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

EPPLER EC 86(-3)-914 AIRFOIL (ec863914-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER EC 86(-3)-914 AIRFOIL (ec863914-il)
Reynolds number: 500,000
Max Cl/Cd: 98.12 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ec863914-il-500000-n5.txt
Download as CSV file: xf-ec863914-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER EC 86(-3)-914 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
  ------ -------- --------- --------- -------- -------- --------
 -16.000  -0.2102   0.15120   0.14841  -0.1062   0.9749   0.0085
 -15.750  -0.2001   0.14859   0.14579  -0.1081   0.9746   0.0088
 -15.500  -0.1903   0.14559   0.14280  -0.1101   0.9743   0.0092
 -15.250  -0.1835   0.14114   0.13835  -0.1126   0.9740   0.0090
  -8.500  -0.3234   0.02575   0.02055  -0.1475   0.9135   0.0074
  -8.250  -0.3017   0.02396   0.01853  -0.1474   0.9126   0.0067
  -8.000  -0.2758   0.02270   0.01703  -0.1480   0.9119   0.0062
  -7.750  -0.2479   0.02174   0.01594  -0.1488   0.9114   0.0061
  -7.500  -0.2558   0.02168   0.01583  -0.1426   0.9047   0.0060
  -7.250  -0.2311   0.02089   0.01496  -0.1427   0.9032   0.0060
  -7.000  -0.2049   0.01982   0.01387  -0.1433   0.9022   0.0061
  -6.750  -0.1756   0.01873   0.01277  -0.1446   0.9014   0.0064
  -6.500  -0.1432   0.01779   0.01180  -0.1466   0.9008   0.0069
  -6.250  -0.1088   0.01699   0.01094  -0.1487   0.9004   0.0069
  -6.000  -0.0735   0.01627   0.01017  -0.1509   0.9000   0.0068
  -5.750  -0.0377   0.01564   0.00948  -0.1531   0.8997   0.0066
  -5.500  -0.0276   0.01555   0.00936  -0.1501   0.8941   0.0065
  -5.250   0.0034   0.01512   0.00889  -0.1513   0.8925   0.0064
  -5.000   0.0370   0.01470   0.00843  -0.1528   0.8916   0.0062
  -4.750   0.0714   0.01431   0.00798  -0.1545   0.8909   0.0061
  -4.500   0.1063   0.01393   0.00757  -0.1563   0.8902   0.0061
  -4.250   0.1414   0.01360   0.00719  -0.1580   0.8897   0.0060
  -4.000   0.1771   0.01327   0.00682  -0.1598   0.8893   0.0060
  -3.750   0.2132   0.01297   0.00648  -0.1616   0.8888   0.0059
  -3.500   0.2496   0.01268   0.00616  -0.1636   0.8884   0.0059
  -3.250   0.2866   0.01241   0.00586  -0.1656   0.8880   0.0060
  -3.000   0.2972   0.01253   0.00596  -0.1623   0.8812   0.0060
  -2.750   0.3321   0.01224   0.00566  -0.1638   0.8798   0.0061
  -2.500   0.3692   0.01191   0.00530  -0.1657   0.8785   0.0066
  -2.250   0.4078   0.01158   0.00497  -0.1679   0.8774   0.0070
  -2.000   0.4606   0.01058   0.00447  -0.1739   0.8767   0.1697
  -1.750   0.4818   0.01029   0.00441  -0.1730   0.8692   0.2248
  -1.500   0.5336   0.00938   0.00400  -0.1788   0.8660   0.3544
  -1.250   0.5725   0.00890   0.00382  -0.1816   0.8615   0.4465
  -1.000   0.6035   0.00858   0.00369  -0.1826   0.8537   0.5060
  -0.750   0.6297   0.00843   0.00376  -0.1822   0.8428   0.5752
  -0.500   0.6560   0.00838   0.00377  -0.1818   0.8320   0.6068
  -0.250   0.6873   0.00827   0.00365  -0.1824   0.8196   0.6232
   0.000   0.7241   0.00811   0.00343  -0.1840   0.8031   0.6351
   0.250   0.7685   0.00796   0.00315  -0.1872   0.7750   0.6445
   0.750   0.8252   0.00841   0.00315  -0.1873   0.7020   0.6562
   1.000   0.8407   0.00881   0.00336  -0.1846   0.6653   0.6605
   1.250   0.8533   0.00926   0.00361  -0.1815   0.6246   0.6652
   1.500   0.8669   0.00976   0.00390  -0.1787   0.5858   0.6703
   1.750   0.8807   0.01028   0.00426  -0.1759   0.5474   0.6756
   2.000   0.8972   0.01079   0.00459  -0.1738   0.5131   0.6815
   2.250   0.9149   0.01126   0.00492  -0.1719   0.4808   0.6861
   2.500   0.9332   0.01173   0.00527  -0.1701   0.4505   0.6903
