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

EPPLER 559 AIRFOIL (e559-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 559 AIRFOIL (e559-il)
Reynolds number: 200,000
Max Cl/Cd: 73.24 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e559-il-200000-n5.txt
Download as CSV file: xf-e559-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 559 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -16.000  -0.5776   0.12045   0.11609  -0.0468   1.0000   0.0203
 -15.750  -0.6123   0.10936   0.10479  -0.0524   1.0000   0.0203
 -15.500  -0.6313   0.10195   0.09724  -0.0561   1.0000   0.0204
 -15.250  -0.6481   0.09525   0.09039  -0.0595   1.0000   0.0206
 -15.000  -0.6591   0.08986   0.08488  -0.0621   1.0000   0.0208
 -14.750  -0.6709   0.08445   0.07933  -0.0648   1.0000   0.0210
 -14.500  -0.6733   0.08105   0.07590  -0.0659   1.0000   0.0213
 -14.250  -0.6750   0.07787   0.07271  -0.0669   1.0000   0.0215
 -14.000  -0.6778   0.07455   0.06936  -0.0680   1.0000   0.0218
 -13.750  -0.6814   0.07121   0.06600  -0.0691   1.0000   0.0220
 -13.500  -0.6857   0.06793   0.06268  -0.0701   1.0000   0.0223
 -13.250  -0.6907   0.06479   0.05949  -0.0708   1.0000   0.0224
 -13.000  -0.6976   0.06150   0.05616  -0.0717   1.0000   0.0228
 -12.750  -0.7044   0.05837   0.05298  -0.0723   1.0000   0.0231
 -12.500  -0.7119   0.05534   0.04991  -0.0727   1.0000   0.0233
 -12.250  -0.7209   0.05231   0.04683  -0.0729   1.0000   0.0237
 -12.000  -0.7310   0.04939   0.04387  -0.0729   1.0000   0.0239
 -11.750  -0.7438   0.04650   0.04094  -0.0724   1.0000   0.0242
 -11.500  -0.7595   0.04374   0.03814  -0.0714   1.0000   0.0244
 -11.250  -0.7827   0.04109   0.03547  -0.0694   1.0000   0.0245
 -11.000  -0.7724   0.03766   0.03188  -0.0745   0.9948   0.0251
 -10.750  -0.7504   0.03332   0.02743  -0.0833   0.9872   0.0259
 -10.500  -0.7224   0.03037   0.02434  -0.0898   0.9810   0.0270
 -10.250  -0.6964   0.02835   0.02219  -0.0934   0.9738   0.0281
 -10.000  -0.6704   0.02668   0.02038  -0.0960   0.9667   0.0293
  -9.500  -0.6163   0.02371   0.01723  -0.1004   0.9518   0.0328
  -9.250  -0.5839   0.02250   0.01590  -0.1028   0.9464   0.0354
  -9.000  -0.5582   0.02132   0.01467  -0.1039   0.9373   0.0384
  -8.750  -0.5248   0.02019   0.01346  -0.1062   0.9322   0.0433
  -8.500  -0.4978   0.01921   0.01244  -0.1070   0.9229   0.0505
  -8.250  -0.4635   0.01816   0.01139  -0.1092   0.9174   0.0634
  -8.000  -0.4336   0.01735   0.01058  -0.1102   0.9081   0.0776
  -7.750  -0.3967   0.01655   0.00978  -0.1126   0.9021   0.0941
  -7.500  -0.3631   0.01593   0.00915  -0.1142   0.8930   0.1100
  -7.250  -0.3225   0.01535   0.00854  -0.1170   0.8861   0.1274
  -7.000  -0.2845   0.01488   0.00804  -0.1193   0.8762   0.1436
  -6.750  -0.2416   0.01442   0.00758  -0.1226   0.8674   0.1634
  -6.500  -0.2029   0.01408   0.00721  -0.1249   0.8551   0.1820
  -6.250  -0.1642   0.01382   0.00687  -0.1272   0.8420   0.1975
