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

EPPLER 858 AIRFOIL (e858-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 858 AIRFOIL (e858-il)
Reynolds number: 100,000
Max Cl/Cd: 34.55 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e858-il-100000-n5.txt
Download as CSV file: xf-e858-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 858 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -15.750  -0.4582   0.11505   0.10911  -0.0623   1.0000   0.0346
 -15.500  -0.5023   0.10320   0.09699  -0.0681   1.0000   0.0347
 -15.250  -0.5300   0.09521   0.08876  -0.0718   1.0000   0.0350
 -15.000  -0.5324   0.09201   0.08558  -0.0726   1.0000   0.0355
 -14.750  -0.5384   0.08832   0.08186  -0.0736   1.0000   0.0361
 -14.500  -0.5464   0.08443   0.07792  -0.0746   1.0000   0.0367
 -14.250  -0.5549   0.08062   0.07402  -0.0755   1.0000   0.0374
 -14.000  -0.5617   0.07724   0.07055  -0.0760   1.0000   0.0380
 -13.750  -0.5700   0.07383   0.06701  -0.0764   1.0000   0.0389
 -13.500  -0.5734   0.07133   0.06450  -0.0761   1.0000   0.0396
 -13.250  -0.5766   0.06913   0.06233  -0.0756   1.0000   0.0403
 -13.000  -0.5817   0.06691   0.06011  -0.0750   1.0000   0.0411
 -12.750  -0.5885   0.06470   0.05789  -0.0741   1.0000   0.0421
 -12.500  -0.5816   0.06203   0.05510  -0.0759   0.9973   0.0439
 -12.250  -0.5618   0.05919   0.05229  -0.0805   0.9917   0.0460
 -12.000  -0.5431   0.05635   0.04934  -0.0844   0.9846   0.0489
 -11.750  -0.5255   0.05373   0.04673  -0.0878   0.9766   0.0514
 -11.500  -0.5101   0.05127   0.04422  -0.0903   0.9653   0.0545
 -11.250  -0.5005   0.04883   0.04179  -0.0918   0.9495   0.0573
 -10.750  -0.4785   0.04415   0.03702  -0.0951   0.9177   0.0652
 -10.500  -0.4671   0.04178   0.03467  -0.0969   0.9021   0.0692
 -10.250  -0.4537   0.03942   0.03226  -0.0991   0.8866   0.0746
 -10.000  -0.4375   0.03711   0.02990  -0.1015   0.8711   0.0807
  -9.750  -0.4188   0.03477   0.02748  -0.1045   0.8556   0.0880
  -9.500  -0.3987   0.03242   0.02504  -0.1080   0.8395   0.0964
  -9.000  -0.3570   0.02852   0.02093  -0.1137   0.8038   0.1169
  -8.750  -0.3359   0.02709   0.01941  -0.1152   0.7856   0.1293
  -8.500  -0.3168   0.02589   0.01813  -0.1159   0.7683   0.1434
  -8.250  -0.2966   0.02483   0.01698  -0.1164   0.7520   0.1599
  -8.000  -0.2762   0.02386   0.01595  -0.1167   0.7367   0.1785
  -7.750  -0.2613   0.02306   0.01518  -0.1156   0.7214   0.1983
  -7.500  -0.2409   0.02238   0.01450  -0.1153   0.7079   0.2234
  -7.250  -0.2196   0.02191   0.01408  -0.1148   0.6943   0.2510
  -7.000  -0.1982   0.02167   0.01384  -0.1140   0.6814   0.2773
  -6.750  -0.1717   0.02159   0.01366  -0.1139   0.6696   0.3017
  -6.500  -0.1498   0.02159   0.01358  -0.1128   0.6573   0.3208
  -6.000  -0.1002   0.02171   0.01342  -0.1116   0.6351   0.3519
  -5.750  -0.0742   0.02173   0.01317  -0.1113   0.6254   0.3661
  -5.500  -0.0503   0.02196   0.01340  -0.1102   0.6148   0.3753
  -5.250  -0.0259   0.02192   0.01308  -0.1097   0.6054   0.3871
  -5.000  -0.0005   0.02217   0.01332  -0.1088   0.5960   0.3944
  -4.750   0.0232   0.02220   0.01317  -0.1080   0.5869   0.4041
  -4.500   0.0496   0.02230   0.01312  -0.1076   0.5790   0.4116
  -4.250   0.0730   0.02244   0.01323  -0.1065   0.5703   0.4187
  -4.000   0.0972   0.02238   0.01292  -0.1060   0.5627   0.4278
  -3.750   0.1227   0.02250   0.01299  -0.1053   0.5554   0.4335
  -3.500   0.1461   0.02261   0.01306  -0.1043   0.5473   0.4398
  -3.250   0.1709   0.02259   0.01285  -0.1038   0.5404   0.4475
