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

EPPLER 580 AIRFOIL (e580-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 580 AIRFOIL (e580-il)
Reynolds number: 500,000
Max Cl/Cd: 107.36 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e580-il-500000.txt
Download as CSV file: xf-e580-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 580 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.1605   0.10002   0.09783  -0.0780   0.9842   0.0281
 -11.250  -0.1748   0.09049   0.08832  -0.0841   0.9824   0.0302
 -11.000  -0.2949   0.10252   0.10028  -0.0671   0.9901   0.0260
 -10.750  -0.2850   0.09804   0.09580  -0.0711   0.9881   0.0266
 -10.500  -0.2768   0.09262   0.09038  -0.0762   0.9863   0.0276
  -9.250  -0.2892   0.03206   0.02931  -0.1242   0.9495   0.0322
  -9.000  -0.2775   0.02868   0.02580  -0.1292   0.9461   0.0329
  -8.750  -0.2749   0.02623   0.02321  -0.1296   0.9363   0.0337
  -8.500  -0.2517   0.02533   0.02185  -0.1337   0.9326   0.0375
  -7.750  -0.2451   0.02387   0.01866  -0.1374   0.9206   0.0246
  -7.500  -0.2114   0.02052   0.01520  -0.1394   0.9185   0.0230
  -7.250  -0.1866   0.01910   0.01364  -0.1392   0.9103   0.0229
  -7.000  -0.1503   0.01779   0.01218  -0.1413   0.9059   0.0232
  -6.750  -0.1189   0.01658   0.01086  -0.1423   0.8991   0.0231
  -6.500  -0.0862   0.01553   0.00972  -0.1435   0.8918   0.0231
  -6.250  -0.0560   0.01470   0.00881  -0.1443   0.8829   0.0233
  -6.000  -0.0232   0.01396   0.00798  -0.1457   0.8744   0.0237
  -5.750   0.0036   0.01340   0.00735  -0.1459   0.8633   0.0240
  -5.500   0.0324   0.01290   0.00676  -0.1465   0.8531   0.0243
  -5.000   0.0881   0.01162   0.00532  -0.1478   0.8316   0.0260
  -4.750   0.1164   0.01127   0.00490  -0.1483   0.8212   0.0275
  -4.500   0.1456   0.01098   0.00452  -0.1489   0.8114   0.0293
  -4.250   0.1733   0.01073   0.00417  -0.1491   0.8005   0.0308
  -4.000   0.2017   0.01043   0.00383  -0.1495   0.7905   0.0355
  -3.750   0.2317   0.00986   0.00342  -0.1506   0.7810   0.0953
  -3.500   0.2649   0.00802   0.00277  -0.1542   0.7708   0.4353
  -3.250   0.2937   0.00806   0.00288  -0.1544   0.7614   0.5010
  -3.000   0.3219   0.00821   0.00296  -0.1544   0.7519   0.5218
  -2.750   0.3498   0.00833   0.00301  -0.1543   0.7424   0.5343
  -2.500   0.3776   0.00860   0.00321  -0.1541   0.7337   0.5513
  -2.250   0.4044   0.00887   0.00347  -0.1537   0.7243   0.5663
  -2.000   0.4321   0.00914   0.00365  -0.1535   0.7160   0.5768
  -1.750   0.4586   0.00924   0.00375  -0.1531   0.7073   0.5829
  -1.500   0.4864   0.00931   0.00375  -0.1531   0.6993   0.5861
  -1.250   0.5143   0.00935   0.00371  -0.1532   0.6911   0.5888
  -1.000   0.5424   0.00940   0.00368  -0.1533   0.6833   0.5915
  -0.750   0.5705   0.00945   0.00363  -0.1535   0.6756   0.5937
  -0.500   0.5979   0.00945   0.00361  -0.1535   0.6682   0.5957
  -0.250   0.6252   0.00949   0.00362  -0.1534   0.6606   0.5978
   0.000   0.6528   0.00955   0.00364  -0.1534   0.6536   0.6000
   0.250   0.6803   0.00960   0.00365  -0.1534   0.6463   0.6023
   0.500   0.7081   0.00968   0.00366  -0.1535   0.6396   0.6046
   0.750   0.7357   0.00973   0.00369  -0.1535   0.6324   0.6072
   1.250   0.7905   0.00984   0.00374  -0.1536   0.6189   0.6115
   1.500   0.8176   0.00992   0.00379  -0.1535   0.6123   0.6135
   1.750   0.8447   0.00998   0.00386  -0.1535   0.6057   0.6156
   2.000   0.8715   0.01006   0.00392  -0.1534   0.5989   0.6181
   2.250   0.8987   0.01017   0.00400  -0.1533   0.5925   0.6207
   2.500   0.9255   0.01024   0.00407  -0.1532   0.5855   0.6232
