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

EPPLER 587 AIRFOIL (e587-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 587 AIRFOIL (e587-il)
Reynolds number: 1,000,000
Max Cl/Cd: 145.92 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e587-il-1000000.txt
Download as CSV file: xf-e587-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 587 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.2550   0.03965   0.03712  -0.1590   0.9305   0.0060
 -11.000  -0.2514   0.03330   0.03033  -0.1689   0.9050   0.0058
 -10.750  -0.2744   0.02896   0.02551  -0.1685   0.8751   0.0055
 -10.500  -0.3240   0.02293   0.01866  -0.1626   0.8514   0.0049
 -10.250  -0.3258   0.02127   0.01675  -0.1597   0.8396   0.0048
 -10.000  -0.3239   0.02007   0.01535  -0.1568   0.8303   0.0048
  -9.750  -0.3183   0.01910   0.01421  -0.1543   0.8220   0.0048
  -9.500  -0.3059   0.01817   0.01314  -0.1528   0.8154   0.0049
  -9.250  -0.2918   0.01724   0.01207  -0.1513   0.8091   0.0049
  -9.000  -0.2745   0.01660   0.01131  -0.1503   0.8036   0.0050
  -8.750  -0.2568   0.01588   0.01051  -0.1492   0.7989   0.0051
  -8.500  -0.2375   0.01527   0.00981  -0.1483   0.7940   0.0053
  -8.250  -0.2176   0.01470   0.00914  -0.1475   0.7892   0.0055
  -8.000  -0.1963   0.01420   0.00855  -0.1469   0.7851   0.0059
  -7.750  -0.1737   0.01377   0.00807  -0.1464   0.7811   0.0062
  -7.500  -0.1514   0.01331   0.00753  -0.1459   0.7771   0.0063
  -7.250  -0.1319   0.01253   0.00665  -0.1452   0.7730   0.0068
  -7.000  -0.1079   0.01213   0.00622  -0.1450   0.7694   0.0071
  -6.750  -0.0833   0.01178   0.00585  -0.1448   0.7660   0.0074
  -6.500  -0.0583   0.01142   0.00545  -0.1447   0.7625   0.0078
  -6.250  -0.0327   0.01112   0.00509  -0.1447   0.7591   0.0082
  -6.000  -0.0063   0.01084   0.00473  -0.1448   0.7553   0.0086
  -5.750   0.0194   0.01043   0.00431  -0.1449   0.7525   0.0095
  -5.500   0.0463   0.01017   0.00402  -0.1450   0.7494   0.0108
  -5.250   0.0733   0.00988   0.00373  -0.1452   0.7462   0.0136
  -5.000   0.1003   0.00951   0.00340  -0.1455   0.7429   0.0256
  -4.750   0.1283   0.00924   0.00318  -0.1460   0.7394   0.0456
  -4.500   0.1561   0.00896   0.00300  -0.1464   0.7367   0.0683
  -4.250   0.1841   0.00863   0.00282  -0.1470   0.7337   0.1043
  -4.000   0.2125   0.00824   0.00262  -0.1477   0.7306   0.1612
  -3.750   0.2416   0.00768   0.00240  -0.1488   0.7275   0.2636
  -3.500   0.2731   0.00666   0.00207  -0.1510   0.7242   0.4838
  -3.250   0.3031   0.00648   0.00207  -0.1518   0.7213   0.5569
  -3.000   0.3323   0.00647   0.00206  -0.1522   0.7185   0.5788
  -2.750   0.3614   0.00648   0.00206  -0.1526   0.7153   0.5935
  -2.500   0.3904   0.00651   0.00208  -0.1529   0.7122   0.6058
  -2.250   0.4197   0.00658   0.00208  -0.1533   0.7090   0.6145
