.02781  -0.0947   0.5357   0.0037
  -0.750   0.5368   0.03062   0.02603  -0.0968   0.5318   0.0037
  -0.500   0.5700   0.02907   0.02434  -0.0986   0.5283   0.0038
  -0.250   0.6034   0.02752   0.02272  -0.1003   0.5242   0.0039
   0.000   0.6364   0.02607   0.02116  -0.1017   0.5202   0.0040
   0.250   0.6690   0.02474   0.01968  -0.1030   0.5166   0.0041
   0.500   0.7014   0.02344   0.01828  -0.1040   0.5128   0.0043
   0.750   0.7338   0.02215   0.01683  -0.1049   0.5089   0.0040
   1.000   0.7656   0.02101   0.01556  -0.1056   0.5054   0.0043
   1.250   0.7969   0.02001   0.01439  -0.1060   0.5019   0.0047
   1.500   0.8287   0.01911   0.01340  -0.1063   0.4980   0.0053
   1.750   0.8610   0.01821   0.01235  -0.1063   0.4941   0.0063
   2.000   0.8890   0.01709   0.01111  -0.1069   0.4905   0.0106
   2.250   0.9196   0.01688   0.01077  -0.1064   0.4869   0.0138
   2.500   0.9497   0.01543   0.00917  -0.1067   0.4832   0.0160
   2.750   0.9783   0.01476   0.00846  -0.1067   0.4796   0.0237
   3.000   1.0067   0.01405   0.00757  -0.1066   0.4761   0.0326
   3.250   1.0352   0.01342   0.00685  -0.1065   0.4723   0.0434
   3.500   1.0634   0.01296   0.00626  -0.1063   0.4682   0.0648
   3.750   1.0909   0.01245   0.00566  -0.1061   0.4642   0.0800
   4.500   1.1736   0.01088   0.00367  -0.1045   0.4518   0.0120
   4.750   1.2000   0.01065   0.00381  -0.1044   0.4473   0.3834
   5.000   1.2330   0.00985   0.00383  -0.1058   0.4401   1.0000
   5.250   1.2579   0.00983   0.00351  -0.1055   0.3901   1.0000
   5.500   1.2831   0.01011   0.00363  -0.1053   0.3676   1.0000
   5.750   1.3087   0.01035   0.00380  -0.1052   0.3515   1.0000
   6.000   1.3024   0.01586   0.00714  -0.1030   0.0065   1.0000
   6.250   1.3230   0.01673   0.00781  -0.1022   0.0052   1.0000
   6.500   1.3499   0.01655   0.00773  -0.1022   0.0041   1.0000
   6.750   1.3732   0.01692   0.00811  -0.1018   0.0042   1.0000
   7.000   1.3963   0.01730   0.00853  -0.1014   0.0043   1.0000
   7.250   1.4189   0.01774   0.00902  -0.1009   0.0047   1.0000
   7.500   1.4410   0.01820   0.00958  -0.1003   0.0051   1.0000
   7.750   1.4637   0.01856   0.01004  -0.0999   0.0056   1.0000
   8.000   1.4860   0.01896   0.01045  -0.0993   0.0060   1.0000
   8.250   1.5071   0.01947   0.01100  -0.0985   0.0064   1.0000
   8.500   1.5273   0.02005   0.01166  -0.0977   0.0067   1.0000
   8.750   1.5456   0.02077   0.01255  -0.0966   0.0072   1.0000
  12.250   1.4798   0.06450   0.05887  -0.0740   0.0024   1.0000
  12.500   1.4781   0.06799   0.06254  -0.0747   0.0022   1.0000
  12.750   1.4801   0.07120   0.06591  -0.0745   0.0021   1.0000
  13.000   1.3321   0.08016   0.07538  -0.0673   0.0025   1.0000
  13.250   1.3187   0.08572   0.08115  -0.0690   0.0024   1.0000
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