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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 50,000
Max Cl/Cd: 44.07 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e58-il-50000-n5.txt
Download as CSV file: xf-e58-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3763   0.12612   0.11977  -0.0198   1.0000   0.0637
  -7.250  -0.3847   0.12551   0.11926  -0.0189   1.0000   0.0640
  -7.000  -0.3948   0.12508   0.11893  -0.0180   1.0000   0.0643
  -6.750  -0.3986   0.12402   0.11795  -0.0192   1.0000   0.0645
  -6.500  -0.3775   0.11532   0.10921  -0.0147   1.0000   0.0693
  -6.250  -0.3804   0.11338   0.10734  -0.0133   1.0000   0.0721
  -6.000  -0.3826   0.11158   0.10560  -0.0130   1.0000   0.0750
  -5.750  -0.3843   0.11031   0.10440  -0.0148   1.0000   0.0774
  -5.500  -0.3803   0.10930   0.10347  -0.0207   1.0000   0.0786
  -5.250  -0.3771   0.10484   0.09908  -0.0180   1.0000   0.0799
  -5.000  -0.3730   0.10129   0.09555  -0.0153   1.0000   0.0827
  -4.750  -0.3667   0.09849   0.09278  -0.0160   1.0000   0.0861
  -4.500  -0.3522   0.09604   0.09035  -0.0223   1.0000   0.0913
  -4.250  -0.3311   0.09258   0.08691  -0.0302   1.0000   0.0932
  -4.000  -0.3292   0.08897   0.08334  -0.0262   1.0000   0.0953
  -3.750  -0.3168   0.08572   0.08006  -0.0272   1.0000   0.0990
  -3.500  -0.2581   0.08147   0.07565  -0.0470   0.9995   0.1067
  -3.250  -0.2474   0.07728   0.07150  -0.0446   0.9962   0.1089
  -2.750  -0.1258   0.06376   0.05749  -0.0726   0.9910   0.0671
  -2.250   0.0381   0.04912   0.04190  -0.1074   0.9902   0.0523
  -2.000   0.1023   0.04511   0.03745  -0.1183   0.9893   0.0576
  -1.750   0.1885   0.03943   0.03074  -0.1333   0.9907   0.0603
  -1.500   0.2471   0.03680   0.02736  -0.1409   0.9902   0.0698
  -1.250   0.3001   0.03470   0.02442  -0.1464   0.9896   0.0759
  -1.000   0.3422   0.03417   0.02344  -0.1498   0.9875   0.0917
  -0.750   0.3796   0.03358   0.02258  -0.1519   0.9842   0.0990
  -0.500   0.4170   0.03324   0.02190  -0.1538   0.9812   0.1109
  -0.250   0.4550   0.03307   0.02152  -0.1561   0.9783   0.1413
   0.000   0.4900   0.03305   0.02147  -0.1580   0.9736   0.2065
   0.250   0.5267   0.03321   0.02194  -0.1608   0.9698   0.3124
   0.500   0.5602   0.03324   0.02223  -0.1625   0.9647   0.3945
   0.750   0.5854   0.03238   0.02250  -0.1619   0.9601   0.8084
   1.000   0.6136   0.03283   0.02264  -0.1624   0.9514   1.0000
   1.250   0.6493   0.03344   0.02294  -0.1644   0.9454   1.0000
   1.500   0.6806   0.03390   0.02322  -0.1654   0.9357   1.0000
   1.750   0.7161   0.03430   0.02347  -0.1672   0.9261   1.0000
   2.000   0.7548   0.03464   0.02369  -0.1694   0.9177   1.0000
   2.500   0.8139   0.03535   0.02435  -0.1706   0.8974   1.0000
   2.750   0.8503   0.03564   0.02468  -0.1724   0.8902   1.0000
   3.000   0.8769   0.03599   0.02509  -0.1725   0.8792   1.0000
   3.250   0.9055   0.03626   0.02544  -0.1728   0.8681   1.0000
   3.500   0.9358   0.03640   0.02573  -0.1732   0.8564   1.0000
   3.750   0.9666   0.03639   0.02585  -0.1736   0.8436   1.0000
   4.000   0.9974   0.03624   0.02587  -0.1737   0.8292   1.0000
   4.250   1.0286   0.03594   0.02582  -0.1737   0.8138   1.0000
   4.500   1.0598   0.03559   0.02569  -0.1736   0.7982   1.0000
   4.750   1.0907   0.03519   0.02555  -0.1734   0.7825   1.0000
   5.000   1.1198   0.03475   0.02543  -0.1727   0.7646   1.0000
   5.250   1.1448   0.03440   0.02538  -0.1712   0.7430   1.0000
   5.500   1.1772   0.03350   0.02481  -0.1704   0.7232   1.0000
   5.750   1.2042   0.03305   0.02470  -0.1690   0.6975   1.0000
   6.000   1.2376   0.03231   0.02440  -0.1683   0.6706   1.0000
   6.250   1.2791   0.03107   0.02350  -0.1682   0.6332   1.0000
   6.500   1.3190   0.02993   0.02203  -0.1666   0.5215   1.0000
   6.750   1.3274   0.03149   0.02258  -0.1621   0.3857   1.0000
   7.000   1.3260   0.03395   0.02423  -0.1576   0.2699   1.0000
   7.250   1.3251   0.03725   0.02645  -0.1541   0.1288   1.0000
   7.500   1.3318   0.04079   0.02937  -0.1515   0.0601   1.0000
   7.750   1.3438   0.04335   0.03198  -0.1492   0.0447   1.0000
   8.000   1.3555   0.04572   0.03459  -0.1468   0.0390   1.0000
   8.250   1.3660   0.04810   0.03714  -0.1444   0.0352   1.0000
   8.500   1.3781   0.05047   0.03978  -0.1421   0.0318   1.0000
   8.750   1.3914   0.05284   0.04240  -0.1401   0.0287   1.0000
   9.000   1.4070   0.05555   0.04527  -0.1383   0.0271   1.0000
   9.250   1.4368   0.05880   0.04882  -0.1377   0.0260   1.0000
   9.500   1.4674   0.06270   0.05327  -0.1373   0.0250   1.0000
   9.750   1.4825   0.06675   0.05788  -0.1355   0.0239   1.0000
  10.000   1.4870   0.07077   0.06240  -0.1328   0.0231   1.0000
  10.250   1.4849   0.07494   0.06704  -0.1298   0.0226   1.0000
  10.500   1.4776   0.07930   0.07183  -0.1268   0.0223   1.0000
  10.750   1.4665   0.08387   0.07681  -0.1239   0.0222   1.0000
  11.000   1.4528   0.08864   0.08196  -0.1216   0.0222   1.0000
  11.250   1.4362   0.09379   0.08747  -0.1198   0.0223   1.0000
  11.500   1.4177   0.09926   0.09327  -0.1189   0.0223   1.0000
  11.750   1.3984   0.10513   0.09942  -0.1191   0.0225   1.0000
  12.000   1.3789   0.11140   0.10597  -0.1204   0.0226   1.0000
  12.250   1.3588   0.11835   0.11315  -0.1231   0.0228   1.0000
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