  -2.000   0.4488   0.00664   0.00212  -0.1537   0.7058   0.6216
  -1.750   0.4777   0.00668   0.00213  -0.1540   0.7028   0.6279
  -1.500   0.5062   0.00670   0.00216  -0.1542   0.6995   0.6341
  -1.250   0.5348   0.00678   0.00221  -0.1545   0.6962   0.6414
  -1.000   0.5635   0.00687   0.00225  -0.1548   0.6926   0.6483
  -0.750   0.5919   0.00694   0.00233  -0.1549   0.6892   0.6538
  -0.500   0.6202   0.00697   0.00237  -0.1552   0.6857   0.6578
  -0.250   0.6486   0.00700   0.00237  -0.1554   0.6819   0.6602
   0.000   0.6771   0.00705   0.00237  -0.1557   0.6781   0.6620
   0.250   0.7054   0.00707   0.00237  -0.1560   0.6742   0.6643
   0.500   0.7333   0.00706   0.00238  -0.1562   0.6703   0.6663
   0.750   0.7611   0.00708   0.00240  -0.1563   0.6659   0.6683
   1.000   0.7888   0.00714   0.00243  -0.1565   0.6615   0.6704
   1.250   0.8166   0.00719   0.00247  -0.1566   0.6573   0.6725
   1.500   0.8444   0.00721   0.00250  -0.1568   0.6529   0.6745
   1.750   0.8719   0.00726   0.00253  -0.1569   0.6483   0.6764
   2.000   0.8992   0.00735   0.00258  -0.1570   0.6434   0.6782
   2.250   0.9266   0.00735   0.00261  -0.1571   0.6384   0.6803
   2.500   0.9532   0.00738   0.00265  -0.1570   0.6325   0.6825
   2.750   0.9794   0.00746   0.00272  -0.1569   0.6269   0.6845
   3.000   1.0064   0.00749   0.00278  -0.1569   0.6210   0.6864
   3.250   1.0323   0.00758   0.00286  -0.1567   0.6149   0.6884
   3.500   1.0585   0.00765   0.00294  -0.1566   0.6089   0.6905
   3.750   1.0843   0.00773   0.00301  -0.1564   0.6018   0.6927
   4.000   1.1092   0.00784   0.00311  -0.1560   0.5946   0.6947
   4.250   1.1344   0.00792   0.00320  -0.1557   0.5864   0.6967
   4.500   1.1583   0.00802   0.00330  -0.1551   0.5782   0.6990
   4.750   1.1821   0.00813   0.00342  -0.1545   0.5689   0.7011
   5.000   1.2053   0.00826   0.00355  -0.1538   0.5592   0.7033
   5.250   1.2262   0.00843   0.00369  -0.1526   0.5486   0.7055
   5.500   1.2464   0.00858   0.00384  -0.1513   0.5366   0.7080
   5.750   1.2660   0.00876   0.00400  -0.1498   0.5240   0.7104
   6.000   1.2841   0.00899   0.00419  -0.1481   0.5103   0.7123
   6.250   1.3018   0.00923   0.00442  -0.1463   0.4958   0.7151
   6.500   1.3189   0.00952   0.00468  -0.1445   0.4803   0.7176
   6.750   1.3342   0.00986   0.00498  -0.1423   0.4626   0.7200
   7.000   1.3492   0.01024   0.00531  -0.1401   0.4447   0.7226
   7.250   1.3624   0.01069   0.00569  -0.1377   0.4255   0.7253
   7.500   1.3749   0.01117   0.00610  -0.1352   0.4059   0.7280
   7.750   1.3869   0.01168   0.00655  -0.1326   0.3851   0.7309
   8.000   1.3976   0.01226   0.00706  -0.1299   0.3646   0.7340
   8.250   1.4077   0.01288   0.00763  -0.1272   0.3458   0.7370
   8.500   1.4176   0.01355   0.00823  -0.1245   0.3259   0.7400
   8.750   1.4270   0.01427   0.00889  -0.1218   0.3061   0.7431
