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EPPLER 58 AIRFOIL (e58-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 58 AIRFOIL (e58-il)
Reynolds number: 1,000,000
Max Cl/Cd: 235.69 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e58-il-1000000.txt
Download as CSV file: xf-e58-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.0921   0.08863   0.08680  -0.1013   0.9675   0.0085
  -8.250  -0.0828   0.08597   0.08414  -0.1028   0.9651   0.0088
  -8.000  -0.0684   0.08293   0.08111  -0.1062   0.9632   0.0105
  -7.750  -0.0531   0.07998   0.07816  -0.1102   0.9618   0.0107
  -7.500  -0.0372   0.07670   0.07488  -0.1138   0.9607   0.0107
  -7.250  -0.0202   0.07329   0.07147  -0.1178   0.9597   0.0108
  -7.000   0.0010   0.06962   0.06779  -0.1229   0.9589   0.0108
  -6.250   0.0310   0.05898   0.05717  -0.1317   0.9481   0.0113
  -6.000   0.0602   0.05546   0.05364  -0.1384   0.9474   0.0115
  -5.750   0.0931   0.05188   0.05004  -0.1460   0.9468   0.0119
  -5.500   0.1288   0.04821   0.04634  -0.1543   0.9462   0.0130
  -5.250   0.1734   0.04298   0.04105  -0.1664   0.9454   0.0140
  -5.000   0.3062   0.01222   0.00830  -0.2157   0.9428   0.0117
  -4.750   0.3291   0.01232   0.00842  -0.2148   0.9402   0.0126
  -4.500   0.3604   0.01194   0.00794  -0.2157   0.9390   0.0139
  -4.250   0.3956   0.01124   0.00710  -0.2174   0.9380   0.0153
  -4.000   0.4338   0.00998   0.00564  -0.2200   0.9372   0.0171
  -3.750   0.4675   0.00966   0.00531  -0.2213   0.9360   0.0189
  -3.500   0.4999   0.00949   0.00511  -0.2222   0.9350   0.0214
  -3.250   0.5334   0.00930   0.00488  -0.2234   0.9339   0.0230
  -3.000   0.5703   0.00825   0.00371  -0.2256   0.9328   0.0264
  -2.750   0.5983   0.00801   0.00346  -0.2257   0.9305   0.0290
  -2.500   0.6250   0.00777   0.00319  -0.2255   0.9274   0.0308
  -2.250   0.6550   0.00748   0.00287  -0.2259   0.9250   0.0326
  -2.000   0.6860   0.00728   0.00265  -0.2266   0.9230   0.0341
  -1.750   0.7186   0.00695   0.00226  -0.2276   0.9211   0.0359
  -1.500   0.7507   0.00674   0.00200  -0.2285   0.9191   0.0375
  -1.250   0.7757   0.00664   0.00189  -0.2278   0.9152   0.0390
  -1.000   0.8040   0.00651   0.00175  -0.2278   0.9119   0.0427
  -0.750   0.8353   0.00619   0.00165  -0.2287   0.9089   0.1249
  -0.500   0.8665   0.00605   0.00156  -0.2293   0.9058   0.1535
  -0.250   0.8911   0.00600   0.00155  -0.2286   0.9004   0.1666
   0.000   0.9192   0.00590   0.00150  -0.2286   0.8959   0.1811
   0.250   0.9492   0.00579   0.00145  -0.2291   0.8918   0.2078
   0.500   0.9758   0.00565   0.00151  -0.2289   0.8864   0.2869
   0.750   1.0063   0.00535   0.00157  -0.2297   0.8809   0.4641
   1.000   1.0349   0.00496   0.00170  -0.2300   0.8697   0.7210
   1.250   1.0555   0.00462   0.00174  -0.2281   0.8557   0.9172
   1.500   1.0799   0.00459   0.00171  -0.2272   0.8409   1.0000
