Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

EPPLER 542 AIRFOIL (e542-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 542 AIRFOIL (e542-il)
Reynolds number: 500,000
Max Cl/Cd: 78.54 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e542-il-500000-n5.txt
Download as CSV file: xf-e542-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 542 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
  ------ -------- --------- --------- -------- -------- --------
 -14.250  -0.6197   0.07926   0.07660  -0.0629   1.0000   0.0035
 -14.000  -0.6521   0.06854   0.06570  -0.0697   1.0000   0.0034
 -13.750  -0.6795   0.06125   0.05821  -0.0727   1.0000   0.0034
 -13.500  -0.7014   0.05609   0.05290  -0.0735   1.0000   0.0034
 -13.250  -0.7260   0.05106   0.04760  -0.0728   1.0000   0.0035
 -13.000  -0.7451   0.04724   0.04358  -0.0712   1.0000   0.0034
 -12.500  -0.7732   0.04108   0.03698  -0.0663   1.0000   0.0035
 -12.250  -0.7850   0.03845   0.03413  -0.0632   1.0000   0.0035
 -11.750  -0.7769   0.03330   0.02853  -0.0617   0.9882   0.0036
 -11.500  -0.7637   0.03179   0.02692  -0.0615   0.9674   0.0038
 -11.250  -0.7368   0.02987   0.02480  -0.0642   0.9523   0.0039
 -10.750  -0.6603   0.02592   0.02035  -0.0732   0.9325   0.0041
 -10.500  -0.6160   0.02427   0.01847  -0.0785   0.9192   0.0042
 -10.250  -0.5755   0.02307   0.01705  -0.0826   0.9007   0.0045
 -10.000  -0.5468   0.02197   0.01571  -0.0840   0.8790   0.0045
  -9.750  -0.5256   0.02133   0.01490  -0.0837   0.8605   0.0046
  -9.250  -0.4910   0.01988   0.01315  -0.0812   0.8313   0.0048
  -9.000  -0.4705   0.01965   0.01280  -0.0804   0.8195   0.0053
  -8.750  -0.4547   0.01893   0.01198  -0.0789   0.8085   0.0055
  -8.500  -0.4390   0.01827   0.01122  -0.0773   0.7976   0.0056
  -8.250  -0.4205   0.01787   0.01076  -0.0761   0.7880   0.0059
  -8.000  -0.4033   0.01739   0.01018  -0.0747   0.7787   0.0062
  -7.750  -0.3858   0.01690   0.00962  -0.0733   0.7704   0.0063
  -7.500  -0.3683   0.01645   0.00909  -0.0719   0.7625   0.0065
  -7.250  -0.3503   0.01601   0.00859  -0.0706   0.7552   0.0069
  -7.000  -0.3321   0.01563   0.00813  -0.0692   0.7475   0.0070
  -6.750  -0.3144   0.01522   0.00765  -0.0678   0.7406   0.0076
  -6.500  -0.2962   0.01485   0.00723  -0.0664   0.7333   0.0084
  -6.250  -0.2780   0.01451   0.00683  -0.0650   0.7270   0.0097
  -6.000  -0.2597   0.01417   0.00646  -0.0636   0.7207   0.0116
  -5.750  -0.2425   0.01381   0.00609  -0.0620   0.7147   0.0166
  -5.500  -0.2246   0.01349   0.00577  -0.0606   0.7093   0.0229
  -5.250  -0.2077   0.01315   0.00546  -0.0589   0.7033   0.0344
  -5.000  -0.1938   0.01272   0.00513  -0.0568   0.6977   0.0595
  -4.750  -0.1816   0.01226   0.00481  -0.0544   0.6923   0.0937
  -4.500  -0.1729   0.01175   0.00448  -0.0513   0.6868   0.1390
  -4.250  -0.1712   0.01127   0.00420  -0.0468   0.6819   0.1936
  -4.000  -0.1689   0.01069   0.00390  -0.0424   0.6774   0.2680
