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EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 59 AIRFOIL (e59-il)
Reynolds number: 50,000
Max Cl/Cd: 43.12 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e59-il-50000.txt
Download as CSV file: xf-e59-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 59 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3947   0.11097   0.10427  -0.0197   1.0000   0.1383
  -7.250  -0.4080   0.11073   0.10415  -0.0186   1.0000   0.1414
  -7.000  -0.4226   0.11100   0.10456  -0.0194   1.0000   0.1426
  -6.750  -0.4055   0.10468   0.09819  -0.0157   1.0000   0.1491
  -6.500  -0.4095   0.10299   0.09659  -0.0155   1.0000   0.1544
  -6.250  -0.4182   0.10307   0.09683  -0.0209   1.0000   0.1572
  -6.000  -0.4071   0.09740   0.09113  -0.0146   1.0000   0.1631
  -5.750  -0.4072   0.09569   0.08950  -0.0172   1.0000   0.1703
  -5.500  -0.4028   0.09212   0.08599  -0.0164   1.0000   0.1750
  -5.250  -0.3968   0.08983   0.08374  -0.0190   1.0000   0.1847
  -5.000  -0.3911   0.08618   0.08013  -0.0173   1.0000   0.1907
  -4.750  -0.3806   0.08332   0.07731  -0.0220   1.0000   0.2018
  -4.500  -0.3674   0.08030   0.07429  -0.0253   1.0000   0.2153
  -4.250  -0.3569   0.07701   0.07102  -0.0255   1.0000   0.2300
  -4.000  -0.3495   0.07381   0.06787  -0.0232   1.0000   0.2464
  -3.750  -0.3327   0.07092   0.06498  -0.0262   1.0000   0.2734
  -3.500  -0.3265   0.06786   0.06194  -0.0225   1.0000   0.2925
  -3.250  -0.3120   0.06489   0.05898  -0.0230   1.0000   0.3203
  -3.000  -0.2999   0.06210   0.05622  -0.0217   1.0000   0.3510
  -2.250  -0.0272   0.04103   0.03323  -0.0880   1.0000   0.2127
  -2.000   0.0595   0.03509   0.02595  -0.1020   1.0000   0.1710
  -1.750   0.1051   0.03220   0.02256  -0.1062   1.0000   0.1660
  -1.500   0.1482   0.03031   0.01997  -0.1095   1.0000   0.1722
  -1.250   0.1819   0.02886   0.01828  -0.1109   1.0000   0.1810
  -1.000   0.2164   0.02766   0.01670  -0.1121   1.0000   0.1900
  -0.750   0.2471   0.02688   0.01572  -0.1127   1.0000   0.2135
  -0.500   0.2764   0.02608   0.01485  -0.1127   1.0000   0.2413
  -0.250   0.3090   0.02516   0.01424  -0.1135   1.0000   0.3182
   0.000   0.3258   0.02265   0.01345  -0.1100   1.0000   1.0000
   0.250   0.3510   0.02311   0.01339  -0.1098   1.0000   1.0000
   0.500   0.3752   0.02360   0.01358  -0.1096   1.0000   1.0000
   0.750   0.3987   0.02412   0.01388  -0.1094   1.0000   1.0000
   1.000   0.4217   0.02468   0.01428  -0.1092   1.0000   1.0000
   1.250   0.4442   0.02529   0.01475  -0.1090   1.0000   1.0000
   1.500   0.4661   0.02594   0.01531  -0.1087   1.0000   1.0000
   1.750   0.4875   0.02664   0.01595  -0.1085   1.0000   1.0000
   2.000   0.5084   0.02741   0.01668  -0.1082   1.0000   1.0000
   2.250   0.5285   0.02824   0.01751  -0.1079   1.0000   1.0000
   2.500   0.5480   0.02915   0.01845  -0.1077   1.0000   1.0000
   2.750   0.5666   0.03015   0.01949  -0.1075   1.0000   1.0000
   3.000   0.5844   0.03126   0.02066  -0.1073   1.0000   1.0000
   3.250   0.6012   0.03251   0.02198  -0.1072   1.0000   1.0000
   3.500   0.6170   0.03388   0.02344  -0.1071   1.0000   1.0000
   3.750   0.6324   0.03539   0.02510  -0.1072   0.9996   1.0000
   4.000   0.6976   0.03730   0.02724  -0.1159   0.9723   1.0000
   4.250   0.7510   0.03845   0.02867  -0.1218   0.9468   1.0000
   4.500   0.8016   0.03916   0.02973  -0.1266   0.9212   1.0000
   4.750   0.8531   0.03943   0.03036  -0.1308   0.8948   1.0000
   5.000   0.9247   0.03763   0.02909  -0.1354   0.8550   1.0000
   5.250   1.0268   0.03075   0.02312  -0.1377   0.7952   1.0000
   5.500   1.0827   0.02511   0.01797  -0.1321   0.6888   1.0000
   5.750   1.1032   0.02827   0.01760  -0.1248   0.2282   1.0000
   6.000   1.1291   0.03115   0.01987  -0.1240   0.1719   1.0000
   6.250   1.1714   0.03355   0.02214  -0.1254   0.1470   1.0000
   6.500   1.2145   0.03620   0.02471  -0.1271   0.1309   1.0000
   6.750   1.2566   0.03945   0.02826  -0.1284   0.1222   1.0000
   7.000   1.2876   0.04274   0.03166  -0.1285   0.1132   1.0000
   7.250   1.3152   0.04648   0.03604  -0.1273   0.1108   1.0000
   7.500   1.3360   0.05036   0.04053  -0.1254   0.1084   1.0000
   7.750   1.3550   0.05410   0.04453  -0.1239   0.1038   1.0000
   8.000   1.3668   0.05860   0.04969  -0.1211   0.1040   1.0000
   8.250   1.3690   0.06386   0.05581  -0.1172   0.1075   1.0000
   8.500   1.3703   0.06941   0.06192  -0.1141   0.1110   1.0000
   8.750   1.3787   0.07546   0.06821  -0.1123   0.1140   1.0000
   9.000   1.3501   0.08097   0.07462  -0.1072   0.1212   1.0000
   9.250   1.3445   0.08725   0.08115  -0.1052   0.1262   1.0000
   9.500   1.2421   0.08963   0.08412  -0.0914   0.1281   1.0000
   9.750   1.1856   0.09330   0.08805  -0.0866   0.1268   1.0000
  10.000   1.1528   0.09939   0.09433  -0.0856   0.1295   1.0000
  10.250   1.0971   0.10713   0.10220  -0.0877   0.1303   1.0000
  10.500   1.0138   0.12495   0.11996  -0.1011   0.1520   1.0000
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