   1.750   1.1054   0.00469   0.00172  -0.2265   0.8162   1.0000
   2.000   1.1245   0.00509   0.00180  -0.2244   0.7529   1.0000
   2.250   1.1322   0.00606   0.00221  -0.2199   0.6477   1.0000
   2.500   1.1458   0.00687   0.00261  -0.2169   0.5701   1.0000
   2.750   1.1581   0.00788   0.00308  -0.2139   0.4608   1.0000
   3.000   1.1759   0.00864   0.00348  -0.2120   0.3911   1.0000
   3.250   1.1955   0.00933   0.00382  -0.2105   0.3206   1.0000
   3.500   1.2146   0.01011   0.00424  -0.2089   0.2553   1.0000
   3.750   1.2350   0.01079   0.00462  -0.2076   0.1941   1.0000
   4.000   1.2563   0.01140   0.00503  -0.2065   0.1564   1.0000
   4.250   1.2791   0.01184   0.00536  -0.2056   0.1309   1.0000
   4.500   1.2973   0.01285   0.00598  -0.2039   0.0598   1.0000
   4.750   1.3169   0.01375   0.00660  -0.2024   0.0174   1.0000
   5.000   1.3395   0.01428   0.00716  -0.2013   0.0128   1.0000
   5.250   1.3624   0.01474   0.00769  -0.2003   0.0118   1.0000
   5.500   1.3850   0.01521   0.00822  -0.1992   0.0109   1.0000
   5.750   1.4072   0.01574   0.00879  -0.1981   0.0097   1.0000
   6.000   1.4284   0.01638   0.00948  -0.1969   0.0087   1.0000
   6.250   1.4464   0.01749   0.01071  -0.1949   0.0078   1.0000
   6.500   1.4619   0.01882   0.01217  -0.1925   0.0074   1.0000
   6.750   1.4816   0.01943   0.01286  -0.1909   0.0072   1.0000
   7.000   1.5006   0.02009   0.01358  -0.1892   0.0068   1.0000
   7.250   1.5179   0.02095   0.01451  -0.1872   0.0064   1.0000
   7.500   1.5340   0.02198   0.01562  -0.1850   0.0060   1.0000
   7.750   1.5500   0.02310   0.01683  -0.1828   0.0057   1.0000
   8.000   1.5665   0.02410   0.01790  -0.1808   0.0054   1.0000
   8.250   1.5826   0.02505   0.01891  -0.1789   0.0051   1.0000
   8.500   1.5974   0.02637   0.02032  -0.1767   0.0049   1.0000
   8.750   1.6121   0.02857   0.02264  -0.1746   0.0046   1.0000
   9.000   1.6328   0.03181   0.02610  -0.1736   0.0045   1.0000
   9.250   1.6579   0.03685   0.03151  -0.1736   0.0044   1.0000
   9.500   1.6620   0.03751   0.03234  -0.1694   0.0043   1.0000
   9.750   1.6684   0.03825   0.03321  -0.1658   0.0042   1.0000
  10.000   1.6745   0.04048   0.03566  -0.1624   0.0041   1.0000
  10.250   1.6783   0.04266   0.03806  -0.1586   0.0039   1.0000
  10.500   1.6780   0.04557   0.04123  -0.1546   0.0038   1.0000
  10.750   1.6736   0.04883   0.04475  -0.1501   0.0037   1.0000
  11.000   1.6636   0.05271   0.04894  -0.1453   0.0036   1.0000
  11.250   1.6502   0.05692   0.05344  -0.1406   0.0035   1.0000
  11.500   1.6279   0.06240   0.05923  -0.1359   0.0036   1.0000
  11.750   1.6066   0.06772   0.06483  -0.1320   0.0036   1.0000
  12.000   1.5791   0.07425   0.07165  -0.1288   0.0036   1.0000
  12.250   1.5541   0.08063   0.07826  -0.1271   0.0036   1.0000
  12.500   1.5298   0.08730   0.08513  -0.1268   0.0037   1.0000
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