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EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 59 AIRFOIL (e59-il)
Reynolds number: 100,000
Max Cl/Cd: 70.65 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e59-il-100000-n5.txt
Download as CSV file: xf-e59-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4062   0.11867   0.11398  -0.0234   1.0000   0.0326
  -7.750  -0.4106   0.11668   0.11205  -0.0222   1.0000   0.0326
  -7.500  -0.4152   0.11459   0.11002  -0.0211   1.0000   0.0326
  -7.250  -0.4158   0.11141   0.10692  -0.0207   1.0000   0.0327
  -7.000  -0.4117   0.10642   0.10195  -0.0179   1.0000   0.0333
  -6.750  -0.4091   0.10314   0.09869  -0.0162   1.0000   0.0337
  -6.500  -0.4089   0.10058   0.09615  -0.0155   1.0000   0.0340
  -6.250  -0.3967   0.09696   0.09254  -0.0179   0.9976   0.0346
  -6.000  -0.3808   0.09335   0.08893  -0.0210   0.9945   0.0359
  -5.750  -0.3632   0.08970   0.08527  -0.0254   0.9905   0.0374
  -5.500  -0.3400   0.08576   0.08131  -0.0322   0.9866   0.0408
  -5.250  -0.2897   0.08048   0.07594  -0.0524   0.9806   0.0436
  -4.750  -0.2486   0.06928   0.06467  -0.0596   0.9743   0.0319
  -4.500  -0.2286   0.06542   0.06078  -0.0621   0.9714   0.0299
  -4.250  -0.1874   0.05962   0.05487  -0.0720   0.9693   0.0284
  -4.000  -0.1373   0.05306   0.04810  -0.0840   0.9674   0.0277
  -3.750  -0.0764   0.04586   0.04055  -0.0978   0.9664   0.0297
  -3.500   0.0020   0.03700   0.03096  -0.1142   0.9686   0.0306
  -3.250   0.0701   0.03023   0.02305  -0.1256   0.9703   0.0330
  -3.000   0.1107   0.02866   0.02124  -0.1295   0.9690   0.0380
  -2.750   0.1595   0.02598   0.01766  -0.1342   0.9686   0.0431
  -2.500   0.1944   0.02473   0.01619  -0.1363   0.9657   0.0498
  -2.250   0.2305   0.02381   0.01481  -0.1380   0.9628   0.0570
  -2.000   0.2668   0.02280   0.01370  -0.1401   0.9606   0.0633
  -1.750   0.3035   0.02211   0.01281  -0.1419   0.9584   0.0690
  -1.500   0.3392   0.02158   0.01221  -0.1437   0.9563   0.0759
  -1.250   0.3679   0.02130   0.01189  -0.1441   0.9514   0.0873
  -1.000   0.4024   0.02096   0.01152  -0.1457   0.9480   0.0985
  -0.750   0.4395   0.02066   0.01131  -0.1477   0.9453   0.1233
  -0.500   0.4736   0.02037   0.01122  -0.1493   0.9418   0.1985
  -0.250   0.5062   0.01978   0.01134  -0.1509   0.9370   0.4225
   0.000   0.5323   0.01891   0.01143  -0.1500   0.9331   0.7697
   0.500   0.5859   0.01882   0.01123  -0.1494   0.9208   1.0000
   0.750   0.6203   0.01894   0.01124  -0.1508   0.9172   1.0000
   1.000   0.6480   0.01909   0.01134  -0.1509   0.9109   1.0000
   1.250   0.6794   0.01919   0.01139  -0.1516   0.9053   1.0000
   1.500   0.7148   0.01916   0.01134  -0.1530   0.9002   1.0000
   1.750   0.7443   0.01912   0.01131  -0.1531   0.8911   1.0000
   2.000   0.7756   0.01899   0.01122  -0.1535   0.8817   1.0000
   2.250   0.8139   0.01866   0.01094  -0.1551   0.8749   1.0000
   2.500   0.8439   0.01842   0.01078  -0.1550   0.8629   1.0000
   2.750   0.8807   0.01760   0.01003  -0.1555   0.8472   1.0000
   3.000   0.9081   0.01706   0.00955  -0.1543   0.8224   1.0000
   3.250   0.9384   0.01667   0.00927  -0.1539   0.7982   1.0000
   3.500   0.9742   0.01626   0.00895  -0.1547   0.7761   1.0000
   3.750   1.0178   0.01565   0.00835  -0.1565   0.7321   1.0000
   4.000   1.0774   0.01525   0.00756  -0.1613   0.6165   1.0000
   4.250   1.1072   0.01609   0.00777  -0.1611   0.5182   1.0000
   4.500   1.1249   0.01716   0.00838  -0.1590   0.4069   1.0000
   5.000   1.1552   0.01997   0.01007  -0.1548   0.2087   1.0000
   5.250   1.1703   0.02161   0.01110  -0.1529   0.1081   1.0000
   5.500   1.1882   0.02291   0.01222  -0.1512   0.0813   1.0000
   5.750   1.2073   0.02397   0.01331  -0.1496   0.0698   1.0000
   6.000   1.2264   0.02497   0.01444  -0.1480   0.0637   1.0000
   6.250   1.2439   0.02616   0.01574  -0.1462   0.0598   1.0000
   6.500   1.2623   0.02723   0.01705  -0.1444   0.0548   1.0000
   6.750   1.2777   0.02868   0.01853  -0.1425   0.0488   1.0000
   7.000   1.2958   0.02994   0.01997  -0.1408   0.0433   1.0000
   7.250   1.3107   0.03188   0.02190  -0.1389   0.0384   1.0000
   7.500   1.3308   0.03369   0.02395  -0.1375   0.0336   1.0000
   7.750   1.3507   0.03589   0.02619  -0.1364   0.0294   1.0000
   8.000   1.3743   0.03842   0.02898  -0.1357   0.0259   1.0000
   8.250   1.3992   0.04127   0.03233  -0.1348   0.0232   1.0000
   8.500   1.4170   0.04363   0.03503  -0.1332   0.0206   1.0000
   8.750   1.4298   0.04573   0.03734  -0.1313   0.0187   1.0000
   9.000   1.4422   0.04942   0.04130  -0.1295   0.0177   1.0000
   9.250   1.4495   0.05363   0.04598  -0.1267   0.0172   1.0000
   9.500   1.4514   0.05735   0.05023  -0.1231   0.0168   1.0000
   9.750   1.4470   0.06111   0.05449  -0.1189   0.0165   1.0000
  10.000   1.4365   0.06466   0.05848  -0.1141   0.0161   1.0000
  10.250   1.4226   0.06839   0.06260  -0.1096   0.0157   1.0000
  10.500   1.4063   0.07248   0.06705  -0.1057   0.0156   1.0000
  10.750   1.3883   0.07682   0.07171  -0.1025   0.0154   1.0000
  11.000   1.3686   0.08160   0.07680  -0.1003   0.0153   1.0000
  11.250   1.3476   0.08691   0.08237  -0.0992   0.0154   1.0000
  11.500   1.3256   0.09287   0.08857  -0.0994   0.0155   1.0000
  11.750   1.3039   0.09935   0.09527  -0.1011   0.0156   1.0000
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