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

EPPLER 431 AIRFOIL (e431-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 431 AIRFOIL (e431-il)
Reynolds number: 1,000,000
Max Cl/Cd: 130.18 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e431-il-1000000-n5.txt
Download as CSV file: xf-e431-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 431 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.2500   0.04219   0.03939  -0.1466   0.7974   0.0034
 -10.250  -0.3320   0.03095   0.02755  -0.1427   0.7786   0.0033
  -9.750  -0.3643   0.02558   0.02160  -0.1345   0.7655   0.0033
  -9.500  -0.3610   0.02357   0.01933  -0.1322   0.7606   0.0033
  -9.250  -0.3535   0.02173   0.01721  -0.1301   0.7559   0.0033
  -9.000  -0.3418   0.02006   0.01529  -0.1285   0.7519   0.0033
  -8.750  -0.3257   0.01878   0.01380  -0.1273   0.7478   0.0032
  -8.500  -0.3083   0.01756   0.01238  -0.1262   0.7440   0.0033
  -8.250  -0.2886   0.01662   0.01127  -0.1253   0.7404   0.0032
  -8.000  -0.2678   0.01574   0.01023  -0.1246   0.7372   0.0033
  -7.750  -0.2456   0.01498   0.00935  -0.1240   0.7339   0.0033
  -7.500  -0.2231   0.01426   0.00850  -0.1234   0.7304   0.0033
  -7.250  -0.1997   0.01368   0.00783  -0.1229   0.7271   0.0033
  -7.000  -0.1764   0.01310   0.00714  -0.1224   0.7240   0.0033
  -6.750  -0.1520   0.01263   0.00660  -0.1221   0.7210   0.0033
  -6.500  -0.1272   0.01216   0.00605  -0.1218   0.7181   0.0034
  -6.250  -0.1022   0.01172   0.00555  -0.1216   0.7149   0.0034
  -6.000  -0.0768   0.01135   0.00512  -0.1214   0.7118   0.0034
  -5.750  -0.0512   0.01103   0.00475  -0.1212   0.7087   0.0035
  -5.500  -0.0253   0.01074   0.00440  -0.1210   0.7058   0.0035
  -5.250   0.0012   0.01042   0.00404  -0.1210   0.7031   0.0036
  -5.000   0.0275   0.01008   0.00366  -0.1209   0.7002   0.0038
  -4.750   0.0543   0.00983   0.00338  -0.1209   0.6970   0.0042
  -4.500   0.0813   0.00965   0.00317  -0.1209   0.6939   0.0044
  -4.250   0.1082   0.00949   0.00296  -0.1210   0.6909   0.0050
  -4.000   0.1359   0.00931   0.00277  -0.1211   0.6882   0.0058
  -3.750   0.1636   0.00914   0.00260  -0.1213   0.6852   0.0071
  -3.500   0.1912   0.00898   0.00245  -0.1214   0.6819   0.0106
  -3.250   0.2186   0.00882   0.00231  -0.1215   0.6787   0.0176
  -3.000   0.2460   0.00868   0.00219  -0.1217   0.6756   0.0284
  -2.750   0.2739   0.00851   0.00208  -0.1219   0.6727   0.0464
  -2.500   0.3015   0.00820   0.00197  -0.1222   0.6696   0.0952
  -2.250   0.3289   0.00784   0.00186  -0.1226   0.6660   0.1716
  -2.000   0.3566   0.00754   0.00177  -0.1230   0.6626   0.2484
  -1.500   0.4130   0.00680   0.00159  -0.1242   0.6559   0.4496
  -1.250   0.4418   0.00626   0.00153  -0.1251   0.6520   0.6147
  -1.000   0.4702   0.00620   0.00157  -0.1253   0.6480   0.6612
  -0.750   0.4983   0.00624   0.00158  -0.1255   0.6439   0.6789
  -0.500   0.5270   0.00625   0.00161  -0.1258   0.6400   0.6915
  -0.250   0.5554   0.00628   0.00164  -0.1260   0.6355   0.7055
   0.000   0.5832   0.00634   0.00170  -0.1261   0.6307   0.7205
   0.250   0.6112   0.00641   0.00176  -0.1262   0.6263   0.7330
   0.500   0.6394   0.00646   0.00181  -0.1263   0.6213   0.7395
   0.750   0.6672   0.00652   0.00183  -0.1265   0.6157   0.7419
   1.000   0.6950   0.00659   0.00185  -0.1266   0.6106   0.7439
   1.250   0.7232   0.00664   0.00188  -0.1269   0.6048   0.7456
   1.500   0.7502   0.00671   0.00192  -0.1268   0.5984   0.7474
   1.750   0.7778   0.00676   0.00197  -0.1270   0.5920   0.7493
   2.000   0.8047   0.00684   0.00203  -0.1269   0.5847   0.7513
   2.250   0.8317   0.00692   0.00209  -0.1269   0.5779   0.7531
   2.500   0.8585   0.00701   0.00215  -0.1269   0.5699   0.7550
   2.750   0.8850   0.00711   0.00223  -0.1268   0.5619   0.7569
   3.000   0.9111   0.00722   0.00231  -0.1267   0.5529   0.7588
   3.250   0.9373   0.00733   0.00239  -0.1266   0.5438   0.7606
