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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 200,000
Max Cl/Cd: 106.57 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e58-il-200000-n5.txt
Download as CSV file: xf-e58-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.1751   0.09660   0.09301  -0.0698   0.9515   0.0129
  -7.000  -0.1698   0.09412   0.09054  -0.0694   0.9488   0.0134
  -6.750  -0.1623   0.09166   0.08809  -0.0702   0.9464   0.0139
  -6.500  -0.1517   0.08904   0.08548  -0.0720   0.9444   0.0145
  -6.250  -0.1569   0.08757   0.08405  -0.0697   0.9391   0.0148
  -6.000  -0.1499   0.08508   0.08158  -0.0707   0.9353   0.0154
  -5.750  -0.1340   0.08187   0.07837  -0.0741   0.9327   0.0164
  -5.500  -0.1135   0.07831   0.07481  -0.0790   0.9309   0.0176
  -5.250  -0.1035   0.07673   0.07325  -0.0833   0.9234   0.0205
  -5.000  -0.0789   0.07313   0.06964  -0.0891   0.9208   0.0207
  -4.000   0.0104   0.05744   0.05389  -0.1071   0.9092   0.0252
  -3.750   0.0952   0.05108   0.04728  -0.1285   0.9083   0.0329
  -3.250   0.2433   0.03349   0.02893  -0.1611   0.9097   0.0215
  -3.000   0.3206   0.02735   0.02203  -0.1759   0.9123   0.0234
  -2.750   0.3572   0.02643   0.02094  -0.1788   0.9100   0.0295
  -2.500   0.4140   0.02332   0.01712  -0.1860   0.9108   0.0269
  -2.250   0.4587   0.02159   0.01485  -0.1897   0.9103   0.0250
  -2.000   0.4965   0.02057   0.01351  -0.1919   0.9092   0.0238
  -1.750   0.5319   0.01988   0.01262  -0.1935   0.9082   0.0230
  -1.500   0.5667   0.01933   0.01197  -0.1951   0.9070   0.0225
  -1.250   0.6023   0.01885   0.01139  -0.1968   0.9060   0.0222
  -1.000   0.6399   0.01834   0.01081  -0.1988   0.9049   0.0222
  -0.750   0.6646   0.01818   0.01058  -0.1983   0.8999   0.0223
  -0.500   0.6971   0.01783   0.01012  -0.1992   0.8962   0.0231
  -0.250   0.7337   0.01742   0.00959  -0.2008   0.8938   0.0245
   0.250   0.7989   0.01660   0.00914  -0.2027   0.8850   0.1663
   0.500   0.8359   0.01605   0.00882  -0.2045   0.8812   0.2301
   0.750   0.8805   0.01491   0.00869  -0.2082   0.8793   0.6068
   1.000   0.8931   0.01441   0.00868  -0.2045   0.8712   0.8751
   1.250   0.9249   0.01394   0.00819  -0.2048   0.8670   1.0000
   1.500   0.9480   0.01397   0.00821  -0.2037   0.8593   1.0000
   1.750   0.9790   0.01374   0.00797  -0.2041   0.8547   1.0000
   2.000   1.0023   0.01378   0.00806  -0.2031   0.8464   1.0000
   2.250   1.0341   0.01349   0.00778  -0.2036   0.8404   1.0000
   2.500   1.0562   0.01355   0.00787  -0.2023   0.8293   1.0000
   2.750   1.0817   0.01348   0.00783  -0.2016   0.8173   1.0000
   3.000   1.1100   0.01329   0.00768  -0.2014   0.8035   1.0000
   3.250   1.1461   0.01277   0.00722  -0.2024   0.7787   1.0000
   3.500   1.2277   0.01152   0.00561  -0.2122   0.6952   1.0000
   3.750   1.2572   0.01207   0.00568  -0.2122   0.6221   1.0000
   4.000   1.2745   0.01281   0.00614  -0.2099   0.5654   1.0000
   4.250   1.2881   0.01363   0.00668  -0.2071   0.5020   1.0000
   4.500   1.2867   0.01522   0.00755  -0.2014   0.3976   1.0000
   4.750   1.2878   0.01697   0.00850  -0.1967   0.2737   1.0000
   5.000   1.3030   0.01796   0.00922  -0.1945   0.2152   1.0000
   5.250   1.3112   0.01983   0.01045  -0.1914   0.0990   1.0000
   5.500   1.3197   0.02189   0.01182  -0.1884   0.0133   1.0000
   5.750   1.3381   0.02278   0.01277  -0.1866   0.0085   1.0000
   6.000   1.3572   0.02357   0.01369  -0.1849   0.0077   1.0000
   6.250   1.3757   0.02442   0.01470  -0.1832   0.0073   1.0000
   6.500   1.3935   0.02537   0.01582  -0.1814   0.0070   1.0000
   6.750   1.4101   0.02644   0.01706  -0.1793   0.0068   1.0000
   7.000   1.4253   0.02763   0.01843  -0.1771   0.0066   1.0000
   7.250   1.4389   0.02899   0.01996  -0.1747   0.0065   1.0000
   7.500   1.4512   0.03050   0.02162  -0.1721   0.0064   1.0000
   7.750   1.4628   0.03218   0.02345  -0.1694   0.0064   1.0000
   8.000   1.4746   0.03399   0.02542  -0.1668   0.0063   1.0000
   8.250   1.4875   0.03596   0.02754  -0.1644   0.0063   1.0000
   8.500   1.5019   0.03795   0.02971  -0.1623   0.0059   1.0000
   8.750   1.5157   0.03982   0.03194  -0.1603   0.0052   1.0000
   9.000   1.5268   0.04192   0.03420  -0.1582   0.0045   1.0000
   9.250   1.5393   0.04480   0.03733  -0.1561   0.0042   1.0000
   9.500   1.5510   0.04862   0.04150  -0.1539   0.0039   1.0000
   9.750   1.5581   0.05205   0.04528  -0.1512   0.0039   1.0000
  10.000   1.5589   0.05602   0.04965  -0.1479   0.0038   1.0000
  10.250   1.5549   0.05999   0.05400  -0.1443   0.0038   1.0000
  10.500   1.5473   0.06405   0.05842  -0.1407   0.0038   1.0000
  10.750   1.5367   0.06831   0.06301  -0.1372   0.0038   1.0000
  11.000   1.5242   0.07267   0.06770  -0.1341   0.0038   1.0000
  11.250   1.5126   0.07690   0.07220  -0.1317   0.0038   1.0000
  11.500   1.4954   0.08212   0.07774  -0.1297   0.0038   1.0000
  11.750   1.4820   0.08690   0.08276  -0.1286   0.0038   1.0000
  12.000   1.4628   0.09292   0.08906  -0.1283   0.0038   1.0000
  12.250   1.4507   0.09804   0.09440  -0.1288   0.0039   1.0000
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