  -6.000  -0.1272   0.01360   0.00657  -0.1290   0.8280   0.2095
  -5.750  -0.0923   0.01344   0.00626  -0.1303   0.8133   0.2203
  -5.500  -0.0597   0.01331   0.00603  -0.1312   0.7981   0.2305
  -5.250  -0.0287   0.01320   0.00582  -0.1318   0.7833   0.2397
  -5.000   0.0010   0.01314   0.00560  -0.1320   0.7683   0.2491
  -4.750   0.0294   0.01306   0.00546  -0.1320   0.7536   0.2569
  -4.500   0.0577   0.01301   0.00527  -0.1320   0.7395   0.2656
  -4.250   0.0856   0.01295   0.00515  -0.1319   0.7259   0.2728
  -4.000   0.1135   0.01293   0.00500  -0.1317   0.7129   0.2810
  -3.500   0.1683   0.01289   0.00479  -0.1313   0.6874   0.2961
  -3.250   0.1956   0.01287   0.00470  -0.1311   0.6754   0.3032
  -3.000   0.2229   0.01287   0.00461  -0.1308   0.6638   0.3107
  -2.750   0.2500   0.01287   0.00453  -0.1305   0.6521   0.3181
  -2.500   0.2771   0.01287   0.00448  -0.1303   0.6410   0.3250
  -2.250   0.3044   0.01291   0.00440  -0.1300   0.6305   0.3328
  -2.000   0.3313   0.01291   0.00438  -0.1297   0.6195   0.3395
  -1.750   0.3584   0.01294   0.00434  -0.1295   0.6095   0.3472
  -1.500   0.3854   0.01297   0.00432  -0.1292   0.5997   0.3540
  -1.250   0.4124   0.01300   0.00431  -0.1289   0.5898   0.3617
  -1.000   0.4393   0.01306   0.00429  -0.1286   0.5805   0.3691
  -0.750   0.4662   0.01309   0.00431  -0.1283   0.5708   0.3762
  -0.500   0.4933   0.01316   0.00431  -0.1281   0.5623   0.3845
  -0.250   0.5200   0.01320   0.00436  -0.1278   0.5533   0.3918
   0.000   0.5469   0.01328   0.00438  -0.1275   0.5448   0.4000
   0.250   0.5736   0.01333   0.00444  -0.1272   0.5361   0.4077
   0.500   0.6004   0.01342   0.00448  -0.1269   0.5285   0.4165
   0.750   0.6271   0.01348   0.00456  -0.1266   0.5204   0.4244
   1.000   0.6536   0.01358   0.00462  -0.1263   0.5129   0.4336
   1.250   0.6802   0.01365   0.00472  -0.1259   0.5049   0.4425
   1.750   0.7332   0.01384   0.00493  -0.1253   0.4909   0.4617
   2.000   0.7594   0.01396   0.00503  -0.1249   0.4839   0.4725
   2.250   0.7858   0.01406   0.00517  -0.1246   0.4771   0.4838
   2.500   0.8120   0.01417   0.00531  -0.1242   0.4704   0.4954
   2.750   0.8380   0.01431   0.00545  -0.1239   0.4646   0.5085
   3.250   0.8898   0.01454   0.00579  -0.1231   0.4516   0.5384
   3.500   0.9156   0.01467   0.00598  -0.1226   0.4460   0.5558
   3.750   0.9412   0.01479   0.00619  -0.1222   0.4400   0.5756
   4.000   0.9663   0.01493   0.00639  -0.1216   0.4345   0.5978
   4.250   0.9914   0.01506   0.00662  -0.1211   0.4290   0.6244
   4.500   1.0160   0.01517   0.00687  -0.1204   0.4231   0.6553
   4.750   1.0398   0.01530   0.00710  -0.1195   0.4181   0.6925
   5.000   1.0626   0.01541   0.00735  -0.1185   0.4131   0.7393
   5.250   1.0831   0.01545   0.00760  -0.1168   0.4076   0.8024
   5.500   1.1012   0.01534   0.00769  -0.1145   0.4027   1.0000
   5.750   1.1265   0.01563   0.00793  -0.1141   0.3980   1.0000