  -3.750  -0.1685   0.00991   0.00352  -0.0378   0.6726   0.3685
  -3.500  -0.1738   0.00876   0.00296  -0.0322   0.6678   0.5144
  -3.250  -0.1683   0.00816   0.00313  -0.0278   0.6635   0.6926
  -3.000  -0.1412   0.00837   0.00331  -0.0276   0.6593   0.7223
  -2.750  -0.1137   0.00861   0.00349  -0.0274   0.6551   0.7388
  -2.500  -0.0858   0.00894   0.00377  -0.0273   0.6512   0.7499
  -2.250  -0.0589   0.00926   0.00398  -0.0270   0.6476   0.7607
  -2.000  -0.0286   0.00981   0.00456  -0.0269   0.6439   0.7662
  -1.750  -0.0006   0.00993   0.00464  -0.0270   0.6398   0.7690
  -1.500   0.0267   0.00989   0.00453  -0.0272   0.6360   0.7704
  -1.250   0.0539   0.00985   0.00441  -0.0274   0.6326   0.7718
  -1.000   0.0813   0.00980   0.00428  -0.0276   0.6294   0.7731
  -0.750   0.1088   0.00973   0.00417  -0.0279   0.6257   0.7745
  -0.500   0.1363   0.00967   0.00405  -0.0282   0.6219   0.7756
  -0.250   0.1640   0.00963   0.00395  -0.0285   0.6186   0.7764
   0.000   0.1916   0.00963   0.00389  -0.0287   0.6153   0.7770
   0.250   0.2195   0.00962   0.00387  -0.0289   0.6120   0.7776
   0.500   0.2474   0.00961   0.00386  -0.0292   0.6084   0.7783
   0.750   0.2751   0.00961   0.00384  -0.0294   0.6048   0.7788
   1.000   0.3027   0.00963   0.00382  -0.0296   0.6014   0.7794
   1.250   0.3302   0.00965   0.00381  -0.0298   0.5981   0.7800
   1.500   0.3580   0.00964   0.00382  -0.0301   0.5943   0.7806
   1.750   0.3854   0.00965   0.00382  -0.0303   0.5900   0.7812
   2.000   0.4126   0.00967   0.00382  -0.0304   0.5855   0.7818
   2.250   0.4397   0.00970   0.00383  -0.0306   0.5810   0.7826
   2.500   0.4670   0.00971   0.00386  -0.0307   0.5761   0.7834
   2.750   0.4940   0.00973   0.00389  -0.0309   0.5713   0.7842
   3.000   0.5208   0.00977   0.00390  -0.0310   0.5667   0.7849
   3.250   0.5479   0.00978   0.00394  -0.0311   0.5614   0.7857
   3.500   0.5746   0.00980   0.00398  -0.0312   0.5557   0.7864
   3.750   0.6008   0.00986   0.00400  -0.0312   0.5503   0.7871
   4.000   0.6276   0.00987   0.00406  -0.0313   0.5441   0.7879
   4.250   0.6535   0.00992   0.00411  -0.0312   0.5372   0.7886
   4.500   0.6796   0.00996   0.00417  -0.0312   0.5303   0.7894
   4.750   0.7051   0.01001   0.00423  -0.0310   0.5220   0.7901
   5.000   0.7305   0.01008   0.00431  -0.0309   0.5136   0.7908
   5.250   0.7546   0.01017   0.00438  -0.0305   0.5034   0.7916
   5.500   0.7792   0.01025   0.00448  -0.0302   0.4917   0.7923
   5.750   0.8020   0.01036   0.00459  -0.0295   0.4780   0.7931
   6.000   0.8233   0.01050   0.00472  -0.0285   0.4612   0.7938
   6.250   0.8420   0.01072   0.00488  -0.0270   0.4399   0.7944
   6.500   0.8583   0.01100   0.00510  -0.0252   0.4131   0.7951
   6.750   0.8718   0.01135   0.00535  -0.0228   0.3851   0.7959
   7.000   0.8817   0.01174   0.00564  -0.0197   0.3580   0.7968
   7.250   0.8901   0.01219   0.00598  -0.0164   0.3308   0.7977
   7.500   0.9002   0.01268   0.00639  -0.0136   0.3049   0.7987