   2.750   0.9527   0.01219   0.00560  -0.1686   0.4218   0.6950
   3.000   0.9731   0.01264   0.00591  -0.1674   0.3941   0.6992
   3.250   0.9927   0.01309   0.00626  -0.1659   0.3674   0.7024
   3.500   1.0131   0.01353   0.00661  -0.1646   0.3434   0.7057
   3.750   1.0349   0.01393   0.00693  -0.1637   0.3232   0.7086
   4.000   1.0574   0.01430   0.00723  -0.1629   0.3062   0.7106
   4.250   1.0803   0.01467   0.00753  -0.1622   0.2908   0.7119
   4.500   1.1033   0.01502   0.00783  -0.1615   0.2770   0.7130
   4.750   1.1259   0.01537   0.00815  -0.1608   0.2645   0.7139
   5.000   1.1484   0.01572   0.00849  -0.1600   0.2535   0.7150
   5.250   1.1705   0.01610   0.00884  -0.1591   0.2426   0.7161
   5.500   1.1929   0.01646   0.00921  -0.1584   0.2325   0.7172
   5.750   1.2151   0.01683   0.00958  -0.1576   0.2242   0.7183
   6.000   1.2368   0.01725   0.00998  -0.1567   0.2155   0.7193
   6.250   1.2585   0.01766   0.01038  -0.1558   0.2051   0.7205
   6.500   1.2798   0.01810   0.01081  -0.1549   0.1951   0.7216
   6.750   1.3005   0.01860   0.01127  -0.1540   0.1831   0.7228
   7.000   1.3210   0.01913   0.01175  -0.1530   0.1706   0.7240
   7.250   1.3423   0.01958   0.01220  -0.1521   0.1612   0.7255
   7.500   1.3627   0.02010   0.01270  -0.1511   0.1499   0.7269
   7.750   1.3805   0.02084   0.01331  -0.1498   0.1286   0.7281
   8.000   1.3822   0.02299   0.01487  -0.1462   0.0477   0.7290
   8.250   1.3865   0.02498   0.01660  -0.1430   0.0058   0.7298
   8.500   1.4034   0.02579   0.01748  -0.1415   0.0043   0.7309
   8.750   1.4206   0.02658   0.01836  -0.1400   0.0037   0.7319
   9.000   1.4377   0.02737   0.01924  -0.1387   0.0034   0.7330
   9.250   1.4540   0.02823   0.02019  -0.1372   0.0032   0.7342
   9.500   1.4695   0.02917   0.02124  -0.1356   0.0029   0.7354
   9.750   1.4838   0.03020   0.02237  -0.1340   0.0027   0.7367
  10.000   1.4974   0.03132   0.02359  -0.1323   0.0025   0.7380
  10.250   1.5092   0.03257   0.02497  -0.1304   0.0024   0.7393
  10.500   1.5191   0.03402   0.02654  -0.1284   0.0023   0.7406
  10.750   1.5257   0.03577   0.02841  -0.1260   0.0021   0.7420
  11.000   1.5282   0.03790   0.03069  -0.1233   0.0020   0.7433
  11.250   1.5384   0.03930   0.03220  -0.1216   0.0019   0.7445
  11.500   1.5466   0.04091   0.03392  -0.1198   0.0018   0.7457
  11.750   1.5536   0.04267   0.03579  -0.1180   0.0017   0.7469
  12.000   1.5613   0.04440   0.03764  -0.1164   0.0016   0.7482
  12.250   1.5645   0.04664   0.04001  -0.1145   0.0015   0.7494
  12.500   1.5654   0.04919   0.04269  -0.1127   0.0015   0.7507
  12.750   1.5672   0.05177   0.04540  -0.1112   0.0014   0.7520
  13.000   1.5675   0.05461   0.04837  -0.1098   0.0014   0.7534
  13.250   1.5648   0.05792   0.05183  -0.1085   0.0013   0.7547
  13.500   1.5636   0.06119   0.05523  -0.1076   0.0013   0.7560
  13.750   1.5598   0.06496   0.05913  -0.1070   0.0012   0.7573
  14.000   1.5559   0.06893   0.06327  -0.1067   0.0012   0.7585
  14.250   1.5497   0.07337   0.06786  -0.1068   0.0012   0.7596
  14.500   1.5423   0.07823   0.07288  -0.1072   0.0012   0.7606
  14.750   1.5337   0.08347   0.07828  -0.1081   0.0011   0.7617
  15.000   1.5258   0.08881   0.08377  -0.1094   0.0011   0.7627
  15.250   1.5146   0.09494   0.09006  -0.1112   0.0011   0.7638
  15.500   1.5045   0.10108   0.09636  -0.1133   0.0011   0.7649
  15.750   1.4930   0.10764   0.10308  -0.1158   0.0011   0.7660
  16.000   1.4820   0.11423   0.10981  -0.1186   0.0010   0.7671
<< Back to EPPLER EC 86(-3)-914 AIRFOIL (ec863914-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER EC 86(-3)-914 AIRFOIL (ec863914-il)