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

EPPLER 557 AIRFOIL (e557-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 557 AIRFOIL (e557-il)
Reynolds number: 500,000
Max Cl/Cd: 102.13 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e557-il-500000.txt
Download as CSV file: xf-e557-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 557 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
  ------ -------- --------- --------- -------- -------- --------
 -15.250  -0.5837   0.11042   0.10769  -0.0507   1.0000   0.0091
 -15.000  -0.6225   0.09771   0.09480  -0.0574   1.0000   0.0089
 -14.750  -0.6503   0.08841   0.08533  -0.0624   1.0000   0.0088
 -14.500  -0.6761   0.08022   0.07695  -0.0667   1.0000   0.0088
 -14.250  -0.6989   0.07321   0.06975  -0.0701   1.0000   0.0085
 -14.000  -0.7133   0.06799   0.06439  -0.0725   1.0000   0.0086
 -13.750  -0.7267   0.06326   0.05953  -0.0743   1.0000   0.0087
 -13.500  -0.7422   0.05863   0.05472  -0.0757   1.0000   0.0085
 -13.250  -0.7516   0.05520   0.05118  -0.0763   1.0000   0.0085
 -13.000  -0.7618   0.05192   0.04778  -0.0765   1.0000   0.0086
 -12.750  -0.7707   0.04901   0.04476  -0.0762   1.0000   0.0086
 -12.500  -0.7783   0.04650   0.04216  -0.0757   1.0000   0.0088
 -12.250  -0.7870   0.04415   0.03971  -0.0744   1.0000   0.0087
 -12.000  -0.7962   0.04207   0.03756  -0.0728   1.0000   0.0088
 -11.750  -0.8067   0.04024   0.03565  -0.0707   1.0000   0.0089
 -11.500  -0.8041   0.03811   0.03341  -0.0711   0.9987   0.0089
 -11.250  -0.7833   0.03568   0.03081  -0.0753   0.9953   0.0094
 -11.000  -0.7624   0.03361   0.02861  -0.0785   0.9918   0.0095
 -10.750  -0.7414   0.03168   0.02653  -0.0815   0.9871   0.0097
 -10.500  -0.7234   0.02936   0.02417  -0.0844   0.9828   0.0101
 -10.250  -0.7026   0.02738   0.02212  -0.0872   0.9763   0.0103
 -10.000  -0.6702   0.02556   0.02025  -0.0919   0.9734   0.0109
  -9.750  -0.6460   0.02373   0.01834  -0.0949   0.9665   0.0114
  -9.500  -0.6120   0.02210   0.01660  -0.0991   0.9632   0.0120
  -9.250  -0.5775   0.02030   0.01472  -0.1036   0.9611   0.0129
  -9.000  -0.5529   0.01918   0.01356  -0.1048   0.9540   0.0136
  -8.750  -0.5181   0.01827   0.01257  -0.1074   0.9512   0.0152
  -8.500  -0.4820   0.01703   0.01131  -0.1109   0.9491   0.0176
  -8.250  -0.4445   0.01599   0.01024  -0.1141   0.9476   0.0206
  -8.000  -0.4182   0.01512   0.00934  -0.1149   0.9405   0.0243
  -7.750  -0.3824   0.01425   0.00846  -0.1174   0.9374   0.0298
  -7.500  -0.3438   0.01350   0.00770  -0.1204   0.9352   0.0376
  -7.250  -0.3028   0.01265   0.00690  -0.1239   0.9333   0.0492
  -7.000  -0.2719   0.01201   0.00630  -0.1252   0.9259   0.0645
  -6.750  -0.2316   0.01121   0.00562  -0.1287   0.9218   0.0933
  -6.500  -0.1864   0.01025   0.00488  -0.1334   0.9188   0.1480
  -6.250  -0.1515   0.00889   0.00404  -0.1366   0.9107   0.2706
  -6.000  -0.1105   0.00846   0.00378  -0.1398   0.9046   0.3436
  -5.750  -0.0757   0.00837   0.00367  -0.1413   0.8959   0.3676
  -5.500  -0.0382   0.00835   0.00357  -0.1432   0.8878   0.3838
