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=5


Details Polar file
Airfoil: EPPLER 580 AIRFOIL (e580-il)
Reynolds number: 500,000
Max Cl/Cd: 101.96 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e580-il-500000-n5.txt
Download as CSV file: xf-e580-il-500000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.4216   0.07156   0.06892  -0.0908   0.9875   0.0110
 -12.000  -0.4471   0.06042   0.05756  -0.1015   0.9852   0.0108
 -11.750  -0.4686   0.05226   0.04915  -0.1093   0.9819   0.0109
 -11.500  -0.4835   0.04642   0.04307  -0.1142   0.9763   0.0109
 -11.000  -0.4935   0.03790   0.03409  -0.1213   0.9646   0.0109
 -10.750  -0.4883   0.03437   0.03028  -0.1249   0.9596   0.0111
 -10.500  -0.4912   0.03240   0.02814  -0.1241   0.9486   0.0112
 -10.250  -0.4882   0.03018   0.02570  -0.1242   0.9382   0.0112
 -10.000  -0.4697   0.02810   0.02338  -0.1263   0.9327   0.0113
  -9.750  -0.4554   0.02662   0.02170  -0.1263   0.9236   0.0114
  -9.250  -0.4070   0.02269   0.01739  -0.1294   0.9126   0.0117
  -9.000  -0.3751   0.02136   0.01595  -0.1317   0.9075   0.0119
  -8.750  -0.3375   0.02022   0.01471  -0.1350   0.9035   0.0122
  -8.500  -0.3070   0.01934   0.01375  -0.1367   0.8952   0.0125
  -8.000  -0.2416   0.01763   0.01181  -0.1405   0.8784   0.0131
  -7.750  -0.2121   0.01684   0.01090  -0.1416   0.8684   0.0134
  -7.500  -0.1836   0.01614   0.01009  -0.1425   0.8578   0.0137
  -7.250  -0.1581   0.01553   0.00938  -0.1426   0.8463   0.0139
  -7.000  -0.1328   0.01499   0.00874  -0.1428   0.8348   0.0141
  -6.750  -0.1073   0.01449   0.00814  -0.1429   0.8237   0.0145
  -6.250  -0.0566   0.01365   0.00711  -0.1430   0.8017   0.0149
  -6.000  -0.0314   0.01313   0.00652  -0.1432   0.7911   0.0153
  -5.750  -0.0056   0.01264   0.00594  -0.1434   0.7809   0.0157
  -5.500   0.0207   0.01225   0.00549  -0.1437   0.7706   0.0161
  -5.250   0.0475   0.01195   0.00512  -0.1439   0.7610   0.0168
  -5.000   0.0745   0.01170   0.00479  -0.1441   0.7512   0.0177
  -4.750   0.1019   0.01146   0.00448  -0.1443   0.7417   0.0185
  -4.500   0.1294   0.01126   0.00419  -0.1445   0.7329   0.0191
  -4.250   0.1570   0.01104   0.00390  -0.1448   0.7235   0.0200
  -4.000   0.1849   0.01085   0.00365  -0.1451   0.7149   0.0216
  -3.750   0.2127   0.01071   0.00344  -0.1453   0.7061   0.0237
  -3.500   0.2410   0.01049   0.00324  -0.1457   0.6978   0.0342
  -3.250   0.2701   0.01002   0.00297  -0.1466   0.6892   0.1035
  -3.000   0.3017   0.00914   0.00259  -0.1487   0.6812   0.2599
  -2.750   0.3332   0.00840   0.00238  -0.1506   0.6731   0.4404
  -2.500   0.3619   0.00840   0.00247  -0.1508   0.6652   0.4948
  -2.000   0.4178   0.00864   0.00266  -0.1508   0.6498   0.5357
  -1.750   0.4456   0.00878   0.00268  -0.1508   0.6424   0.5452
  -1.500   0.4734   0.00882   0.00271  -0.1508   0.6352   0.5491
  -1.250   0.5010   0.00888   0.00270  -0.1508   0.6281   0.5511
  -1.000   0.5288   0.00894   0.00270  -0.1509   0.6213   0.5533
  -0.750   0.5565   0.00900   0.00270  -0.1509   0.6143   0.5557
  -0.500   0.5842   0.00907   0.00271  -0.1509   0.6078   0.5580
  -0.250   0.6120   0.00913   0.00272  -0.1510   0.6010   0.5600
   0.250   0.6669   0.00925   0.00278  -0.1510   0.5881   0.5639
   0.500   0.6940   0.00933   0.00283  -0.1509   0.5815   0.5659
   0.750   0.7214   0.00941   0.00288  -0.1509   0.5754   0.5678
   1.000   0.7487   0.00948   0.00293  -0.1509   0.5688   0.5698
   1.250   0.7755   0.00959   0.00299  -0.1508   0.5627   0.5720
   1.500   0.8030   0.00966   0.00305  -0.1508   0.5563   0.5743
   1.750   0.8297   0.00977   0.00312  -0.1506   0.5498   0.5765
   2.000   0.8567   0.00985   0.00320  -0.1506   0.5436   0.5783
   2.250   0.8832   0.00994   0.00329  -0.1504   0.5367   0.5801
   2.500   0.9095   0.01005   0.00339  -0.1502   0.5304   0.5820