   7.750   0.9103   0.01321   0.00684  -0.0109   0.2819   0.7996
   8.000   0.9205   0.01374   0.00731  -0.0082   0.2595   0.8006
   8.250   0.9288   0.01437   0.00785  -0.0054   0.2367   0.8017
   8.500   0.9381   0.01498   0.00840  -0.0029   0.2160   0.8027
   8.750   0.9448   0.01573   0.00905  -0.0001   0.1941   0.8039
   9.000   0.9515   0.01652   0.00976   0.0026   0.1726   0.8052
   9.250   0.9583   0.01735   0.01052   0.0050   0.1532   0.8064
   9.500   0.9651   0.01825   0.01134   0.0073   0.1359   0.8074
   9.750   0.9747   0.01906   0.01213   0.0092   0.1237   0.8082
  10.000   0.9824   0.01997   0.01302   0.0111   0.1102   0.8091
  10.250   0.9910   0.02089   0.01392   0.0129   0.0993   0.8100
  10.500   0.9983   0.02191   0.01493   0.0147   0.0875   0.8110
  10.750   1.0066   0.02293   0.01596   0.0163   0.0786   0.8119
  11.000   1.0142   0.02404   0.01706   0.0179   0.0698   0.8129
  11.250   1.0218   0.02519   0.01821   0.0193   0.0617   0.8139
  11.500   1.0302   0.02633   0.01936   0.0206   0.0547   0.8149
  11.750   1.0361   0.02766   0.02067   0.0220   0.0464   0.8159
  12.000   1.0438   0.02893   0.02196   0.0231   0.0408   0.8170
  12.250   1.0500   0.03033   0.02336   0.0243   0.0347   0.8180
  12.500   1.0562   0.03177   0.02481   0.0253   0.0294   0.8191
  12.750   1.0634   0.03319   0.02625   0.0262   0.0262   0.8201
  13.000   1.0701   0.03470   0.02778   0.0270   0.0224   0.8213
  13.250   1.0757   0.03633   0.02943   0.0278   0.0186   0.8225
  13.500   1.0828   0.03789   0.03103   0.0284   0.0164   0.8235
  14.000   1.0954   0.04126   0.03450   0.0295   0.0126   0.8255
  14.250   1.1016   0.04301   0.03630   0.0299   0.0110   0.8265
  14.500   1.1076   0.04482   0.03818   0.0302   0.0099   0.8275
  14.750   1.1137   0.04667   0.04011   0.0305   0.0091   0.8285
  15.000   1.1194   0.04861   0.04213   0.0306   0.0080   0.8296
  15.250   1.1244   0.05068   0.04427   0.0307   0.0075   0.8307
  15.500   1.1287   0.05285   0.04652   0.0307   0.0060   0.8318
  15.750   1.1329   0.05512   0.04886   0.0306   0.0056   0.8329
  16.000   1.1364   0.05752   0.05135   0.0304   0.0049   0.8341
  16.250   1.1400   0.05996   0.05387   0.0301   0.0045   0.8352
  16.500   1.1415   0.06269   0.05668   0.0296   0.0040   0.8363
  16.750   1.1435   0.06547   0.05955   0.0291   0.0035   0.8374
  17.000   1.1435   0.06855   0.06273   0.0284   0.0031   0.8384
  17.250   1.1435   0.07173   0.06600   0.0276   0.0029   0.8394
  17.750   1.1401   0.07871   0.07320   0.0255   0.0022   0.8415
  18.000   1.1393   0.08222   0.07683   0.0242   0.0022   0.8426
  18.250   1.1376   0.08594   0.08067   0.0228   0.0022   0.8437
  18.500   1.1337   0.09008   0.08494   0.0212   0.0021   0.8447
  18.750   1.1262   0.09486   0.08983   0.0192   0.0018   0.8458
  19.000   1.1193   0.09968   0.09478   0.0171   0.0017   0.8470
<< Back to EPPLER 542 AIRFOIL (e542-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 542 AIRFOIL (e542-il)