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

EPPLER 338 AIRFOIL (e338-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 338 AIRFOIL (e338-il)
Reynolds number: 1,000,000
Max Cl/Cd: 108.24 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e338-il-1000000-n5.txt
Download as CSV file: xf-e338-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 338 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2737   0.09332   0.09035  -0.0460   0.6446   0.0061
  -9.750  -0.2728   0.08968   0.08668  -0.0475   0.6334   0.0061
  -9.250  -0.2771   0.08123   0.07819  -0.0515   0.6134   0.0060
  -8.750  -0.2290   0.06337   0.06040  -0.0484   0.5696   0.0055
  -8.500  -0.2363   0.05886   0.05590  -0.0510   0.5641   0.0054
  -8.250  -0.2534   0.05364   0.05067  -0.0553   0.5602   0.0053
  -7.750  -0.2898   0.04608   0.04302  -0.0544   0.5511   0.0051
  -6.750  -0.4224   0.01959   0.01412  -0.0339   0.5572   0.0036
  -6.500  -0.4041   0.01742   0.01155  -0.0321   0.5487   0.0036
  -6.250  -0.3816   0.01629   0.01017  -0.0310   0.5398   0.0037
  -6.000  -0.3579   0.01544   0.00914  -0.0302   0.5317   0.0037
  -5.750  -0.3340   0.01468   0.00820  -0.0294   0.5227   0.0037
  -5.500  -0.3113   0.01354   0.00685  -0.0284   0.5144   0.0041
  -5.250  -0.2864   0.01313   0.00634  -0.0277   0.5056   0.0042
  -5.000  -0.2607   0.01287   0.00603  -0.0273   0.4975   0.0044
  -4.750  -0.2353   0.01256   0.00564  -0.0268   0.4904   0.0046
  -4.500  -0.2100   0.01222   0.00523  -0.0262   0.4836   0.0047
  -4.250  -0.1848   0.01190   0.00484  -0.0256   0.4764   0.0049
  -4.000  -0.1597   0.01156   0.00443  -0.0250   0.4698   0.0051
  -3.750  -0.1342   0.01134   0.00414  -0.0244   0.4624   0.0053
  -3.500  -0.1087   0.01110   0.00385  -0.0239   0.4562   0.0054
  -3.250  -0.0831   0.01088   0.00357  -0.0233   0.4499   0.0056
  -2.750  -0.0323   0.01043   0.00304  -0.0222   0.4389   0.0060
  -2.500  -0.0065   0.01027   0.00284  -0.0217   0.4328   0.0062
  -2.250   0.0194   0.01014   0.00268  -0.0213   0.4270   0.0067
  -2.000   0.0456   0.01002   0.00253  -0.0209   0.4218   0.0074
  -1.750   0.0717   0.00990   0.00237  -0.0205   0.4167   0.0081
  -1.500   0.0977   0.00979   0.00224  -0.0200   0.4119   0.0096
  -1.250   0.1241   0.00967   0.00211  -0.0197   0.4078   0.0118
  -1.000   0.1504   0.00958   0.00202  -0.0193   0.4030   0.0156
  -0.750   0.1765   0.00950   0.00194  -0.0189   0.3979   0.0249
  -0.500   0.2024   0.00936   0.00188  -0.0185   0.3940   0.0482
  -0.250   0.2278   0.00919   0.00184  -0.0181   0.3903   0.0912
   0.000   0.2522   0.00894   0.00179  -0.0175   0.3863   0.1686
   0.250   0.2406   0.00662   0.00170  -0.0099   0.3838   0.8350
   0.500   0.2651   0.00678   0.00190  -0.0089   0.3798   0.8791
   0.750   0.2913   0.00692   0.00200  -0.0083   0.3764   0.8922
   1.000   0.3176   0.00705   0.00211  -0.0078   0.3729   0.9003
   1.250   0.3435   0.00718   0.00220  -0.0072   0.3693   0.9070
   1.500   0.3688   0.00736   0.00235  -0.0064   0.3655   0.9134
   1.750   0.3961   0.00743   0.00238  -0.0062   0.3627   0.9160
   2.000   0.4242   0.00747   0.00239  -0.0063   0.3594   0.9168
   2.250   0.4522   0.00752   0.00240  -0.0064   0.3560   0.9175
   2.500   0.4799   0.00759   0.00242  -0.0065   0.3525   0.9183
   2.750   0.5077   0.00766   0.00245  -0.0065   0.3491   0.9191
   3.000   0.5358   0.00771   0.00248  -0.0067   0.3465   0.9196
   3.250   0.5637   0.00776   0.00251  -0.0068   0.3434   0.9202
   3.500   0.5912   0.00782   0.00255  -0.0068   0.3398   0.9209
   3.750   0.6185   0.00791   0.00261  -0.0068   0.3364   0.9215