   2.750   0.9361   0.01014   0.00350  -0.1501   0.5234   0.5842
   3.000   0.9616   0.01028   0.00361  -0.1497   0.5166   0.5866
   3.250   0.9882   0.01038   0.00372  -0.1496   0.5093   0.5890
   3.500   1.0134   0.01052   0.00385  -0.1492   0.5020   0.5911
   3.750   1.0394   0.01064   0.00396  -0.1489   0.4944   0.5931
   4.000   1.0640   0.01079   0.00412  -0.1484   0.4865   0.5951
   4.250   1.0895   0.01091   0.00427  -0.1481   0.4784   0.5973
   4.500   1.1133   0.01109   0.00444  -0.1475   0.4698   0.5998
   4.750   1.1380   0.01124   0.00460  -0.1470   0.4605   0.6024
   5.000   1.1615   0.01142   0.00479  -0.1463   0.4510   0.6049
   5.250   1.1837   0.01164   0.00498  -0.1454   0.4405   0.6073
   5.500   1.2062   0.01183   0.00517  -0.1445   0.4296   0.6096
   5.750   1.2275   0.01204   0.00539  -0.1433   0.4192   0.6118
   6.000   1.2475   0.01229   0.00564  -0.1420   0.4081   0.6144
   6.250   1.2680   0.01255   0.00590  -0.1408   0.3961   0.6174
   6.500   1.2886   0.01283   0.00617  -0.1396   0.3846   0.6205
   6.750   1.3081   0.01315   0.00648  -0.1382   0.3721   0.6235
   7.000   1.3270   0.01350   0.00682  -0.1368   0.3597   0.6262
   7.250   1.3454   0.01387   0.00718  -0.1353   0.3471   0.6290
   7.750   1.3811   0.01468   0.00798  -0.1322   0.3200   0.6354
   8.000   1.3981   0.01513   0.00841  -0.1305   0.3062   0.6389
   8.250   1.4137   0.01565   0.00891  -0.1287   0.2905   0.6424
   8.500   1.4258   0.01633   0.00952  -0.1263   0.2693   0.6462
   8.750   1.4359   0.01714   0.01023  -0.1237   0.2452   0.6504
   9.000   1.4444   0.01805   0.01104  -0.1210   0.2199   0.6546
   9.250   1.4529   0.01900   0.01189  -0.1184   0.1988   0.6586
   9.500   1.4628   0.01989   0.01275  -0.1160   0.1810   0.6630
   9.750   1.4709   0.02092   0.01372  -0.1136   0.1627   0.6679
  10.250   1.4857   0.02317   0.01590  -0.1088   0.1316   0.6779
  10.500   1.4934   0.02435   0.01707  -0.1066   0.1189   0.6840
  10.750   1.5000   0.02564   0.01835  -0.1045   0.1064   0.6903
  11.000   1.5053   0.02708   0.01978  -0.1023   0.0937   0.6974
  11.250   1.5082   0.02877   0.02144  -0.1001   0.0801   0.7049
  11.500   1.5125   0.03041   0.02309  -0.0982   0.0695   0.7131
  11.750   1.5178   0.03206   0.02478  -0.0965   0.0613   0.7219
  12.000   1.5215   0.03392   0.02666  -0.0949   0.0530   0.7317
  12.250   1.5236   0.03598   0.02875  -0.0933   0.0447   0.7425
  12.500   1.5255   0.03815   0.03096  -0.0919   0.0374   0.7549
  12.750   1.5281   0.04034   0.03322  -0.0906   0.0315   0.7698
  13.000   1.5306   0.04261   0.03557  -0.0895   0.0264   0.7888
  13.250   1.5313   0.04510   0.03816  -0.0885   0.0213   0.8150
  13.500   1.5306   0.04766   0.04086  -0.0873   0.0165   0.8682
  13.750   1.5254   0.05035   0.04367  -0.0857   0.0123   1.0000
  14.000   1.5246   0.05340   0.04676  -0.0852   0.0097   1.0000
  14.250   1.5252   0.05637   0.04980  -0.0849   0.0084   1.0000
  14.500   1.5263   0.05938   0.05288  -0.0847   0.0077   1.0000
  14.750   1.5264   0.06258   0.05617  -0.0846   0.0071   1.0000
  15.000   1.5269   0.06581   0.05949  -0.0847   0.0066   1.0000
  15.250   1.5279   0.06903   0.06281  -0.0848   0.0064   1.0000
  15.500   1.5282   0.07241   0.06629  -0.0851   0.0061   1.0000
  15.750   1.5280   0.07596   0.06994  -0.0856   0.0059   1.0000
  16.000   1.5270   0.07967   0.07375  -0.0861   0.0057   1.0000
  16.250   1.5252   0.08356   0.07774  -0.0868   0.0055   1.0000
  16.500   1.5228   0.08761   0.08189  -0.0876   0.0053   1.0000
  16.750   1.5196   0.09185   0.08624  -0.0886   0.0052   1.0000
  17.000   1.5153   0.09635   0.09085  -0.0898   0.0051   1.0000
  17.250   1.5104   0.10101   0.09562  -0.0911   0.0050   1.0000
<< Back to EPPLER 580 AIRFOIL (e580-il)

Polar data table (+)

Polar graphs


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