  -5.250  -0.0069   0.00837   0.00352  -0.1438   0.8772   0.3963
  -5.000   0.0269   0.00843   0.00346  -0.1450   0.8679   0.4066
  -4.750   0.0572   0.00845   0.00343  -0.1454   0.8571   0.4141
  -4.500   0.0870   0.00852   0.00339  -0.1456   0.8465   0.4211
  -4.250   0.1177   0.00854   0.00332  -0.1461   0.8366   0.4269
  -4.000   0.1458   0.00861   0.00334  -0.1461   0.8255   0.4331
  -3.750   0.1745   0.00875   0.00339  -0.1461   0.8151   0.4425
  -3.250   0.2309   0.00885   0.00338  -0.1460   0.7943   0.4523
  -3.000   0.2594   0.00887   0.00330  -0.1461   0.7842   0.4555
  -2.750   0.2875   0.00886   0.00319  -0.1461   0.7740   0.4582
  -2.500   0.3148   0.00882   0.00313  -0.1459   0.7636   0.4609
  -2.250   0.3428   0.00884   0.00309  -0.1458   0.7539   0.4638
  -2.000   0.3702   0.00886   0.00306  -0.1457   0.7436   0.4669
  -1.750   0.3980   0.00888   0.00302  -0.1456   0.7338   0.4700
  -1.500   0.4259   0.00894   0.00297  -0.1456   0.7242   0.4728
  -1.250   0.4532   0.00891   0.00293  -0.1454   0.7142   0.4756
  -1.000   0.4807   0.00894   0.00292  -0.1453   0.7050   0.4786
  -0.750   0.5079   0.00897   0.00293  -0.1451   0.6950   0.4817
  -0.500   0.5353   0.00901   0.00293  -0.1450   0.6857   0.4849
  -0.250   0.5627   0.00907   0.00293  -0.1449   0.6762   0.4881
   0.000   0.5901   0.00912   0.00293  -0.1447   0.6666   0.4909
   0.250   0.6172   0.00915   0.00294  -0.1445   0.6575   0.4941
   0.500   0.6442   0.00918   0.00298  -0.1443   0.6477   0.4975
   0.750   0.6712   0.00926   0.00303  -0.1441   0.6385   0.5011
   1.000   0.6982   0.00933   0.00307  -0.1439   0.6288   0.5046
   1.250   0.7253   0.00942   0.00311  -0.1437   0.6194   0.5078
   1.500   0.7518   0.00946   0.00315  -0.1435   0.6099   0.5112
   1.750   0.7786   0.00952   0.00323  -0.1432   0.6003   0.5147
   2.000   0.8050   0.00963   0.00331  -0.1429   0.5910   0.5186
   2.250   0.8316   0.00972   0.00338  -0.1426   0.5809   0.5227
   2.500   0.8579   0.00980   0.00347  -0.1423   0.5712   0.5266
   2.750   0.8838   0.00990   0.00356  -0.1419   0.5611   0.5305
   3.000   0.9099   0.00999   0.00367  -0.1416   0.5508   0.5346
   3.250   0.9357   0.01014   0.00378  -0.1412   0.5408   0.5390
   3.500   0.9614   0.01024   0.00389  -0.1408   0.5302   0.5432
   3.750   0.9870   0.01034   0.00402  -0.1403   0.5199   0.5478
   4.000   1.0120   0.01051   0.00417  -0.1398   0.5098   0.5530
   4.250   1.0376   0.01064   0.00430  -0.1393   0.4992   0.5581
   4.500   1.0626   0.01076   0.00446  -0.1388   0.4889   0.5630
   5.000   1.1120   0.01109   0.00480  -0.1376   0.4674   0.5743
   5.250   1.1363   0.01125   0.00499  -0.1370   0.4565   0.5800
   5.500   1.1597   0.01146   0.00520  -0.1362   0.4453   0.5868
   5.750   1.1837   0.01163   0.00540  -0.1355   0.4334   0.5936
   6.000   1.2072   0.01182   0.00563  -0.1347   0.4218   0.6007
   6.250   1.2299   0.01206   0.00586  -0.1338   0.4099   0.6085
   6.500   1.2517   0.01230   0.00612  -0.1327   0.3973   0.6162
   6.750   1.2740   0.01254   0.00637  -0.1317   0.3842   0.6254