   4.000   0.6460   0.00797   0.00266  -0.0069   0.3335   0.9221
   4.250   0.6736   0.00804   0.00272  -0.0069   0.3305   0.9227
   4.500   0.7011   0.00811   0.00280  -0.0070   0.3273   0.9234
   4.750   0.7281   0.00821   0.00287  -0.0070   0.3236   0.9242
   5.000   0.7548   0.00832   0.00297  -0.0069   0.3198   0.9250
   5.250   0.7823   0.00839   0.00305  -0.0070   0.3169   0.9257
   5.500   0.8095   0.00847   0.00314  -0.0070   0.3130   0.9264
   5.750   0.8363   0.00858   0.00324  -0.0070   0.3092   0.9271
   6.000   0.8628   0.00871   0.00335  -0.0069   0.3045   0.9279
   6.250   0.8898   0.00880   0.00345  -0.0070   0.2994   0.9287
   6.500   0.9157   0.00896   0.00358  -0.0069   0.2919   0.9295
   6.750   0.9422   0.00908   0.00370  -0.0068   0.2857   0.9303
   7.000   0.9679   0.00925   0.00384  -0.0067   0.2781   0.9311
   7.250   0.9940   0.00938   0.00399  -0.0066   0.2728   0.9319
   7.500   1.0199   0.00953   0.00414  -0.0065   0.2672   0.9327
   7.750   1.0452   0.00972   0.00431  -0.0064   0.2618   0.9335
   8.000   1.0705   0.00989   0.00449  -0.0062   0.2542   0.9344
   8.250   1.0946   0.01012   0.00469  -0.0058   0.2452   0.9354
   8.500   1.1182   0.01034   0.00491  -0.0054   0.2364   0.9365
   8.750   1.1416   0.01058   0.00514  -0.0049   0.2274   0.9375
   9.000   1.1640   0.01086   0.00540  -0.0042   0.2181   0.9387
   9.500   1.2073   0.01147   0.00598  -0.0027   0.1978   0.9412
   9.750   1.2279   0.01181   0.00630  -0.0019   0.1881   0.9426
  10.000   1.2476   0.01218   0.00666  -0.0008   0.1784   0.9442
  10.250   1.2662   0.01258   0.00704   0.0004   0.1678   0.9460
  10.500   1.2843   0.01299   0.00743   0.0016   0.1577   0.9478
  10.750   1.2995   0.01346   0.00787   0.0033   0.1464   0.9498
  11.000   1.3100   0.01393   0.00833   0.0060   0.1355   0.9528
  11.250   1.3203   0.01442   0.00882   0.0086   0.1268   0.9565
  11.500   1.3307   0.01495   0.00934   0.0111   0.1182   0.9611
  11.750   1.3368   0.01572   0.01007   0.0140   0.1054   0.9672
  12.250   1.3607   0.01765   0.01196   0.0161   0.0821   0.9820
  12.500   1.3784   0.01875   0.01306   0.0155   0.0721   0.9914
  12.750   1.3754   0.02048   0.01473   0.0177   0.0577   1.0000
  13.000   1.3756   0.02211   0.01636   0.0194   0.0514   1.0000
  13.250   1.3787   0.02371   0.01799   0.0205   0.0482   1.0000
  13.500   1.3811   0.02551   0.01985   0.0213   0.0458   1.0000
  13.750   1.3815   0.02763   0.02201   0.0219   0.0419   1.0000
  14.000   1.3841   0.02967   0.02412   0.0223   0.0404   1.0000
  14.250   1.3771   0.03269   0.02717   0.0226   0.0361   1.0000
  14.500   1.3775   0.03512   0.02966   0.0226   0.0339   1.0000
  14.750   1.3815   0.03723   0.03187   0.0226   0.0355   1.0000
  15.000   1.3823   0.03970   0.03442   0.0225   0.0354   1.0000
  15.250   1.3779   0.04272   0.03752   0.0223   0.0329   1.0000
  15.500   1.3657   0.04664   0.04146   0.0220   0.0292   1.0000
  15.750   1.3634   0.04957   0.04448   0.0216   0.0294   1.0000
  16.000   1.3496   0.05390   0.04882   0.0209   0.0243   1.0000
  16.250   1.3462   0.05714   0.05214   0.0203   0.0242   1.0000
  16.750   1.3269   0.06534   0.06045   0.0184   0.0189   1.0000
  17.000   1.3169   0.06963   0.06478   0.0172   0.0167   1.0000
  17.250   1.3106   0.07353   0.06876   0.0161   0.0160   1.0000
  17.500   1.3079   0.07704   0.07234   0.0151   0.0160   1.0000
  17.750   1.2996   0.08132   0.07669   0.0137   0.0142   1.0000
  18.000   1.2904   0.08580   0.08122   0.0122   0.0126   1.0000
<< Back to EPPLER 338 AIRFOIL (e338-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 338 AIRFOIL (e338-il)