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

EPPLER 396 AIRFOIL (e396-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 396 AIRFOIL (e396-il)
Reynolds number: 100,000
Max Cl/Cd: 54.68 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e396-il-100000.txt
Download as CSV file: xf-e396-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 396 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2821   0.11359   0.10938  -0.0504   0.9778   0.1070
  -8.500  -0.2904   0.11074   0.10657  -0.0588   0.9714   0.1122
  -8.250  -0.3111   0.10876   0.10468  -0.0642   0.9617   0.1129
  -8.000  -0.2566   0.10290   0.09871  -0.0597   0.9621   0.1205
  -7.750  -0.2683   0.10034   0.09620  -0.0649   0.9543   0.1259
  -7.500  -0.2870   0.09781   0.09375  -0.0701   0.9454   0.1268
  -7.250  -0.3178   0.08161   0.07758  -0.0770   0.9369   0.0732
  -7.000  -0.3091   0.07426   0.07019  -0.0847   0.9314   0.0698
  -6.750  -0.3038   0.07458   0.07054  -0.0818   0.9244   0.0767
  -6.500  -0.3108   0.04406   0.03821  -0.1170   0.9176   0.0602
  -6.250  -0.2921   0.04376   0.03812  -0.1163   0.9120   0.0659
  -6.000  -0.2576   0.03845   0.03161  -0.1219   0.9078   0.0725
  -5.750  -0.2220   0.03667   0.02969  -0.1246   0.9043   0.0825
  -5.500  -0.1790   0.03480   0.02748  -0.1285   0.9018   0.0955
  -5.250  -0.1638   0.03387   0.02637  -0.1274   0.8950   0.1040
  -5.000  -0.1282   0.03266   0.02491  -0.1295   0.8910   0.1174
  -4.750  -0.0872   0.03179   0.02387  -0.1323   0.8881   0.1348
  -4.500  -0.0690   0.03142   0.02340  -0.1314   0.8817   0.1471
  -4.250  -0.0366   0.03094   0.02281  -0.1326   0.8772   0.1662
  -4.000   0.0034   0.03039   0.02221  -0.1351   0.8741   0.1893
  -3.750   0.0215   0.03043   0.02225  -0.1341   0.8676   0.2069
  -3.500   0.0525   0.03026   0.02200  -0.1350   0.8628   0.2315
  -3.250   0.0924   0.02997   0.02168  -0.1373   0.8596   0.2602
  -3.000   0.1077   0.03026   0.02205  -0.1358   0.8527   0.2796
  -2.750   0.1390   0.03022   0.02200  -0.1367   0.8479   0.3073
  -2.500   0.1790   0.03003   0.02182  -0.1389   0.8449   0.3391
  -2.250   0.1906   0.03054   0.02235  -0.1369   0.8372   0.3603
  -2.000   0.2238   0.03052   0.02236  -0.1379   0.8328   0.3910
  -1.750   0.2650   0.03033   0.02220  -0.1401   0.8300   0.4260
  -1.500   0.2711   0.03104   0.02299  -0.1373   0.8213   0.4462
  -1.250   0.3070   0.03097   0.02296  -0.1386   0.8174   0.4819
  -1.000   0.3290   0.03135   0.02341  -0.1380   0.8116   0.5124
  -0.750   0.3507   0.03167   0.02381  -0.1373   0.8052   0.5447
  -0.500   0.3884   0.03149   0.02374  -0.1385   0.8019   0.5872
  -0.250   0.3962   0.03232   0.02467  -0.1359   0.7936   0.6187
   0.000   0.4267   0.03229   0.02478  -0.1360   0.7891   0.6652
   0.250   0.4644   0.03195   0.02457  -0.1367   0.7865   0.7237
   0.500   0.4607   0.03299   0.02578  -0.1322   0.7764   0.7688
   0.750   0.4831   0.03248   0.02550  -0.1297   0.7730   0.8634
   1.000   0.4933   0.03339   0.02640  -0.1285   0.7634   1.0000
   1.250   0.5359   0.03351   0.02631  -0.1313   0.7595   1.0000
   1.500   0.5776   0.03358   0.02622  -0.1336   0.7562   1.0000
   1.750   0.5850   0.03491   0.02748  -0.1318   0.7462   1.0000
   2.000   0.6290   0.03476   0.02722  -0.1341   0.7435   1.0000
   2.250   0.6320   0.03634   0.02876  -0.1317   0.7329   1.0000
   2.500   0.6737   0.03620   0.02855  -0.1335   0.7297   1.0000
   2.750   0.6790   0.03774   0.03006  -0.1314   0.7194   1.0000
   3.000   0.7187   0.03761   0.02988  -0.1329   0.7158   1.0000
   3.250   0.7265   0.03907   0.03132  -0.1311   0.7059   1.0000
   3.500   0.7642   0.03897   0.03121  -0.1323   0.7018   1.0000
   3.750   0.8008   0.03891   0.03114  -0.1332   0.6977   1.0000
   4.250   0.8409   0.04046   0.03272  -0.1319   0.6824   1.0000
   4.500   0.8573   0.04140   0.03368  -0.1308   0.6734   1.0000