   3.500   0.9625   0.00747   0.00249  -0.1262   0.5334   0.7625
   3.750   0.9872   0.00761   0.00261  -0.1258   0.5212   0.7645
   4.000   1.0115   0.00777   0.00274  -0.1253   0.5079   0.7665
   4.250   1.0350   0.00796   0.00289  -0.1247   0.4935   0.7684
   4.500   1.0580   0.00816   0.00305  -0.1240   0.4787   0.7704
   4.750   1.0800   0.00840   0.00323  -0.1231   0.4624   0.7725
   5.000   1.1014   0.00865   0.00342  -0.1221   0.4459   0.7747
   5.250   1.1219   0.00893   0.00363  -0.1209   0.4286   0.7768
   5.500   1.1404   0.00924   0.00386  -0.1193   0.4103   0.7790
   5.750   1.1572   0.00952   0.00410  -0.1174   0.3919   0.7814
   6.000   1.1729   0.00986   0.00437  -0.1153   0.3726   0.7838
   6.250   1.1887   0.01022   0.00468  -0.1132   0.3543   0.7863
   6.500   1.2039   0.01062   0.00500  -0.1111   0.3362   0.7888
   7.000   1.2328   0.01149   0.00574  -0.1068   0.2997   0.7939
   7.250   1.2470   0.01195   0.00614  -0.1047   0.2827   0.7963
   7.500   1.2607   0.01242   0.00657  -0.1025   0.2664   0.7991
   7.750   1.2726   0.01298   0.00706  -0.1001   0.2478   0.8022
   8.000   1.2852   0.01353   0.00757  -0.0978   0.2324   0.8055
   8.250   1.2967   0.01415   0.00813  -0.0955   0.2162   0.8087
   8.750   1.3197   0.01546   0.00937  -0.0911   0.1874   0.8146
   9.000   1.3300   0.01621   0.01008  -0.0889   0.1731   0.8181
   9.250   1.3398   0.01703   0.01086  -0.0866   0.1588   0.8218
   9.500   1.3509   0.01783   0.01163  -0.0847   0.1474   0.8255
   9.750   1.3597   0.01877   0.01253  -0.0825   0.1335   0.8294
  10.000   1.3695   0.01969   0.01344  -0.0806   0.1225   0.8338
  10.250   1.3777   0.02074   0.01446  -0.0785   0.1098   0.8386
  10.500   1.3868   0.02180   0.01548  -0.0767   0.0998   0.8429
  10.750   1.3967   0.02280   0.01650  -0.0750   0.0911   0.8478
  11.000   1.4069   0.02384   0.01755  -0.0734   0.0835   0.8533
  11.500   1.4236   0.02623   0.01995  -0.0701   0.0679   0.8657
  11.750   1.4328   0.02741   0.02115  -0.0687   0.0613   0.8736
  12.000   1.4395   0.02876   0.02253  -0.0670   0.0543   0.8839
  12.500   1.4515   0.03147   0.02533  -0.0635   0.0421   0.9210
  13.000   1.4662   0.03449   0.02843  -0.0613   0.0326   1.0000
  13.500   1.4820   0.03769   0.03166  -0.0594   0.0260   1.0000
  13.750   1.4881   0.03950   0.03348  -0.0584   0.0223   1.0000
  14.000   1.4938   0.04140   0.03540  -0.0575   0.0191   1.0000
  14.250   1.5007   0.04324   0.03727  -0.0568   0.0167   1.0000
  14.500   1.5065   0.04521   0.03927  -0.0561   0.0146   1.0000
  14.750   1.5131   0.04715   0.04126  -0.0555   0.0133   1.0000
  15.000   1.5191   0.04921   0.04335  -0.0550   0.0120   1.0000
  15.250   1.5249   0.05132   0.04552  -0.0546   0.0108   1.0000
  15.500   1.5297   0.05359   0.04784  -0.0542   0.0098   1.0000
  15.750   1.5349   0.05586   0.05016  -0.0539   0.0087   1.0000
  16.000   1.5383   0.05837   0.05273  -0.0536   0.0078   1.0000
  16.250   1.5426   0.06083   0.05525  -0.0535   0.0070   1.0000
  16.500   1.5460   0.06348   0.05796  -0.0535   0.0062   1.0000
  16.750   1.5480   0.06633   0.06087  -0.0535   0.0055   1.0000
  17.000   1.5504   0.06917   0.06377  -0.0536   0.0048   1.0000
  17.250   1.5507   0.07234   0.06701  -0.0539   0.0041   1.0000
  17.500   1.5524   0.07541   0.07017  -0.0543   0.0039   1.0000
  17.750   1.5520   0.07882   0.07365  -0.0547   0.0033   1.0000
  18.000   1.5513   0.08233   0.07723  -0.0554   0.0029   1.0000
  18.250   1.5510   0.08583   0.08082  -0.0561   0.0027   1.0000
  18.500   1.5496   0.08955   0.08463  -0.0569   0.0025   1.0000
  18.750   1.5478   0.09340   0.08857  -0.0579   0.0023   1.0000
  19.000   1.5447   0.09751   0.09277  -0.0591   0.0020   1.0000
  19.250   1.5413   0.10170   0.09704  -0.0604   0.0019   1.0000
<< Back to EPPLER 431 AIRFOIL (e431-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 431 AIRFOIL (e431-il)