   6.000   1.1515   0.01587   0.00822  -0.1137   0.3924   1.0000
   6.250   1.1759   0.01615   0.00849  -0.1131   0.3870   1.0000
   6.500   1.2000   0.01647   0.00876  -0.1125   0.3823   1.0000
   6.750   1.2241   0.01673   0.00910  -0.1119   0.3767   1.0000
   7.000   1.2473   0.01703   0.00940  -0.1112   0.3713   1.0000
   7.250   1.2702   0.01737   0.00971  -0.1104   0.3665   1.0000
   7.500   1.2931   0.01766   0.01008  -0.1096   0.3608   1.0000
   7.750   1.3149   0.01798   0.01043  -0.1086   0.3552   1.0000
   8.000   1.3362   0.01835   0.01077  -0.1076   0.3503   1.0000
   8.250   1.3572   0.01865   0.01118  -0.1064   0.3444   1.0000
   8.500   1.3763   0.01900   0.01155  -0.1050   0.3387   1.0000
   8.750   1.3948   0.01938   0.01195  -0.1035   0.3334   1.0000
   9.000   1.4130   0.01974   0.01239  -0.1019   0.3270   1.0000
   9.250   1.4297   0.02016   0.01282  -0.1002   0.3210   1.0000
   9.500   1.4467   0.02058   0.01332  -0.0985   0.3148   1.0000
   9.750   1.4622   0.02103   0.01382  -0.0966   0.3079   1.0000
  10.000   1.4769   0.02154   0.01435  -0.0947   0.3017   1.0000
  10.250   1.4917   0.02204   0.01495  -0.0928   0.2943   1.0000
  10.500   1.5036   0.02267   0.01557  -0.0906   0.2875   1.0000
  10.750   1.5174   0.02325   0.01627  -0.0888   0.2797   1.0000
  11.250   1.5390   0.02473   0.01785  -0.0845   0.2638   1.0000
  11.500   1.5469   0.02565   0.01879  -0.0822   0.2560   1.0000
  11.750   1.5559   0.02659   0.01982  -0.0802   0.2469   1.0000
  12.000   1.5623   0.02772   0.02099  -0.0781   0.2383   1.0000
  12.250   1.5670   0.02901   0.02232  -0.0760   0.2289   1.0000
  12.500   1.5718   0.03040   0.02376  -0.0741   0.2190   1.0000
  12.750   1.5735   0.03208   0.02547  -0.0721   0.2095   1.0000
  13.000   1.5737   0.03398   0.02740  -0.0704   0.1993   1.0000
  13.250   1.5741   0.03601   0.02946  -0.0688   0.1891   1.0000
  13.500   1.5714   0.03841   0.03190  -0.0674   0.1796   1.0000
  13.750   1.5673   0.04107   0.03459  -0.0662   0.1704   1.0000
  14.000   1.5641   0.04382   0.03739  -0.0653   0.1611   1.0000
  14.250   1.5577   0.04702   0.04063  -0.0646   0.1529   1.0000
  14.500   1.5519   0.05031   0.04397  -0.0642   0.1445   1.0000
  14.750   1.5451   0.05387   0.04758  -0.0640   0.1369   1.0000
  15.000   1.5366   0.05779   0.05155  -0.0640   0.1296   1.0000
  15.250   1.5297   0.06165   0.05548  -0.0643   0.1226   1.0000
  15.500   1.5198   0.06603   0.05991  -0.0648   0.1162   1.0000
  15.750   1.5126   0.07022   0.06417  -0.0654   0.1099   1.0000
  16.000   1.5026   0.07490   0.06891  -0.0663   0.1040   1.0000
  16.250   1.4947   0.07941   0.07350  -0.0673   0.0986   1.0000
  16.500   1.4860   0.08415   0.07831  -0.0685   0.0932   1.0000
  16.750   1.4765   0.08912   0.08334  -0.0699   0.0886   1.0000
  17.000   1.4698   0.09374   0.08804  -0.0712   0.0835   1.0000
<< Back to EPPLER 559 AIRFOIL (e559-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 559 AIRFOIL (e559-il)