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

EPPLER 639 AIRFOIL (e639-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 639 AIRFOIL (e639-il)
Reynolds number: 100,000
Max Cl/Cd: 32.77 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e639-il-100000.txt
Download as CSV file: xf-e639-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 639 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.2563   0.09603   0.09183  -0.0314   1.0000   0.0723
  -8.250  -0.2565   0.09339   0.08929  -0.0320   1.0000   0.0742
  -8.000  -0.2615   0.09103   0.08707  -0.0330   1.0000   0.0763
  -7.750  -0.2747   0.08924   0.08547  -0.0338   1.0000   0.0776
  -7.500  -0.2904   0.08774   0.08417  -0.0385   0.9873   0.0786
  -7.250  -0.2669   0.08147   0.07774  -0.0532   0.9530   0.0804
  -7.000  -0.2269   0.07607   0.07247  -0.0506   0.9402   0.0861
  -6.750  -0.2020   0.07228   0.06820  -0.0668   0.9077   0.0945
  -6.500  -0.1718   0.06552   0.06152  -0.0683   0.8869   0.0979
  -6.250  -0.1500   0.06223   0.05809  -0.0697   0.8596   0.1032
  -6.000  -0.1462   0.06051   0.05580  -0.0734   0.8349   0.1111
  -5.750  -0.1279   0.05629   0.05169  -0.0720   0.8139   0.1144
  -5.500  -0.1142   0.05419   0.04938  -0.0718   0.7950   0.1229
  -5.250  -0.1035   0.05128   0.04627  -0.0715   0.7786   0.1299
  -5.000  -0.0919   0.04981   0.04441  -0.0711   0.7631   0.1430
  -4.750  -0.0755   0.04681   0.04147  -0.0697   0.7483   0.1494
  -4.500  -0.0616   0.04467   0.03916  -0.0688   0.7347   0.1639
  -4.250  -0.0469   0.04287   0.03720  -0.0676   0.7221   0.1840
  -4.000  -0.0335   0.04089   0.03511  -0.0661   0.7109   0.2106
  -3.750  -0.0197   0.03913   0.03331  -0.0642   0.6997   0.2436
  -3.000   0.0207   0.03404   0.02817  -0.0570   0.6693   0.3655
  -2.750   0.0952   0.03138   0.02304  -0.0615   0.6610   0.1254
  -2.500   0.1259   0.02864   0.01974  -0.0600   0.6532   0.0926
  -2.250   0.1512   0.02696   0.01785  -0.0591   0.6439   0.0873
  -2.000   0.1788   0.02576   0.01626  -0.0582   0.6366   0.0856
  -1.750   0.2058   0.02484   0.01502  -0.0572   0.6274   0.0824
  -1.500   0.2348   0.02403   0.01392  -0.0565   0.6209   0.0810
  -1.250   0.2610   0.02326   0.01312  -0.0558   0.6123   0.0813
  -1.000   0.2885   0.02241   0.01221  -0.0553   0.6060   0.0833
  -0.750   0.3124   0.02199   0.01187  -0.0544   0.5981   0.0891
  -0.500   0.4640   0.01782   0.01017  -0.0760   0.5873   1.0000
  -0.250   0.4855   0.01810   0.01017  -0.0749   0.5802   1.0000
   0.000   0.5086   0.01831   0.01007  -0.0739   0.5749   1.0000
   0.250   0.5289   0.01869   0.01036  -0.0728   0.5680   1.0000
   0.500   0.5510   0.01897   0.01046  -0.0718   0.5623   1.0000
   0.750   0.5737   0.01930   0.01060  -0.0709   0.5578   1.0000
   1.000   0.5933   0.01977   0.01105  -0.0697   0.5513   1.0000
   1.250   0.6157   0.02007   0.01121  -0.0687   0.5462   1.0000
   1.500   0.6375   0.02048   0.01150  -0.0677   0.5415   1.0000
   1.750   0.6566   0.02102   0.01205  -0.0665   0.5357   1.0000
   2.000   0.6788   0.02140   0.01233  -0.0656   0.5312   1.0000
   2.250   0.7018   0.02181   0.01261  -0.0647   0.5273   1.0000
   2.500   0.7184   0.02253   0.01342  -0.0633   0.5212   1.0000
   2.750   0.7402   0.02295   0.01377  -0.0623   0.5167   1.0000
   3.000   0.7642   0.02332   0.01402  -0.0615   0.5133   1.0000
   3.250   0.7788   0.02428   0.01510  -0.0599   0.5079   1.0000
   3.500   0.7982   0.02490   0.01573  -0.0587   0.5031   1.0000
   3.750   0.8221   0.02524   0.01598  -0.0580   0.4994   1.0000
   4.000   0.8398   0.02607   0.01684  -0.0566   0.4951   1.0000
   4.250   0.8531   0.02717   0.01808  -0.0549   0.4899   1.0000
   4.500   0.8747   0.02770   0.01858  -0.0540   0.4860   1.0000
   4.750   0.9009   0.02800   0.01878  -0.0536   0.4829   1.0000
   5.000   0.9063   0.02966   0.02065  -0.0511   0.4771   1.0000
   5.250   0.9215   0.03068   0.02175  -0.0496   0.4726   1.0000
   5.500   0.9463   0.03105   0.02208  -0.0491   0.4694   1.0000
   5.750   0.9656   0.03187   0.02292  -0.0480   0.4656   1.0000
   6.000   0.9610   0.03429   0.02557  -0.0449   0.4590   1.0000
   6.250   0.9815   0.03498   0.02628  -0.0439   0.4554   1.0000
   6.500   1.0138   0.03490   0.02616  -0.0441   0.4528   1.0000
   6.750   0.9832   0.03939   0.03095  -0.0393   0.4449   1.0000
   7.000   0.9989   0.04040   0.03201  -0.0380   0.4406   1.0000
   7.250   1.0357   0.03992   0.03151  -0.0383   0.4382   1.0000
   7.500   0.9020   0.05279   0.04462  -0.0294   0.4246   1.0000
   7.750   0.9739   0.04930   0.04116  -0.0307   0.4237   1.0000
   8.000   1.0381   0.04651   0.03839  -0.0320   0.4226   1.0000
   8.250   0.7648   0.07610   0.06794  -0.0298   0.4019   1.0000
   8.500   0.7618   0.07945   0.07132  -0.0299   0.3966   1.0000
   8.750   0.8188   0.07650   0.06843  -0.0281   0.3927   1.0000
   9.000   0.7634   0.08567   0.07760  -0.0299   0.3866   1.0000
   9.250   0.7720   0.08781   0.07980  -0.0296   0.3808   1.0000
   9.500   0.8485   0.08264   0.07469  -0.0269   0.3769   1.0000
   9.750   0.7688   0.09460   0.08666  -0.0301   0.3697   1.0000
  10.000   0.8189   0.09214   0.08426  -0.0281   0.3633   1.0000
  10.250   0.7849   0.09908   0.09123  -0.0298   0.3562   1.0000
  10.500   0.8060   0.09994   0.09214  -0.0292   0.3501   1.0000
<< Back to EPPLER 639 AIRFOIL (e639-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 639 AIRFOIL (e639-il)