   9.000   1.4353   0.01508   0.00962  -0.1191   0.2868   0.7461
   9.250   1.4417   0.01599   0.01047  -0.1162   0.2668   0.7498
   9.500   1.4487   0.01695   0.01136  -0.1135   0.2461   0.7536
   9.750   1.4558   0.01795   0.01230  -0.1109   0.2275   0.7576
  10.000   1.4594   0.01920   0.01344  -0.1080   0.2062   0.7613
  10.250   1.4668   0.02029   0.01449  -0.1057   0.1903   0.7655
  10.500   1.4729   0.02149   0.01565  -0.1034   0.1727   0.7698
  10.750   1.4766   0.02292   0.01699  -0.1009   0.1530   0.7745
  11.000   1.4809   0.02436   0.01836  -0.0987   0.1364   0.7791
  11.250   1.4859   0.02580   0.01976  -0.0966   0.1216   0.7844
  11.500   1.4901   0.02737   0.02127  -0.0945   0.1064   0.7904
  11.750   1.4964   0.02884   0.02272  -0.0928   0.0949   0.7966
  12.000   1.5018   0.03041   0.02428  -0.0911   0.0836   0.8037
  12.250   1.5086   0.03194   0.02580  -0.0896   0.0743   0.8113
  12.500   1.5143   0.03358   0.02745  -0.0881   0.0649   0.8203
  12.750   1.5197   0.03530   0.02918  -0.0866   0.0567   0.8307
  13.000   1.5251   0.03704   0.03095  -0.0852   0.0494   0.8439
  13.250   1.5291   0.03891   0.03285  -0.0838   0.0425   0.8623
  13.500   1.5329   0.04061   0.03467  -0.0822   0.0375   0.9008
  13.750   1.5358   0.04235   0.03650  -0.0806   0.0321   1.0000
  14.000   1.5415   0.04435   0.03851  -0.0797   0.0282   1.0000
  14.250   1.5467   0.04644   0.04061  -0.0789   0.0244   1.0000
  14.500   1.5518   0.04859   0.04278  -0.0782   0.0214   1.0000
  14.750   1.5571   0.05078   0.04499  -0.0775   0.0185   1.0000
  15.000   1.5623   0.05301   0.04726  -0.0769   0.0165   1.0000
  15.250   1.5668   0.05537   0.04966  -0.0765   0.0146   1.0000
  15.500   1.5716   0.05773   0.05206  -0.0760   0.0130   1.0000
  15.750   1.5752   0.06029   0.05467  -0.0757   0.0116   1.0000
  16.000   1.5793   0.06285   0.05728  -0.0755   0.0103   1.0000
  16.250   1.5825   0.06554   0.06002  -0.0753   0.0092   1.0000
  16.500   1.5852   0.06836   0.06291  -0.0752   0.0083   1.0000
  16.750   1.5875   0.07128   0.06588  -0.0752   0.0073   1.0000
  17.000   1.5892   0.07434   0.06900  -0.0754   0.0065   1.0000
  17.250   1.5905   0.07750   0.07224  -0.0756   0.0058   1.0000
  17.500   1.5897   0.08101   0.07582  -0.0759   0.0051   1.0000
  17.750   1.5910   0.08426   0.07914  -0.0764   0.0046   1.0000
  18.000   1.5895   0.08798   0.08294  -0.0770   0.0040   1.0000
  18.250   1.5871   0.09189   0.08693  -0.0777   0.0036   1.0000
  18.500   1.5859   0.09569   0.09081  -0.0785   0.0032   1.0000
  18.750   1.5827   0.09983   0.09503  -0.0795   0.0028   1.0000
  19.000   1.5784   0.10419   0.09948  -0.0807   0.0024   1.0000
  19.250   1.5746   0.10851   0.10390  -0.0820   0.0021   1.0000
<< Back to EPPLER 587 AIRFOIL (e587-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 587 AIRFOIL (e587-il)