   3.000   0.9788   0.01040   0.00422  -0.1530   0.5720   0.6276
   3.250   1.0049   0.01051   0.00432  -0.1528   0.5652   0.6299
   3.500   1.0309   0.01060   0.00444  -0.1525   0.5582   0.6325
   3.750   1.0567   0.01071   0.00455  -0.1522   0.5510   0.6352
   4.000   1.0826   0.01083   0.00468  -0.1520   0.5439   0.6379
   4.250   1.1082   0.01095   0.00479  -0.1517   0.5363   0.6404
   4.500   1.1334   0.01106   0.00491  -0.1513   0.5289   0.6429
   4.750   1.1579   0.01117   0.00506  -0.1508   0.5206   0.6454
   5.000   1.1823   0.01130   0.00522  -0.1502   0.5126   0.6484
   5.250   1.2061   0.01145   0.00537  -0.1496   0.5037   0.6516
   5.500   1.2301   0.01159   0.00553  -0.1490   0.4948   0.6547
   5.750   1.2529   0.01179   0.00571  -0.1481   0.4859   0.6575
   6.000   1.2760   0.01191   0.00589  -0.1474   0.4760   0.6604
   6.250   1.2980   0.01209   0.00610  -0.1464   0.4663   0.6635
   6.500   1.3183   0.01231   0.00630  -0.1451   0.4560   0.6670
   6.750   1.3388   0.01249   0.00653  -0.1438   0.4450   0.6709
   7.000   1.3581   0.01271   0.00676  -0.1423   0.4341   0.6745
   7.250   1.3766   0.01299   0.00705  -0.1407   0.4231   0.6781
   7.500   1.3951   0.01328   0.00735  -0.1391   0.4108   0.6821
   7.750   1.4140   0.01359   0.00768  -0.1377   0.3987   0.6865
   8.000   1.4316   0.01393   0.00803  -0.1360   0.3856   0.6907
   8.250   1.4478   0.01433   0.00843  -0.1341   0.3706   0.6951
   8.500   1.4616   0.01484   0.00889  -0.1319   0.3524   0.7003
   9.000   1.4876   0.01598   0.00997  -0.1273   0.3151   0.7114
   9.250   1.5011   0.01658   0.01055  -0.1252   0.2984   0.7181
   9.500   1.5127   0.01727   0.01122  -0.1229   0.2801   0.7244
   9.750   1.5226   0.01807   0.01198  -0.1204   0.2611   0.7320
  10.000   1.5316   0.01895   0.01283  -0.1178   0.2436   0.7399
  10.250   1.5406   0.01986   0.01373  -0.1154   0.2255   0.7495
  10.500   1.5479   0.02089   0.01474  -0.1129   0.2069   0.7599
  10.750   1.5520   0.02216   0.01596  -0.1100   0.1870   0.7727
  11.000   1.5561   0.02348   0.01726  -0.1074   0.1685   0.7888
  11.250   1.5598   0.02487   0.01866  -0.1048   0.1513   0.8109
  11.500   1.5630   0.02623   0.02009  -0.1022   0.1374   0.8475
  11.750   1.5614   0.02751   0.02149  -0.0988   0.1250   1.0000
  12.000   1.5641   0.02934   0.02326  -0.0969   0.1107   1.0000
  12.250   1.5667   0.03124   0.02512  -0.0950   0.0984   1.0000
  12.500   1.5687   0.03327   0.02712  -0.0934   0.0872   1.0000
  12.750   1.5699   0.03546   0.02929  -0.0918   0.0769   1.0000
  13.000   1.5701   0.03783   0.03163  -0.0903   0.0669   1.0000
  13.250   1.5696   0.04038   0.03416  -0.0890   0.0569   1.0000
  13.750   1.5624   0.04640   0.04012  -0.0867   0.0360   1.0000
  14.000   1.5542   0.05013   0.04381  -0.0858   0.0256   1.0000
  14.250   1.5464   0.05395   0.04763  -0.0851   0.0185   1.0000
  14.500   1.5411   0.05765   0.05137  -0.0847   0.0149   1.0000
  14.750   1.5394   0.06104   0.05484  -0.0845   0.0132   1.0000
  15.000   1.5359   0.06476   0.05863  -0.0845   0.0120   1.0000
  15.250   1.5337   0.06840   0.06237  -0.0846   0.0113   1.0000
  15.500   1.5324   0.07200   0.06608  -0.0849   0.0108   1.0000
  15.750   1.5302   0.07582   0.07000  -0.0853   0.0104   1.0000
  16.000   1.5271   0.07986   0.07414  -0.0859   0.0101   1.0000
  16.250   1.5222   0.08422   0.07860  -0.0867   0.0097   1.0000
  16.500   1.5160   0.08889   0.08338  -0.0877   0.0095   1.0000
  16.750   1.5083   0.09391   0.08851  -0.0889   0.0092   1.0000
<< Back to EPPLER 580 AIRFOIL (e580-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 580 AIRFOIL (e580-il)