  -3.000   0.1955   0.02256   0.01268  -0.1033   0.5338   0.4534
  -2.750   0.2190   0.02262   0.01274  -0.1024   0.5267   0.4577
  -2.500   0.2441   0.02263   0.01265  -0.1018   0.5207   0.4625
  -2.250   0.2707   0.02263   0.01244  -0.1017   0.5154   0.4680
  -2.000   0.2932   0.02256   0.01225  -0.1010   0.5089   0.4736
  -1.750   0.3171   0.02256   0.01223  -0.1002   0.5028   0.4770
  -1.500   0.3428   0.02259   0.01216  -0.0998   0.4976   0.4806
  -1.250   0.3675   0.02265   0.01215  -0.0992   0.4924   0.4850
  -1.000   0.3903   0.02269   0.01214  -0.0984   0.4867   0.4899
  -0.750   0.4147   0.02272   0.01204  -0.0980   0.4817   0.4949
  -0.500   0.4408   0.02275   0.01196  -0.0977   0.4774   0.4988
  -0.250   0.4669   0.02285   0.01200  -0.0975   0.4734   0.5025
   0.000   0.4884   0.02297   0.01215  -0.0964   0.4685   0.5068
   0.250   0.5114   0.02307   0.01222  -0.0956   0.4637   0.5118
   0.500   0.5361   0.02316   0.01220  -0.0952   0.4594   0.5168
   0.750   0.5629   0.02326   0.01215  -0.0952   0.4557   0.5213
   1.000   0.5858   0.02341   0.01233  -0.0943   0.4518   0.5251
   1.250   0.6068   0.02360   0.01256  -0.0932   0.4477   0.5297
   1.500   0.6294   0.02378   0.01273  -0.0924   0.4439   0.5351
   1.750   0.6536   0.02395   0.01283  -0.0919   0.4404   0.5408
   2.000   0.6791   0.02410   0.01292  -0.0917   0.4373   0.5455
   2.250   0.7070   0.02429   0.01303  -0.0918   0.4344   0.5501
   2.500   0.7236   0.02456   0.01340  -0.0900   0.4307   0.5554
   2.750   0.7428   0.02482   0.01369  -0.0887   0.4269   0.5614
   3.000   0.7642   0.02507   0.01393  -0.0878   0.4235   0.5676
   3.250   0.7866   0.02529   0.01417  -0.0870   0.4206   0.5730
   3.500   0.8112   0.02552   0.01437  -0.0866   0.4180   0.5792
   3.750   0.8391   0.02578   0.01456  -0.0869   0.4157   0.5864
   4.000   0.8556   0.02613   0.01499  -0.0851   0.4129   0.5928
   4.250   0.8674   0.02650   0.01548  -0.0826   0.4098   0.5993
   4.500   0.8819   0.02687   0.01591  -0.0806   0.4067   0.6074
   4.750   0.8997   0.02722   0.01629  -0.0791   0.4038   0.6153
   5.000   0.9207   0.02753   0.01663  -0.0782   0.4012   0.6234
   5.250   0.9452   0.02784   0.01691  -0.0779   0.3988   0.6334
   5.500   0.9723   0.02814   0.01721  -0.0781   0.3969   0.6435
   5.750   0.9892   0.02866   0.01779  -0.0766   0.3947   0.6554
   6.000   0.9937   0.02936   0.01868  -0.0733   0.3920   0.6663
   6.250   1.0012   0.03005   0.01951  -0.0705   0.3892   0.6800
   6.500   1.0114   0.03065   0.02024  -0.0682   0.3865   0.6949
   6.750   1.0260   0.03114   0.02082  -0.0665   0.3839   0.7133
   7.000   1.0445   0.03151   0.02130  -0.0653   0.3817   0.7377
   7.250   1.0665   0.03177   0.02166  -0.0646   0.3798   0.7728
   7.500   1.0949   0.03189   0.02196  -0.0649   0.3781   0.8455
   7.750   1.1078   0.03271   0.02294  -0.0637   0.3759   1.0000
   8.000   1.0923   0.03440   0.02479  -0.0589   0.3731   1.0000
   8.250   1.0793   0.03623   0.02674  -0.0548   0.3700   1.0000
   8.500   1.0735   0.03792   0.02850  -0.0519   0.3671   1.0000
   8.750   1.0799   0.03914   0.02973  -0.0503   0.3645   1.0000
   9.000   1.0974   0.03990   0.03045  -0.0497   0.3626   1.0000
   9.250   1.1224   0.04033   0.03083  -0.0498   0.3610   1.0000
   9.500   1.1530   0.04057   0.03100  -0.0504   0.3596   1.0000
  10.000   0.9091   0.06438   0.05558  -0.0388   0.3408   1.0000
  10.250   0.9555   0.06215   0.05327  -0.0388   0.3415   1.0000
<< Back to EPPLER 858 AIRFOIL (e858-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 858 AIRFOIL (e858-il)