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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 100,000
Max Cl/Cd: 68.42 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e58-il-100000-n5.txt
Download as CSV file: xf-e58-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 58 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.3067   0.10775   0.10318  -0.0343   0.9799   0.0295
  -6.500  -0.3028   0.10526   0.10072  -0.0343   0.9769   0.0300
  -6.250  -0.2998   0.10287   0.09836  -0.0345   0.9734   0.0305
  -6.000  -0.2893   0.09993   0.09543  -0.0368   0.9704   0.0310
  -5.750  -0.2805   0.09710   0.09262  -0.0386   0.9671   0.0315
  -5.500  -0.2719   0.09415   0.08970  -0.0404   0.9629   0.0325
  -5.250  -0.2564   0.09082   0.08637  -0.0441   0.9594   0.0331
  -5.000  -0.2388   0.08733   0.08287  -0.0485   0.9561   0.0339
  -4.750  -0.2247   0.08402   0.07958  -0.0518   0.9510   0.0347
  -4.500  -0.1985   0.07993   0.07544  -0.0585   0.9472   0.0355
  -4.250  -0.1337   0.07396   0.06933  -0.0790   0.9443   0.0380
  -3.250   0.0335   0.05169   0.04659  -0.1138   0.9336   0.0267
  -3.000   0.1204   0.04331   0.03775  -0.1332   0.9351   0.0266
  -2.750   0.2066   0.03633   0.03007  -0.1505   0.9377   0.0308
  -2.500   0.2908   0.03080   0.02336  -0.1647   0.9416   0.0345
  -2.250   0.3432   0.02837   0.02015  -0.1706   0.9419   0.0407
  -2.000   0.3789   0.02742   0.01887  -0.1727   0.9395   0.0477
  -1.750   0.4162   0.02642   0.01758  -0.1749   0.9377   0.0538
  -1.500   0.4511   0.02592   0.01693  -0.1766   0.9355   0.0603
  -1.250   0.4867   0.02558   0.01648  -0.1784   0.9335   0.0690
  -1.000   0.5235   0.02534   0.01610  -0.1803   0.9318   0.0750
  -0.750   0.5558   0.02520   0.01597  -0.1815   0.9288   0.0896
  -0.500   0.5830   0.02521   0.01606  -0.1816   0.9241   0.1407
  -0.250   0.6151   0.02526   0.01610  -0.1827   0.9208   0.1874
   0.000   0.6497   0.02529   0.01612  -0.1842   0.9183   0.2206
   0.250   0.6765   0.02532   0.01623  -0.1843   0.9127   0.2561
   0.500   0.7098   0.02511   0.01631  -0.1858   0.9079   0.3447
   0.750   0.7393   0.02395   0.01633  -0.1856   0.9047   0.8435
   1.000   0.7631   0.02389   0.01619  -0.1847   0.8957   1.0000
   1.250   0.8007   0.02377   0.01595  -0.1864   0.8911   1.0000
   1.500   0.8272   0.02378   0.01590  -0.1861   0.8822   1.0000
   1.750   0.8649   0.02356   0.01563  -0.1877   0.8774   1.0000
   2.000   0.8897   0.02359   0.01567  -0.1870   0.8678   1.0000
   2.250   0.9291   0.02320   0.01528  -0.1888   0.8632   1.0000
   2.500   0.9532   0.02317   0.01529  -0.1879   0.8523   1.0000
   2.750   0.9800   0.02303   0.01524  -0.1874   0.8418   1.0000
   3.000   1.0102   0.02272   0.01499  -0.1874   0.8321   1.0000
   3.250   1.0436   0.02222   0.01459  -0.1878   0.8231   1.0000
   3.500   1.0678   0.02208   0.01460  -0.1867   0.8101   1.0000
   3.750   1.0918   0.02198   0.01462  -0.1856   0.7966   1.0000
   4.250   1.1415   0.02173   0.01466  -0.1835   0.7635   1.0000
   4.500   1.1682   0.02156   0.01468  -0.1828   0.7446   1.0000
   4.750   1.2045   0.02075   0.01401  -0.1831   0.7152   1.0000
   5.000   1.2683   0.01898   0.01219  -0.1877   0.6572   1.0000
   5.250   1.3191   0.01928   0.01151  -0.1905   0.4837   1.0000
   5.500   1.3295   0.02079   0.01237  -0.1871   0.3846   1.0000
   5.750   1.3343   0.02270   0.01348  -0.1833   0.2665   1.0000
   6.000   1.3428   0.02468   0.01474  -0.1803   0.1624   1.0000
   6.250   1.3495   0.02716   0.01643  -0.1772   0.0548   1.0000
   6.500   1.3628   0.02884   0.01797  -0.1746   0.0278   1.0000
   6.750   1.3774   0.03030   0.01951  -0.1723   0.0220   1.0000
   7.000   1.3905   0.03188   0.02136  -0.1697   0.0196   1.0000
   7.250   1.4001   0.03378   0.02349  -0.1666   0.0182   1.0000
   7.500   1.4092   0.03565   0.02559  -0.1635   0.0172   1.0000
   7.750   1.4198   0.03744   0.02762  -0.1607   0.0160   1.0000
   8.000   1.4306   0.03945   0.02984  -0.1580   0.0148   1.0000
   8.250   1.4437   0.04173   0.03232  -0.1556   0.0143   1.0000
   8.500   1.4622   0.04428   0.03509  -0.1539   0.0138   1.0000
   8.750   1.4862   0.04726   0.03835  -0.1530   0.0133   1.0000
   9.000   1.5108   0.05060   0.04204  -0.1523   0.0130   1.0000
   9.250   1.5258   0.05385   0.04567  -0.1504   0.0123   1.0000
   9.500   1.5316   0.05696   0.04912  -0.1476   0.0115   1.0000
   9.750   1.5348   0.06047   0.05300  -0.1445   0.0112   1.0000
  10.000   1.5329   0.06413   0.05711  -0.1411   0.0109   1.0000
  10.250   1.5269   0.06809   0.06144  -0.1376   0.0107   1.0000
  10.500   1.5177   0.07222   0.06595  -0.1339   0.0107   1.0000
  10.750   1.5060   0.07648   0.07058  -0.1306   0.0107   1.0000
  11.000   1.4928   0.08086   0.07532  -0.1277   0.0108   1.0000
  11.250   1.4766   0.08574   0.08054  -0.1253   0.0108   1.0000
  11.500   1.4591   0.09089   0.08600  -0.1237   0.0108   1.0000
  11.750   1.4409   0.09645   0.09187  -0.1230   0.0109   1.0000
  12.000   1.4222   0.10238   0.09809  -0.1233   0.0110   1.0000
  12.250   1.4027   0.10893   0.10491  -0.1250   0.0111   1.0000
  12.500   1.3833   0.11608   0.11229  -0.1279   0.0111   1.0000
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