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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 500,000
Max Cl/Cd: 166.55 at α=1.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e58-il-500000-n5.txt
Download as CSV file: xf-e58-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.000   0.0039   0.07156   0.06900  -0.1196   0.9450   0.0057
  -6.500   0.0060   0.06727   0.06474  -0.1184   0.9348   0.0056
  -6.250   0.0260   0.06339   0.06086  -0.1237   0.9332   0.0056
  -6.000   0.0513   0.05925   0.05671  -0.1305   0.9320   0.0056
  -5.750   0.0698   0.05553   0.05299  -0.1355   0.9289   0.0056
  -5.500   0.0722   0.05295   0.05043  -0.1360   0.9219   0.0062
  -5.250   0.1031   0.04965   0.04710  -0.1427   0.9205   0.0066
  -4.750   0.3255   0.01498   0.01053  -0.2132   0.9227   0.0075
  -4.500   0.3656   0.01329   0.00839  -0.2167   0.9226   0.0085
  -4.250   0.4022   0.01218   0.00699  -0.2191   0.9223   0.0103
  -4.000   0.4343   0.01184   0.00657  -0.2202   0.9215   0.0123
  -3.750   0.4678   0.01138   0.00600  -0.2215   0.9208   0.0147
  -3.500   0.5014   0.01101   0.00556  -0.2229   0.9199   0.0184
  -3.250   0.5354   0.01069   0.00518  -0.2242   0.9190   0.0233
  -3.000   0.5687   0.01053   0.00494  -0.2254   0.9181   0.0271
  -2.750   0.6043   0.01006   0.00440  -0.2272   0.9172   0.0300
  -2.500   0.6232   0.00996   0.00429  -0.2254   0.9121   0.0328
  -2.250   0.6536   0.00966   0.00393  -0.2260   0.9090   0.0342
  -2.000   0.6874   0.00933   0.00354  -0.2272   0.9065   0.0352
  -1.750   0.7228   0.00900   0.00315  -0.2288   0.9042   0.0365
  -1.500   0.7481   0.00888   0.00299  -0.2282   0.8997   0.0366
  -1.250   0.7753   0.00875   0.00282  -0.2280   0.8952   0.0369
  -1.000   0.8069   0.00856   0.00260  -0.2288   0.8919   0.0374
  -0.750   0.8397   0.00839   0.00240  -0.2298   0.8895   0.0392
  -0.500   0.8629   0.00828   0.00244  -0.2289   0.8836   0.0829
  -0.250   0.8931   0.00811   0.00239  -0.2295   0.8789   0.1362
   0.000   0.9212   0.00801   0.00235  -0.2295   0.8726   0.1552
   0.250   0.9492   0.00794   0.00231  -0.2295   0.8659   0.1667
   0.500   0.9773   0.00788   0.00230  -0.2296   0.8596   0.1798
   0.750   1.0059   0.00780   0.00228  -0.2298   0.8515   0.2004
   1.000   1.0339   0.00771   0.00232  -0.2299   0.8423   0.2508
   1.250   1.0667   0.00734   0.00246  -0.2313   0.8321   0.4903
   1.500   1.0972   0.00696   0.00264  -0.2320   0.8151   0.7540
   1.750   1.1159   0.00670   0.00251  -0.2295   0.7755   0.9674
   2.000   1.1297   0.00748   0.00269  -0.2262   0.6728   1.0000
   2.250   1.1397   0.00842   0.00315  -0.2224   0.5958   1.0000
   2.500   1.1533   0.00920   0.00357  -0.2195   0.5236   1.0000
   2.750   1.1681   0.00999   0.00401  -0.2169   0.4553   1.0000
   3.000   1.1881   0.01051   0.00436  -0.2154   0.4202   1.0000
   3.250   1.2092   0.01097   0.00469  -0.2142   0.3902   1.0000
   3.500   1.2255   0.01186   0.00517  -0.2121   0.3097   1.0000
   3.750   1.2421   0.01277   0.00573  -0.2102   0.2473   1.0000
   4.000   1.2584   0.01379   0.00631  -0.2082   0.1672   1.0000
   4.250   1.2785   0.01441   0.00679  -0.2069   0.1368   1.0000
   4.500   1.2950   0.01547   0.00747  -0.2050   0.0679   1.0000
   4.750   1.3122   0.01648   0.00819  -0.2031   0.0177   1.0000
   5.000   1.3329   0.01706   0.00881  -0.2017   0.0115   1.0000
   5.250   1.3539   0.01758   0.00942  -0.2004   0.0090   1.0000
   5.500   1.3744   0.01814   0.01005  -0.1990   0.0076   1.0000
   5.750   1.3937   0.01885   0.01085  -0.1974   0.0065   1.0000
   6.000   1.4119   0.01972   0.01183  -0.1956   0.0059   1.0000
   6.250   1.4308   0.02045   0.01265  -0.1939   0.0054   1.0000
   6.500   1.4492   0.02117   0.01346  -0.1922   0.0047   1.0000
   6.750   1.4669   0.02196   0.01429  -0.1905   0.0042   1.0000
   7.000   1.4802   0.02329   0.01572  -0.1879   0.0039   1.0000
   7.250   1.4948   0.02444   0.01698  -0.1856   0.0037   1.0000
   7.500   1.5089   0.02567   0.01834  -0.1831   0.0035   1.0000
   7.750   1.5223   0.02706   0.01986  -0.1806   0.0033   1.0000
   8.000   1.5352   0.02863   0.02157  -0.1781   0.0031   1.0000
   8.250   1.5486   0.03030   0.02340  -0.1756   0.0030   1.0000
   8.500   1.5627   0.03216   0.02541  -0.1733   0.0029   1.0000
   8.750   1.5772   0.03400   0.02742  -0.1712   0.0027   1.0000
   9.000   1.5909   0.03557   0.02918  -0.1691   0.0026   1.0000
   9.250   1.6023   0.03693   0.03064  -0.1670   0.0024   1.0000
   9.500   1.6119   0.03993   0.03386  -0.1645   0.0023   1.0000
   9.750   1.6224   0.04170   0.03592  -0.1618   0.0021   1.0000
  10.000   1.6295   0.04476   0.03931  -0.1588   0.0020   1.0000
  10.250   1.6318   0.04841   0.04332  -0.1552   0.0020   1.0000
  10.500   1.6280   0.05270   0.04801  -0.1511   0.0019   1.0000
  10.750   1.6194   0.05702   0.05268  -0.1469   0.0019   1.0000
  11.000   1.6084   0.06137   0.05735  -0.1429   0.0018   1.0000
  11.250   1.5920   0.06646   0.06277  -0.1390   0.0018   1.0000
  11.500   1.5741   0.07171   0.06831  -0.1357   0.0018   1.0000
  11.750   1.5542   0.07737   0.07426  -0.1332   0.0017   1.0000
  12.000   1.5347   0.08315   0.08028  -0.1317   0.0017   1.0000
  12.250   1.5135   0.08955   0.08691  -0.1313   0.0017   1.0000
  12.500   1.4935   0.09614   0.09371  -0.1321   0.0018   1.0000
  12.750   1.4704   0.10419   0.10199  -0.1344   0.0017   1.0000
  13.000   1.4502   0.11254   0.11053  -0.1382   0.0017   1.0000
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