   4.750   0.8715   0.04256   0.03486  -0.1296   0.6644   1.0000
   5.000   0.9059   0.04235   0.03469  -0.1300   0.6591   1.0000
   5.250   0.9546   0.04115   0.03356  -0.1317   0.6570   1.0000
   5.500   1.0067   0.03957   0.03205  -0.1335   0.6555   1.0000
   5.750   0.9725   0.04393   0.03643  -0.1281   0.6355   1.0000
   6.000   1.0116   0.04310   0.03570  -0.1286   0.6307   1.0000
   6.250   1.0640   0.04114   0.03384  -0.1300   0.6289   1.0000
   6.500   1.1070   0.03983   0.03262  -0.1307   0.6249   1.0000
   6.750   1.1908   0.03593   0.02891  -0.1357   0.6265   1.0000
   7.000   1.1959   0.03690   0.02995  -0.1326   0.6146   1.0000
   7.250   1.2788   0.03335   0.02660  -0.1382   0.6130   1.0000
   7.500   1.2832   0.03407   0.02742  -0.1345   0.6009   1.0000
   7.750   1.3034   0.03395   0.02741  -0.1328   0.5902   1.0000
   8.000   1.3947   0.02998   0.02358  -0.1396   0.5845   1.0000
   8.250   1.4177   0.02955   0.02330  -0.1381   0.5714   1.0000
   8.500   1.4451   0.02890   0.02276  -0.1370   0.5578   1.0000
   8.750   1.4750   0.02810   0.02204  -0.1363   0.5426   1.0000
   9.000   1.5015   0.02746   0.02145  -0.1351   0.5252   1.0000
   9.250   1.5060   0.02771   0.02179  -0.1309   0.5071   1.0000
   9.500   1.5165   0.02777   0.02189  -0.1275   0.4869   1.0000
   9.750   1.5252   0.02800   0.02217  -0.1241   0.4651   1.0000
  10.000   1.5276   0.02862   0.02279  -0.1201   0.4419   1.0000
  10.250   1.5265   0.02957   0.02374  -0.1159   0.4170   1.0000
  10.500   1.5250   0.03072   0.02483  -0.1120   0.3905   1.0000
  10.750   1.5210   0.03220   0.02622  -0.1081   0.3629   1.0000
  11.000   1.5148   0.03402   0.02794  -0.1044   0.3346   1.0000
  11.250   1.5072   0.03617   0.02995  -0.1010   0.3063   1.0000
  11.500   1.4986   0.03862   0.03225  -0.0978   0.2786   1.0000
  11.750   1.4894   0.04132   0.03479  -0.0949   0.2523   1.0000
  12.000   1.4800   0.04426   0.03762  -0.0924   0.2266   1.0000
  12.250   1.4706   0.04740   0.04065  -0.0902   0.2026   1.0000
  12.500   1.4612   0.05072   0.04377  -0.0882   0.1811   1.0000
  12.750   1.4529   0.05415   0.04715  -0.0866   0.1598   1.0000
  13.000   1.4450   0.05769   0.05051  -0.0851   0.1417   1.0000
  13.250   1.4390   0.06122   0.05401  -0.0839   0.1246   1.0000
  13.500   1.4346   0.06473   0.05746  -0.0829   0.1095   1.0000
  13.750   1.4314   0.06819   0.06088  -0.0820   0.0967   1.0000
  14.000   1.4307   0.07147   0.06410  -0.0811   0.0856   1.0000
  14.250   1.4327   0.07448   0.06698  -0.0803   0.0757   1.0000
  14.500   1.4328   0.07788   0.07057  -0.0798   0.0681   1.0000
  14.750   1.4384   0.08074   0.07346  -0.0790   0.0609   1.0000
  15.000   1.4406   0.08387   0.07660  -0.0787   0.0553   1.0000
  15.250   1.4501   0.08667   0.07954  -0.0777   0.0500   1.0000
  15.500   1.4503   0.09018   0.08323  -0.0776   0.0463   1.0000
  15.750   1.4629   0.09283   0.08581  -0.0768   0.0419   1.0000
  16.000   1.4595   0.09714   0.09049  -0.0770   0.0403   1.0000
  16.250   1.4550   0.10160   0.09525  -0.0775   0.0385   1.0000
  16.500   1.4502   0.10603   0.09991  -0.0782   0.0370   1.0000
  16.750   1.4454   0.11030   0.10437  -0.0793   0.0353   1.0000
  17.000   1.4503   0.11397   0.10802  -0.0796   0.0332   1.0000
  17.250   1.4364   0.12017   0.11449  -0.0816   0.0327   1.0000
  17.500   1.4210   0.12638   0.12100  -0.0843   0.0326   1.0000
  17.750   1.4038   0.13318   0.12808  -0.0877   0.0325   1.0000
  18.000   1.3866   0.14035   0.13552  -0.0915   0.0325   1.0000
  18.250   1.3685   0.14800   0.14341  -0.0961   0.0326   1.0000
  18.500   1.3492   0.15635   0.15197  -0.1014   0.0327   1.0000
  18.750   1.3312   0.16488   0.16071  -0.1071   0.0329   1.0000
<< Back to EPPLER 396 AIRFOIL (e396-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 396 AIRFOIL (e396-il)