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EPPLER 857 AIRFOIL (e857-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 857 AIRFOIL (e857-il)
Reynolds number: 50,000
Max Cl/Cd: 3.79 at α=10.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e857-il-50000.txt
Download as CSV file: xf-e857-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 857 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3196   0.13724   0.13182  -0.0201   1.0000   0.2974
  -9.500  -0.3007   0.13331   0.12794  -0.0176   1.0000   0.3014
  -9.250  -0.5544   0.09128   0.08613  -0.0469   1.0000   0.1326
  -9.000  -0.5808   0.08840   0.08335  -0.0454   1.0000   0.1317
  -8.750  -0.6130   0.08340   0.07835  -0.0477   0.9958   0.1297
  -8.500  -0.6516   0.07403   0.06865  -0.0560   0.9845   0.1264
  -8.250  -0.6643   0.06677   0.06086  -0.0618   0.9729   0.1266
  -8.000  -0.6598   0.06115   0.05463  -0.0660   0.9615   0.1294
  -7.750  -0.6471   0.05640   0.04919  -0.0693   0.9500   0.1342
  -7.500  -0.6178   0.05415   0.04689  -0.0713   0.9381   0.1436
  -7.250  -0.5929   0.05145   0.04397  -0.0731   0.9263   0.1533
  -7.000  -0.5662   0.04911   0.04137  -0.0748   0.9152   0.1665
  -6.750  -0.5273   0.04724   0.03950  -0.0777   0.9054   0.1861
  -6.500  -0.5053   0.04585   0.03812  -0.0778   0.8931   0.2047
  -6.250  -0.4812   0.04449   0.03662  -0.0784   0.8819   0.2290
  -6.000  -0.4407   0.04394   0.03652  -0.0802   0.8725   0.2658
  -5.750  -0.4258   0.04417   0.03693  -0.0783   0.8599   0.2956
  -5.500  -0.3985   0.04540   0.03844  -0.0770   0.8491   0.3340
  -5.250  -0.3671   0.04786   0.04115  -0.0746   0.8384   0.3722
  -5.000  -0.3544   0.04936   0.04263  -0.0714   0.8272   0.4047
  -4.750  -0.3154   0.05217   0.04550  -0.0684   0.8183   0.4397
  -4.500  -0.3159   0.05310   0.04636  -0.0644   0.8075   0.4620
  -4.250  -0.2802   0.05531   0.04858  -0.0607   0.7997   0.4872
  -4.000  -0.2848   0.05596   0.04916  -0.0567   0.7899   0.5048
  -3.750  -0.2578   0.05693   0.05004  -0.0539   0.7821   0.5281
  -3.500  -0.2552   0.05741   0.05043  -0.0508   0.7742   0.5458
  -3.250  -0.2477   0.05784   0.05075  -0.0482   0.7669   0.5641
  -3.000  -0.2149   0.05838   0.05118  -0.0462   0.7611   0.5872
  -2.750  -0.2292   0.05877   0.05149  -0.0435   0.7538   0.6009
  -2.500  -0.2214   0.05943   0.05211  -0.0397   0.7474   0.6161
  -2.250  -0.1878   0.05931   0.05180  -0.0398   0.7419   0.6377
  -2.000  -0.2009   0.06011   0.05262  -0.0358   0.7377   0.6455
  -1.750  -0.2006   0.06054   0.05294  -0.0347   0.7335   0.6575
  -1.500  -0.1932   0.06094   0.05330  -0.0323   0.7291   0.6680
  -1.250  -0.1646   0.06093   0.05312  -0.0333   0.7240   0.6824
  -1.000  -0.1531   0.06150   0.05357  -0.0327   0.7209   0.6934
  -0.750  -0.1515   0.06226   0.05427  -0.0316   0.7202   0.7017
  -0.500  -0.1462   0.06307   0.05503  -0.0307   0.7202   0.7104
