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

EPPLER 560 AIRFOIL (e560-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 560 AIRFOIL (e560-il)
Reynolds number: 100,000
Max Cl/Cd: 54.55 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e560-il-100000-n5.txt
Download as CSV file: xf-e560-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 560 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.5008   0.08832   0.08269  -0.0627   1.0000   0.0383
 -12.250  -0.5290   0.08065   0.07491  -0.0660   1.0000   0.0385
 -12.000  -0.5480   0.07528   0.06947  -0.0678   1.0000   0.0388
 -11.750  -0.5645   0.07072   0.06485  -0.0690   1.0000   0.0391
 -11.500  -0.5795   0.06679   0.06086  -0.0696   1.0000   0.0395
 -11.250  -0.5954   0.06312   0.05714  -0.0698   1.0000   0.0401
 -11.000  -0.6109   0.05996   0.05396  -0.0693   1.0000   0.0404
 -10.750  -0.6287   0.05705   0.05102  -0.0684   1.0000   0.0407
 -10.500  -0.6512   0.05428   0.04825  -0.0667   1.0000   0.0409
 -10.250  -0.6759   0.05175   0.04572  -0.0648   0.9994   0.0409
 -10.000  -0.6619   0.04790   0.04174  -0.0710   0.9911   0.0422
  -9.750  -0.6422   0.04365   0.03740  -0.0792   0.9826   0.0439
  -9.500  -0.6198   0.03924   0.03282  -0.0876   0.9731   0.0464
  -9.250  -0.5957   0.03585   0.02924  -0.0937   0.9635   0.0498
  -9.000  -0.5695   0.03298   0.02624  -0.0990   0.9548   0.0551
  -8.750  -0.5373   0.03050   0.02366  -0.1042   0.9481   0.0635
  -8.500  -0.5095   0.02858   0.02164  -0.1073   0.9391   0.0737
  -8.250  -0.4725   0.02696   0.01989  -0.1113   0.9339   0.0875
  -8.000  -0.4436   0.02580   0.01863  -0.1132   0.9246   0.1002
  -7.750  -0.4053   0.02474   0.01748  -0.1165   0.9195   0.1154
  -7.500  -0.3753   0.02401   0.01673  -0.1180   0.9102   0.1318
  -7.250  -0.3369   0.02341   0.01607  -0.1208   0.9048   0.1503
  -7.000  -0.3057   0.02293   0.01550  -0.1219   0.8959   0.1636
  -6.750  -0.2685   0.02235   0.01478  -0.1240   0.8899   0.1757
  -6.500  -0.2349   0.02184   0.01408  -0.1253   0.8817   0.1873
  -6.250  -0.1988   0.02145   0.01367  -0.1270   0.8745   0.1975
  -6.000  -0.1631   0.02104   0.01316  -0.1285   0.8667   0.2082
  -5.750  -0.1263   0.02061   0.01254  -0.1303   0.8585   0.2196
  -5.500  -0.0894   0.02033   0.01222  -0.1319   0.8503   0.2294
  -5.250  -0.0516   0.01998   0.01176  -0.1338   0.8415   0.2394
  -5.000  -0.0149   0.01967   0.01127  -0.1355   0.8319   0.2501
  -4.750   0.0240   0.01940   0.01095  -0.1374   0.8230   0.2588
  -4.500   0.0579   0.01912   0.01050  -0.1386   0.8114   0.2687
  -4.250   0.0950   0.01892   0.01024  -0.1402   0.8014   0.2772
  -4.000   0.1304   0.01869   0.00985  -0.1415   0.7901   0.2865
  -3.750   0.1624   0.01856   0.00966  -0.1422   0.7780   0.2947
  -3.500   0.1969   0.01838   0.00934  -0.1433   0.7667   0.3036
  -3.250   0.2304   0.01826   0.00912  -0.1442   0.7554   0.3120
  -3.000   0.2605   0.01815   0.00891  -0.1446   0.7430   0.3203
  -2.750   0.2915   0.01808   0.00876  -0.1451   0.7315   0.3283
  -2.500   0.3238   0.01799   0.00852  -0.1458   0.7208   0.3369
  -2.250   0.3515   0.01796   0.00846  -0.1456   0.7087   0.3444
  -2.000   0.3813   0.01792   0.00828  -0.1459   0.6976   0.3533
  -1.750   0.4111   0.01789   0.00820  -0.1461   0.6873   0.3608
  -1.500   0.4387   0.01789   0.00810  -0.1460   0.6758   0.3698
  -1.250   0.4671   0.01789   0.00808  -0.1460   0.6657   0.3773
  -1.000   0.4957   0.01790   0.00797  -0.1460   0.6554   0.3868
  -0.750   0.5225   0.01793   0.00801  -0.1457   0.6452   0.3942
  -0.250   0.5775   0.01801   0.00799  -0.1454   0.6257   0.4117
   0.000   0.6060   0.01807   0.00795  -0.1454   0.6171   0.4219
   0.250   0.6318   0.01813   0.00804  -0.1449   0.6075   0.4302
   0.500   0.6594   0.01821   0.00806  -0.1448   0.5990   0.4404
   0.750   0.6855   0.01829   0.00816  -0.1444   0.5900   0.4499
   1.000   0.7126   0.01839   0.00823  -0.1442   0.5820   0.4602
   1.250   0.7389   0.01850   0.00834  -0.1438   0.5734   0.4713
