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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 500,000
Max Cl/Cd: 175.11 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e58-il-500000.txt
Download as CSV file: xf-e58-il-500000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.1845   0.09231   0.09000  -0.0726   0.9599   0.0156
  -7.000  -0.1698   0.08946   0.08715  -0.0764   0.9586   0.0157
  -6.750  -0.1854   0.08853   0.08626  -0.0713   0.9508   0.0158
  -6.500  -0.1689   0.08504   0.08277  -0.0754   0.9484   0.0158
  -6.250  -0.1476   0.08096   0.07868  -0.0809   0.9469   0.0159
  -6.000  -0.1228   0.07695   0.07466  -0.0870   0.9458   0.0159
  -5.750  -0.1330   0.07553   0.07327  -0.0836   0.9380   0.0159
  -5.500  -0.1246   0.07030   0.06806  -0.0858   0.9359   0.0167
  -5.250  -0.1003   0.06698   0.06472  -0.0903   0.9345   0.0171
  -5.000  -0.0709   0.06339   0.06111  -0.0963   0.9335   0.0178
  -4.750  -0.0603   0.06082   0.05854  -0.0980   0.9279   0.0183
  -4.500  -0.0279   0.05673   0.05439  -0.1053   0.9245   0.0195
  -4.250   0.0381   0.05083   0.04838  -0.1213   0.9233   0.0229
  -3.250   0.3460   0.02017   0.01555  -0.1858   0.9329   0.0235
  -3.000   0.3922   0.01735   0.01233  -0.1910   0.9332   0.0273
  -2.750   0.4335   0.01635   0.01114  -0.1939   0.9329   0.0303
  -2.500   0.4568   0.01585   0.01051  -0.1930   0.9282   0.0322
  -2.250   0.4939   0.01497   0.00948  -0.1949   0.9258   0.0360
  -2.000   0.5345   0.01410   0.00854  -0.1976   0.9244   0.0391
  -1.750   0.5737   0.01354   0.00795  -0.1999   0.9232   0.0428
  -1.500   0.6159   0.01293   0.00730  -0.2027   0.9221   0.0461
  -1.250   0.6574   0.01240   0.00671  -0.2054   0.9212   0.0486
  -1.000   0.6989   0.01190   0.00622  -0.2081   0.9205   0.0566
  -0.750   0.7398   0.01140   0.00596  -0.2109   0.9201   0.1427
  -0.500   0.7792   0.01107   0.00571  -0.2133   0.9196   0.1740
  -0.250   0.7919   0.01121   0.00588  -0.2101   0.9125   0.1866
   0.000   0.8302   0.01080   0.00556  -0.2122   0.9111   0.2098
   0.500   0.9166   0.00922   0.00505  -0.2190   0.9088   0.6416
   0.750   0.9450   0.00826   0.00470  -0.2185   0.9069   1.0000
   1.000   0.9699   0.00822   0.00464  -0.2177   0.9018   1.0000
   1.250   0.9997   0.00807   0.00448  -0.2179   0.8976   1.0000
   1.500   1.0354   0.00776   0.00417  -0.2193   0.8945   1.0000
   1.750   1.0617   0.00759   0.00402  -0.2187   0.8866   1.0000
   2.000   1.0951   0.00723   0.00365  -0.2195   0.8775   1.0000
   2.250   1.1213   0.00714   0.00358  -0.2189   0.8677   1.0000
   2.500   1.1476   0.00707   0.00352  -0.2183   0.8554   1.0000
   2.750   1.1755   0.00695   0.00341  -0.2180   0.8340   1.0000
   3.000   1.2065   0.00689   0.00316  -0.2182   0.7822   1.0000
   3.250   1.2146   0.00804   0.00344  -0.2135   0.6373   1.0000
   3.500   1.2212   0.00919   0.00402  -0.2090   0.5366   1.0000
   3.750   1.2318   0.01022   0.00457  -0.2056   0.4477   1.0000
   4.000   1.2478   0.01105   0.00507  -0.2034   0.3847   1.0000
   4.250   1.2625   0.01212   0.00565  -0.2010   0.2909   1.0000
   4.500   1.2805   0.01298   0.00621  -0.1993   0.2348   1.0000
   4.750   1.2981   0.01396   0.00679  -0.1976   0.1644   1.0000
   5.000   1.3144   0.01515   0.00751  -0.1957   0.0841   1.0000
   5.250   1.3304   0.01645   0.00842  -0.1936   0.0256   1.0000
   5.500   1.3509   0.01718   0.00918  -0.1920   0.0184   1.0000
   5.750   1.3714   0.01784   0.00995  -0.1906   0.0166   1.0000
   6.000   1.3908   0.01863   0.01083  -0.1889   0.0152   1.0000
   6.250   1.4084   0.01963   0.01193  -0.1868   0.0139   1.0000
   6.500   1.4200   0.02133   0.01379  -0.1836   0.0124   1.0000
   6.750   1.4354   0.02255   0.01511  -0.1813   0.0119   1.0000
   7.000   1.4524   0.02360   0.01629  -0.1792   0.0115   1.0000
   7.250   1.4690   0.02487   0.01767  -0.1771   0.0111   1.0000
   7.500   1.4865   0.02638   0.01931  -0.1751   0.0106   1.0000
   7.750   1.5065   0.02818   0.02125  -0.1736   0.0103   1.0000
   8.000   1.5331   0.03060   0.02385  -0.1733   0.0102   1.0000
   8.250   1.5666   0.03415   0.02769  -0.1740   0.0105   1.0000
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