  -0.250  -0.2162   0.06624   0.05848  -0.0262   0.7831   0.7094
   0.000  -0.1660   0.06575   0.05773  -0.0282   0.7450   0.7228
   0.250  -0.2654   0.06745   0.05978  -0.0204   0.8641   0.7190
   0.500  -0.2380   0.06868   0.06089  -0.0219   0.8564   0.7297
   0.750  -0.2216   0.06930   0.06138  -0.0225   0.8510   0.7387
   1.000  -0.2034   0.06973   0.06172  -0.0226   0.8410   0.7476
   1.250  -0.1686   0.07201   0.06384  -0.0259   0.8362   0.7582
   1.500  -0.1671   0.07131   0.06310  -0.0238   0.8273   0.7653
   1.750  -0.1342   0.07308   0.06472  -0.0267   0.8195   0.7758
   2.000  -0.1237   0.07353   0.06510  -0.0261   0.8122   0.7842
   2.250  -0.0982   0.07470   0.06616  -0.0277   0.8026   0.7948
   2.500  -0.0636   0.07756   0.06890  -0.0306   0.7983   0.8070
   2.750  -0.0645   0.07659   0.06792  -0.0285   0.7876   0.8152
   3.250  -0.0264   0.07894   0.07018  -0.0297   0.7721   0.8406
   3.500   0.0013   0.08058   0.07179  -0.0314   0.7637   0.8566
   3.750   0.0289   0.08300   0.07421  -0.0333   0.7596   0.8763
   4.000   0.0352   0.08267   0.07395  -0.0326   0.7483   0.8950
   4.250   0.0891   0.08641   0.07777  -0.0397   0.7422   0.9347
   4.500   0.0967   0.08609   0.07746  -0.0407   0.7301   1.0000
   4.750   0.1428   0.08973   0.08091  -0.0473   0.7238   1.0000
   5.000   0.1509   0.09057   0.08166  -0.0485   0.7158   1.0000
   5.250   0.1834   0.09306   0.08399  -0.0527   0.7073   1.0000
   5.500   0.2330   0.09814   0.08884  -0.0589   0.7032   1.0000
   5.750   0.2249   0.09712   0.08777  -0.0573   0.6912   1.0000
   6.000   0.2634   0.10083   0.09129  -0.0613   0.6856   1.0000
   6.250   0.2649   0.10171   0.09208  -0.0609   0.6781   1.0000
   6.500   0.2874   0.10387   0.09412  -0.0625   0.6698   1.0000
   6.750   0.3265   0.10840   0.09849  -0.0658   0.6655   1.0000
   7.000   0.3162   0.10781   0.09787  -0.0638   0.6548   1.0000
   7.250   0.3448   0.11090   0.10086  -0.0657   0.6485   1.0000
   7.500   0.3697   0.11467   0.10453  -0.0674   0.6452   1.0000
   7.750   0.3642   0.11427   0.10412  -0.0659   0.6339   1.0000
   8.000   0.3957   0.11797   0.10775  -0.0678   0.6286   1.0000
   8.250   0.3933   0.11881   0.10856  -0.0669   0.6211   1.0000
   8.500   0.4107   0.12095   0.11067  -0.0676   0.6131   1.0000
   8.750   0.4435   0.12529   0.11495  -0.0696   0.6089   1.0000
   9.000   0.4335   0.12518   0.11485  -0.0682   0.6002   1.0000
   9.250   0.4556   0.12787   0.11751  -0.0692   0.5929   1.0000
   9.500   0.4915   0.13320   0.12281  -0.0715   0.5894   1.0000
   9.750   0.4732   0.13178   0.12141  -0.0696   0.5796   1.0000
  10.000   0.4977   0.13495   0.12456  -0.0708   0.5732   1.0000
  10.250   0.5172   0.13878   0.12839  -0.0719   0.5693   1.0000
  10.500   0.5126   0.13859   0.12822  -0.0712   0.5590   1.0000
  10.750   0.5392   0.14236   0.13199  -0.0724   0.5537   1.0000
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