   1.500   0.7658   0.01861   0.00845  -0.1435   0.5659   0.4821
   1.750   0.7915   0.01874   0.00861  -0.1431   0.5575   0.4940
   2.000   0.8189   0.01888   0.00870  -0.1429   0.5507   0.5073
   2.250   0.8436   0.01902   0.00894  -0.1423   0.5426   0.5202
   2.500   0.8707   0.01917   0.00907  -0.1421   0.5360   0.5349
   2.750   0.8955   0.01934   0.00932  -0.1415   0.5283   0.5505
   3.000   0.9215   0.01950   0.00952  -0.1411   0.5216   0.5683
   3.250   0.9473   0.01968   0.00974  -0.1407   0.5152   0.5886
   3.500   0.9718   0.01987   0.01002  -0.1400   0.5081   0.6114
   3.750   0.9978   0.02002   0.01023  -0.1395   0.5023   0.6384
   4.000   1.0204   0.02021   0.01058  -0.1385   0.4955   0.6706
   4.250   1.0432   0.02035   0.01086  -0.1374   0.4894   0.7113
   4.500   1.0648   0.02044   0.01107  -0.1359   0.4843   0.7666
   4.750   1.0774   0.02043   0.01134  -0.1326   0.4780   0.8615
   5.000   1.1022   0.02059   0.01151  -0.1320   0.4721   1.0000
   5.250   1.1296   0.02096   0.01181  -0.1320   0.4669   1.0000
   5.500   1.1531   0.02137   0.01225  -0.1315   0.4604   1.0000
   5.750   1.1790   0.02173   0.01257  -0.1312   0.4548   1.0000
   6.000   1.2045   0.02213   0.01295  -0.1309   0.4496   1.0000
   6.250   1.2270   0.02257   0.01346  -0.1302   0.4433   1.0000
   6.500   1.2521   0.02296   0.01383  -0.1298   0.4380   1.0000
   6.750   1.2760   0.02339   0.01427  -0.1293   0.4328   1.0000
   7.000   1.2970   0.02388   0.01485  -0.1283   0.4267   1.0000
   7.250   1.3213   0.02428   0.01524  -0.1277   0.4214   1.0000
   7.500   1.3435   0.02476   0.01575  -0.1269   0.4161   1.0000
   7.750   1.3628   0.02528   0.01639  -0.1257   0.4101   1.0000
   8.000   1.3860   0.02570   0.01680  -0.1250   0.4048   1.0000
   8.250   1.4058   0.02622   0.01740  -0.1238   0.3993   1.0000
   8.500   1.4231   0.02678   0.01807  -0.1222   0.3931   1.0000
   8.750   1.4453   0.02720   0.01850  -0.1214   0.3878   1.0000
   9.000   1.4605   0.02782   0.01923  -0.1195   0.3819   1.0000
   9.250   1.4749   0.02838   0.01988  -0.1175   0.3757   1.0000
   9.500   1.4965   0.02878   0.02027  -0.1166   0.3704   1.0000
   9.750   1.5032   0.02956   0.02123  -0.1135   0.3639   1.0000
  10.000   1.5167   0.03015   0.02189  -0.1114   0.3577   1.0000
  10.250   1.5310   0.03075   0.02254  -0.1095   0.3518   1.0000
  10.500   1.5367   0.03162   0.02357  -0.1066   0.3449   1.0000
  10.750   1.5515   0.03218   0.02415  -0.1048   0.3389   1.0000
  11.000   1.5545   0.03326   0.02541  -0.1018   0.3320   1.0000
  11.250   1.5621   0.03414   0.02637  -0.0994   0.3253   1.0000
  11.500   1.5687   0.03517   0.02749  -0.0970   0.3188   1.0000
  11.750   1.5705   0.03648   0.02894  -0.0944   0.3116   1.0000
  12.000   1.5781   0.03753   0.03003  -0.0924   0.3049   1.0000
  12.250   1.5747   0.03934   0.03203  -0.0898   0.2973   1.0000
  12.500   1.5811   0.04056   0.03327  -0.0880   0.2904   1.0000
  12.750   1.5739   0.04294   0.03586  -0.0857   0.2823   1.0000
  13.000   1.5772   0.04454   0.03746  -0.0841   0.2750   1.0000
  13.250   1.5676   0.04747   0.04061  -0.0825   0.2667   1.0000
  13.500   1.5673   0.04963   0.04279  -0.0812   0.2590   1.0000
  13.750   1.5566   0.05307   0.04640  -0.0801   0.2506   1.0000
  14.000   1.5521   0.05597   0.04936  -0.0793   0.2426   1.0000
  14.250   1.5415   0.05978   0.05330  -0.0789   0.2340   1.0000
  14.500   1.5329   0.06355   0.05715  -0.0787   0.2258   1.0000
  14.750   1.5243   0.06745   0.06114  -0.0787   0.2173   1.0000
  15.000   1.5117   0.07216   0.06596  -0.0791   0.2089   1.0000
  15.250   1.5066   0.07582   0.06961  -0.0794   0.2007   1.0000
  15.500   1.4908   0.08140   0.07536  -0.0805   0.1923   1.0000
  15.750   1.4840   0.08559   0.07957  -0.0812   0.1842   1.0000
  16.000   1.4716   0.09090   0.08497  -0.0825   0.1760   1.0000
  16.250   1.4637   0.09555   0.08966  -0.0837   0.1682   1.0000
<< Back to EPPLER 560 AIRFOIL (e560-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 560 AIRFOIL (e560-il)