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

EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 200,000
Max Cl/Cd: 77.49 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e851-il-200000.txt
Download as CSV file: xf-e851-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E851 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3907   0.08745   0.08426  -0.0382   1.0000   0.0331
  -8.750  -0.3965   0.08409   0.08095  -0.0377   1.0000   0.0336
  -8.500  -0.4042   0.08080   0.07771  -0.0371   1.0000   0.0341
  -8.250  -0.4143   0.07752   0.07449  -0.0363   1.0000   0.0345
  -8.000  -0.5190   0.08168   0.07841  -0.0453   1.0000   0.0298
  -7.750  -0.5348   0.07895   0.07565  -0.0446   1.0000   0.0299
  -7.250  -0.5648   0.06912   0.06588  -0.0430   1.0000   0.0309
  -7.000  -0.5688   0.06594   0.06272  -0.0417   1.0000   0.0315
  -6.750  -0.5715   0.06282   0.05957  -0.0409   1.0000   0.0321
  -6.500  -0.5713   0.05970   0.05638  -0.0405   1.0000   0.0329
  -6.250  -0.5686   0.05625   0.05282  -0.0406   1.0000   0.0339
  -6.000  -0.5624   0.05258   0.04898  -0.0409   1.0000   0.0352
  -5.750  -0.5453   0.04836   0.04446  -0.0430   0.9988   0.0374
  -5.500  -0.5146   0.04315   0.03839  -0.0478   0.9954   0.0431
  -5.000  -0.4617   0.03556   0.03049  -0.0516   0.9910   0.0477
  -4.750  -0.4223   0.02847   0.02235  -0.0502   0.9902   0.0161
  -4.500  -0.3921   0.02501   0.01837  -0.0510   0.9886   0.0148
  -4.250  -0.3607   0.02242   0.01535  -0.0515   0.9870   0.0141
  -4.000  -0.3287   0.02040   0.01294  -0.0519   0.9852   0.0140
  -3.750  -0.2963   0.01877   0.01113  -0.0526   0.9835   0.0150
  -3.500  -0.2627   0.01771   0.00992  -0.0538   0.9819   0.0171
  -3.250  -0.2296   0.01579   0.00819  -0.0554   0.9805   0.0794
  -3.000  -0.2044   0.01369   0.00791  -0.0566   0.9784   0.4965
  -2.750  -0.1763   0.01376   0.00796  -0.0567   0.9749   0.5524
  -2.500  -0.1461   0.01396   0.00815  -0.0572   0.9719   0.5993
  -2.250  -0.1133   0.01418   0.00833  -0.0583   0.9695   0.6290
  -2.000  -0.0867   0.01428   0.00838  -0.0582   0.9659   0.6511
  -1.750  -0.0602   0.01439   0.00843  -0.0579   0.9619   0.6773
  -1.500  -0.0297   0.01451   0.00857  -0.0585   0.9588   0.6990
  -1.250   0.0051   0.01463   0.00865  -0.0601   0.9565   0.7135
  -1.000   0.0298   0.01465   0.00866  -0.0598   0.9520   0.7256
  -0.750   0.0591   0.01469   0.00870  -0.0604   0.9480   0.7388
  -0.500   0.0929   0.01475   0.00874  -0.0619   0.9451   0.7527
  -0.250   0.1278   0.01483   0.00884  -0.0635   0.9426   0.7679
   0.000   0.1490   0.01483   0.00889  -0.0625   0.9365   0.7838
   0.250   0.1817   0.01484   0.00897  -0.0636   0.9330   0.8016
   0.500   0.2182   0.01484   0.00906  -0.0654   0.9305   0.8231
   0.750   0.2366   0.01482   0.00916  -0.0636   0.9234   0.8493
   1.000   0.2694   0.01471   0.00919  -0.0645   0.9197   0.8854
   1.250   0.3193   0.01453   0.00918  -0.0690   0.9178   0.9604
   1.500   0.3494   0.01459   0.00923  -0.0702   0.9106   1.0000
   1.750   0.3916   0.01453   0.00919  -0.0733   0.9065   1.0000
   2.000   0.4296   0.01447   0.00923  -0.0755   0.9009   1.0000
   2.250   0.4705   0.01422   0.00906  -0.0779   0.8944   1.0000
   2.500   0.5337   0.01339   0.00837  -0.0841   0.8912   1.0000
   2.750   0.5920   0.01216   0.00728  -0.0886   0.8791   1.0000
   3.000   0.6405   0.01119   0.00653  -0.0913   0.8655   1.0000
   3.250   0.6733   0.01069   0.00614  -0.0913   0.8481   1.0000
   3.500   0.7052   0.01006   0.00561  -0.0906   0.8193   1.0000
   3.750   0.7309   0.00961   0.00516  -0.0887   0.7655   1.0000
   4.000   0.7532   0.00972   0.00474  -0.0861   0.6160   1.0000
   4.250   0.7489   0.01154   0.00524  -0.0797   0.3779   1.0000
   4.500   0.7460   0.01395   0.00628  -0.0747   0.1318   1.0000
   4.750   0.7569   0.01579   0.00750  -0.0718   0.0454   1.0000
   5.000   0.7728   0.01736   0.00897  -0.0694   0.0176   1.0000
   5.250   0.7883   0.01942   0.01111  -0.0670   0.0131   1.0000
   5.500   0.8117   0.02108   0.01290  -0.0658   0.0119   1.0000
   5.750   0.8392   0.02347   0.01561  -0.0651   0.0111   1.0000
   6.000   0.8675   0.02629   0.01874  -0.0645   0.0111   1.0000
   6.250   0.8925   0.02958   0.02245  -0.0632   0.0114   1.0000
   6.500   0.9119   0.03332   0.02666  -0.0612   0.0120   1.0000
   6.750   0.9243   0.03793   0.03174  -0.0585   0.0129   1.0000
   7.000   0.9546   0.04634   0.04093  -0.0540   0.0389   1.0000
   7.250   0.9669   0.05069   0.04537  -0.0525   0.0379   1.0000
   7.500   0.9604   0.05424   0.04956  -0.0479   0.0343   1.0000
   7.750   0.9635   0.05759   0.05326  -0.0447   0.0324   1.0000
   8.000   0.9634   0.06131   0.05725  -0.0419   0.0312   1.0000
   8.250   0.9601   0.06508   0.06124  -0.0393   0.0304   1.0000
   8.500   0.9535   0.06879   0.06515  -0.0368   0.0297   1.0000
   8.750   0.9429   0.07230   0.06880  -0.0341   0.0292   1.0000
   9.000   0.9274   0.07564   0.07226  -0.0313   0.0290   1.0000
   9.250   0.9100   0.07935   0.07609  -0.0296   0.0288   1.0000
   9.500   0.8912   0.08370   0.08054  -0.0294   0.0288   1.0000
   9.750   0.8722   0.08892   0.08584  -0.0310   0.0289   1.0000
<< Back to EPPLER E851 AIRFOIL (e851-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E851 AIRFOIL (e851-il)