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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 50,000
Max Cl/Cd: 46.12 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e58-il-50000.txt
Download as CSV file: xf-e58-il-50000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.3873   0.11614   0.10992  -0.0142   1.0000   0.1248
  -6.500  -0.3991   0.11601   0.10990  -0.0134   1.0000   0.1268
  -6.250  -0.4076   0.11659   0.11060  -0.0168   1.0000   0.1279
  -6.000  -0.3954   0.10977   0.10373  -0.0118   1.0000   0.1315
  -5.750  -0.3927   0.10707   0.10106  -0.0109   1.0000   0.1362
  -5.500  -0.3936   0.10597   0.10004  -0.0139   1.0000   0.1409
  -5.250  -0.3894   0.10312   0.09726  -0.0159   1.0000   0.1432
  -5.000  -0.3840   0.09913   0.09330  -0.0124   1.0000   0.1485
  -4.750  -0.3729   0.09761   0.09180  -0.0197   1.0000   0.1554
  -4.500  -0.3680   0.09350   0.08774  -0.0174   1.0000   0.1586
  -4.250  -0.3590   0.09043   0.08469  -0.0177   1.0000   0.1658
  -4.000  -0.3436   0.08710   0.08138  -0.0223   1.0000   0.1723
  -3.750  -0.3197   0.08406   0.07831  -0.0296   1.0000   0.1846
  -3.500  -0.3120   0.08070   0.07498  -0.0274   1.0000   0.1956
  -3.250  -0.2979   0.07725   0.07155  -0.0283   1.0000   0.2053
  -3.000  -0.2760   0.07381   0.06805  -0.0319   1.0000   0.2197
  -2.750  -0.2369   0.07011   0.06426  -0.0416   1.0000   0.2414
  -2.500  -0.2178   0.06670   0.06085  -0.0426   1.0000   0.2570
  -2.250  -0.1936   0.06358   0.05770  -0.0450   1.0000   0.2792
  -2.000  -0.1614   0.06028   0.05436  -0.0497   1.0000   0.3129
  -1.750  -0.1491   0.05766   0.05179  -0.0474   1.0000   0.3470
  -1.250   0.1691   0.04094   0.03243  -0.1197   1.0000   0.1698
  -1.000   0.2331   0.03745   0.02819  -0.1280   1.0000   0.1572
  -0.750   0.2779   0.03569   0.02600  -0.1320   1.0000   0.1619
  -0.500   0.3184   0.03456   0.02444  -0.1351   1.0000   0.1770
  -0.250   0.3589   0.03352   0.02298  -0.1377   1.0000   0.1904
   0.000   0.3948   0.03272   0.02188  -0.1392   1.0000   0.2213
   0.250   0.4338   0.03160   0.02101  -0.1413   1.0000   0.3560
   0.500   0.4660   0.03114   0.02121  -0.1423   1.0000   0.5226
   0.750   0.4786   0.03007   0.02102  -0.1392   1.0000   1.0000
   1.000   0.5047   0.03083   0.02117  -0.1396   1.0000   1.0000
   1.250   0.5272   0.03164   0.02167  -0.1395   1.0000   1.0000
   1.500   0.5487   0.03249   0.02231  -0.1393   1.0000   1.0000
   1.750   0.5695   0.03341   0.02307  -0.1391   1.0000   1.0000
   2.000   0.5897   0.03440   0.02394  -0.1389   1.0000   1.0000
   2.250   0.6092   0.03545   0.02492  -0.1386   1.0000   1.0000
   2.500   0.6279   0.03659   0.02600  -0.1384   1.0000   1.0000
   2.750   0.6461   0.03780   0.02720  -0.1382   1.0000   1.0000
   3.000   0.6636   0.03911   0.02851  -0.1381   1.0000   1.0000
   3.250   0.6804   0.04053   0.02998  -0.1380   1.0000   1.0000
   3.500   0.6964   0.04205   0.03154  -0.1379   1.0000   1.0000
   3.750   0.7117   0.04371   0.03326  -0.1379   1.0000   1.0000
   4.000   0.7635   0.04588   0.03557  -0.1446   0.9753   1.0000
   4.250   0.8245   0.04732   0.03722  -0.1515   0.9459   1.0000
   4.500   0.8696   0.04827   0.03837  -0.1555   0.9208   1.0000
   4.750   0.9134   0.04890   0.03921  -0.1589   0.8958   1.0000
   5.000   0.9613   0.04899   0.03964  -0.1621   0.8699   1.0000
   5.250   1.0148   0.04823   0.03921  -0.1652   0.8425   1.0000
   5.500   1.0689   0.04677   0.03814  -0.1676   0.8148   1.0000
   5.750   1.1174   0.04531   0.03718  -0.1689   0.7882   1.0000
   6.000   1.1692   0.04313   0.03553  -0.1698   0.7608   1.0000
   6.250   1.2408   0.03756   0.03068  -0.1697   0.7278   1.0000
   6.500   1.3143   0.03057   0.02437  -0.1671   0.6578   1.0000
   6.750   1.3434   0.02913   0.02220  -0.1612   0.4993   1.0000
   7.000   1.3358   0.03140   0.02341  -0.1540   0.3595   1.0000
   7.250   1.3191   0.03580   0.02623  -0.1469   0.1998   1.0000
   7.500   1.3450   0.04075   0.03012  -0.1455   0.1166   1.0000
   7.750   1.3944   0.04412   0.03331  -0.1478   0.0919   1.0000
   8.000   1.4613   0.04925   0.03873  -0.1523   0.0823   1.0000
   8.250   1.4978   0.05389   0.04389  -0.1525   0.0795   1.0000
   8.500   1.5238   0.05848   0.04873  -0.1519   0.0761   1.0000
   8.750   1.5395   0.06393   0.05458  -0.1501   0.0740   1.0000
   9.000   1.5471   0.06851   0.05976  -0.1469   0.0738   1.0000
   9.250   1.5569   0.07423   0.06594  -0.1445   0.0746   1.0000
   9.500   1.5401   0.07733   0.06998  -0.1378   0.0769   1.0000
   9.750   1.5205   0.08239   0.07572  -0.1324   0.0793   1.0000
  10.000   1.5014   0.08734   0.08111  -0.1279   0.0812   1.0000
  10.250   1.4787   0.09195   0.08603  -0.1235   0.0825   1.0000
  10.500   1.4574   0.09675   0.09109  -0.1201   0.0838   1.0000
  10.750   1.4389   0.10206   0.09660  -0.1178   0.0851   1.0000
  11.000   1.4397   0.10849   0.10315  -0.1171   0.0878   1.0000
  11.250   1.4000   0.11355   0.10850  -0.1160   0.0884   1.0000
  11.500   1.3600   0.12037   0.11553  -0.1178   0.0891   1.0000
  11.750   1.3001   0.13301   0.12834  -0.1273   0.0933   1.0000
  12.000   1.2843   0.14286   0.13819  -0.1330   0.0974   1.0000
  12.250   1.2914   0.14865   0.14400  -0.1334   0.0994   1.0000
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