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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 200,000
Max Cl/Cd: 108.16 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e58-il-200000.txt
Download as CSV file: xf-e58-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.500  -0.4085   0.10751   0.10447  -0.0082   1.0000   0.0255
  -6.250  -0.4142   0.10572   0.10272  -0.0063   1.0000   0.0260
  -6.000  -0.4176   0.10381   0.10084  -0.0052   1.0000   0.0262
  -5.750  -0.4175   0.10148   0.09851  -0.0050   1.0000   0.0268
  -5.500  -0.4159   0.09898   0.09604  -0.0053   1.0000   0.0274
  -5.250  -0.4123   0.09634   0.09342  -0.0062   1.0000   0.0280
  -5.000  -0.4066   0.09349   0.09059  -0.0077   1.0000   0.0288
  -4.750  -0.3980   0.09047   0.08758  -0.0100   1.0000   0.0296
  -4.500  -0.3860   0.08724   0.08436  -0.0132   1.0000   0.0306
  -4.250  -0.3629   0.08346   0.08057  -0.0198   0.9996   0.0322
  -4.000  -0.2767   0.07711   0.07406  -0.0452   0.9945   0.0340
  -3.750  -0.2081   0.06737   0.06423  -0.0621   0.9792   0.0356
  -3.500  -0.1779   0.06416   0.06100  -0.0655   0.9738   0.0379
  -3.250  -0.0662   0.05474   0.05121  -0.0949   0.9722   0.0480
  -3.000  -0.0509   0.05279   0.04931  -0.0937   0.9653   0.0508
  -2.750   0.0339   0.04609   0.04226  -0.1122   0.9644   0.0623
  -2.500   0.0971   0.04202   0.03791  -0.1232   0.9627   0.0755
  -2.000   0.2830   0.02691   0.02058  -0.1541   0.9718   0.0502
  -1.750   0.3209   0.02564   0.01901  -0.1564   0.9677   0.0537
  -1.500   0.3601   0.02467   0.01771  -0.1587   0.9650   0.0567
  -1.250   0.3994   0.02410   0.01692  -0.1610   0.9624   0.0599
  -1.000   0.4430   0.02345   0.01624  -0.1643   0.9602   0.0669
  -0.750   0.4816   0.02321   0.01594  -0.1665   0.9558   0.0726
  -0.500   0.5177   0.02291   0.01568  -0.1683   0.9502   0.0853
  -0.250   0.5617   0.02270   0.01572  -0.1715   0.9469   0.1685
   0.000   0.5940   0.02266   0.01574  -0.1725   0.9400   0.2130
   0.250   0.6339   0.02251   0.01567  -0.1749   0.9351   0.2484
   0.500   0.6831   0.02216   0.01555  -0.1791   0.9323   0.3122
   0.750   0.7061   0.02068   0.01551  -0.1778   0.9234   1.0000
   1.000   0.7546   0.02038   0.01505  -0.1814   0.9197   1.0000
   1.250   0.7828   0.02028   0.01488  -0.1813   0.9108   1.0000
   1.500   0.8243   0.02006   0.01462  -0.1838   0.9075   1.0000
   1.750   0.8533   0.02001   0.01455  -0.1839   0.9002   1.0000
   2.000   0.8914   0.01974   0.01428  -0.1856   0.8957   1.0000
   2.250   0.9380   0.01916   0.01372  -0.1888   0.8929   1.0000
   2.500   0.9661   0.01883   0.01346  -0.1884   0.8829   1.0000
   2.750   1.0114   0.01814   0.01282  -0.1912   0.8799   1.0000
   3.000   1.0372   0.01792   0.01267  -0.1904   0.8701   1.0000
   3.250   1.0840   0.01694   0.01179  -0.1932   0.8665   1.0000
   3.500   1.1169   0.01604   0.01102  -0.1932   0.8547   1.0000
   3.750   1.1489   0.01515   0.01023  -0.1929   0.8421   1.0000
   4.000   1.1795   0.01442   0.00961  -0.1924   0.8273   1.0000
   4.250   1.2047   0.01405   0.00936  -0.1912   0.8070   1.0000
   4.500   1.3001   0.01202   0.00662  -0.2019   0.6446   1.0000
   4.750   1.3130   0.01319   0.00719  -0.1985   0.5439   1.0000
   5.000   1.3180   0.01449   0.00793  -0.1939   0.4500   1.0000
   5.250   1.3231   0.01589   0.00880  -0.1896   0.3602   1.0000
   5.500   1.3286   0.01740   0.00970  -0.1856   0.2642   1.0000
   5.750   1.3374   0.01908   0.01080  -0.1824   0.1633   1.0000
   6.000   1.3429   0.02169   0.01260  -0.1787   0.0501   1.0000
   6.250   1.3589   0.02303   0.01392  -0.1763   0.0352   1.0000
   6.500   1.3750   0.02419   0.01512  -0.1741   0.0294   1.0000
   6.750   1.3865   0.02586   0.01691  -0.1711   0.0272   1.0000
   7.000   1.4005   0.02735   0.01854  -0.1684   0.0260   1.0000
   7.250   1.4159   0.02910   0.02041  -0.1661   0.0249   1.0000
   7.500   1.4359   0.03111   0.02257  -0.1645   0.0241   1.0000
   7.750   1.4580   0.03287   0.02444  -0.1634   0.0223   1.0000
   8.000   1.4853   0.03519   0.02685  -0.1634   0.0207   1.0000
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