   7.000   1.2954   0.01278   0.00666  -0.1306   0.3710   0.6348
   7.250   1.3150   0.01306   0.00695  -0.1291   0.3567   0.6451
   7.500   1.3337   0.01336   0.00727  -0.1275   0.3419   0.6565
   7.750   1.3518   0.01370   0.00762  -0.1258   0.3262   0.6683
   8.000   1.3692   0.01408   0.00801  -0.1239   0.3097   0.6814
   8.250   1.3856   0.01452   0.00844  -0.1220   0.2918   0.6963
   8.500   1.4007   0.01502   0.00893  -0.1199   0.2728   0.7136
   8.750   1.4150   0.01556   0.00946  -0.1177   0.2534   0.7336
   9.000   1.4286   0.01612   0.01004  -0.1154   0.2342   0.7573
   9.250   1.4407   0.01673   0.01066  -0.1129   0.2162   0.7869
   9.500   1.4508   0.01734   0.01133  -0.1101   0.1995   0.8293
   9.750   1.4532   0.01771   0.01186  -0.1056   0.1859   1.0000
  10.000   1.4644   0.01855   0.01263  -0.1034   0.1703   1.0000
  10.250   1.4748   0.01944   0.01346  -0.1011   0.1558   1.0000
  10.500   1.4849   0.02037   0.01435  -0.0988   0.1430   1.0000
  10.750   1.4944   0.02136   0.01530  -0.0966   0.1313   1.0000
  11.000   1.5027   0.02245   0.01635  -0.0944   0.1205   1.0000
  11.250   1.5104   0.02361   0.01749  -0.0922   0.1104   1.0000
  11.500   1.5192   0.02475   0.01863  -0.0902   0.1015   1.0000
  11.750   1.5260   0.02607   0.01995  -0.0882   0.0933   1.0000
  12.000   1.5313   0.02755   0.02140  -0.0862   0.0854   1.0000
  12.250   1.5387   0.02894   0.02283  -0.0845   0.0788   1.0000
  12.500   1.5421   0.03070   0.02458  -0.0826   0.0723   1.0000
  12.750   1.5485   0.03228   0.02620  -0.0811   0.0668   1.0000
  13.000   1.5510   0.03426   0.02820  -0.0795   0.0614   1.0000
  13.250   1.5557   0.03614   0.03012  -0.0783   0.0566   1.0000
  13.500   1.5565   0.03844   0.03244  -0.0770   0.0523   1.0000
  13.750   1.5606   0.04052   0.03458  -0.0760   0.0483   1.0000
  14.000   1.5585   0.04331   0.03739  -0.0750   0.0446   1.0000
  14.250   1.5630   0.04552   0.03968  -0.0743   0.0417   1.0000
  14.500   1.5619   0.04841   0.04260  -0.0737   0.0388   1.0000
  14.750   1.5613   0.05136   0.04563  -0.0732   0.0363   1.0000
  15.000   1.5624   0.05421   0.04855  -0.0730   0.0339   1.0000
  15.250   1.5582   0.05780   0.05220  -0.0729   0.0319   1.0000
  15.500   1.5567   0.06119   0.05568  -0.0730   0.0301   1.0000
  15.750   1.5565   0.06449   0.05908  -0.0733   0.0284   1.0000
  16.000   1.5520   0.06849   0.06313  -0.0738   0.0269   1.0000
  16.250   1.5448   0.07296   0.06769  -0.0746   0.0256   1.0000
  16.500   1.5445   0.07657   0.07142  -0.0753   0.0243   1.0000
  16.750   1.5414   0.08068   0.07561  -0.0762   0.0231   1.0000
  17.000   1.5342   0.08550   0.08051  -0.0776   0.0221   1.0000
  17.250   1.5249   0.09077   0.08587  -0.0792   0.0212   1.0000
  17.500   1.5232   0.09492   0.09014  -0.0805   0.0202   1.0000
  17.750   1.5193   0.09949   0.09481  -0.0821   0.0193   1.0000
  18.000   1.5131   0.10451   0.09992  -0.0840   0.0186   1.0000
  18.250   1.5031   0.11023   0.10573  -0.0863   0.0179   1.0000
<< Back to EPPLER 557 AIRFOIL (e557-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 557 AIRFOIL (e557-il)