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

EPPLER 587 AIRFOIL (e587-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 587 AIRFOIL (e587-il)
Reynolds number: 200,000
Max Cl/Cd: 75.22 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e587-il-200000.txt
Download as CSV file: xf-e587-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 587 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.0716   0.10841   0.10504  -0.1005   0.9537   0.0475
 -11.500  -0.0868   0.10221   0.09887  -0.1067   0.9505   0.0492
 -11.250  -0.0812   0.09665   0.09331  -0.1095   0.9486   0.0500
 -11.000  -0.0582   0.09370   0.09034  -0.1104   0.9471   0.0512
 -10.750  -0.0428   0.08993   0.08656  -0.1129   0.9458   0.0536
 -10.500  -0.0401   0.08414   0.08077  -0.1179   0.9447   0.0567
 -10.250  -0.1597   0.09102   0.08760  -0.1110   0.9507   0.0506
 -10.000  -0.1403   0.08816   0.08472  -0.1130   0.9482   0.0520
  -9.750  -0.0262   0.06970   0.06635  -0.1268   0.9335   0.0644
  -9.500  -0.0346   0.06127   0.05792  -0.1348   0.9313   0.0670
  -9.250  -0.0444   0.05472   0.05136  -0.1389   0.9252   0.0658
  -9.000  -0.0736   0.04712   0.04363  -0.1460   0.9186   0.0667
  -8.750  -0.0854   0.04220   0.03860  -0.1490   0.9124   0.0654
  -8.500  -0.1202   0.03956   0.03571  -0.1487   0.9011   0.0675
  -8.250  -0.2002   0.02573   0.02058  -0.1415   0.8872   0.0291
  -8.000  -0.1862   0.02221   0.01671  -0.1421   0.8842   0.0289
  -7.750  -0.1899   0.02085   0.01513  -0.1381   0.8750   0.0289
  -7.500  -0.1866   0.03194   0.02557  -0.1429   0.8808   0.0287
  -7.250  -0.1741   0.03001   0.02337  -0.1410   0.8744   0.0285
  -7.000  -0.1463   0.02754   0.02057  -0.1416   0.8714   0.0279
  -6.750  -0.1141   0.02535   0.01806  -0.1425   0.8692   0.0276
  -6.500  -0.0995   0.02432   0.01686  -0.1401   0.8634   0.0280
  -6.250  -0.0779   0.02327   0.01565  -0.1389   0.8586   0.0286
  -6.000  -0.0485   0.02197   0.01426  -0.1392   0.8557   0.0300
  -5.750  -0.0179   0.02088   0.01320  -0.1400   0.8533   0.0328
  -5.500  -0.0054   0.02045   0.01278  -0.1376   0.8475   0.0368
  -5.250   0.0129   0.01990   0.01224  -0.1364   0.8424   0.0426
  -5.000   0.0401   0.01887   0.01128  -0.1370   0.8392   0.0561
  -4.750   0.0707   0.01760   0.01023  -0.1386   0.8368   0.1008
  -4.500   0.0826   0.01679   0.00995  -0.1372   0.8306   0.1946
  -4.250   0.1047   0.01524   0.00993  -0.1382   0.8263   0.5255
  -4.000   0.1354   0.01568   0.01039  -0.1380   0.8233   0.6055
  -3.750   0.1680   0.01616   0.01075  -0.1380   0.8210   0.6352
  -3.500   0.1859   0.01684   0.01138  -0.1356   0.8162   0.6519
  -3.250   0.2045   0.01741   0.01191  -0.1333   0.8110   0.6663
  -3.000   0.2325   0.01779   0.01218  -0.1326   0.8078   0.6806
  -2.750   0.2608   0.01818   0.01251  -0.1316   0.8054   0.6910
  -2.500   0.2837   0.01856   0.01285  -0.1301   0.8019   0.7005
  -2.250   0.2979   0.01904   0.01327  -0.1277   0.7956   0.7123
  -2.000   0.3156   0.01950   0.01376  -0.1242   0.7920   0.7205
  -1.750   0.3453   0.01965   0.01382  -0.1239   0.7896   0.7319
  -1.500   0.3761   0.01974   0.01383  -0.1238   0.7875   0.7383
  -1.250   0.3849   0.02008   0.01415  -0.1213   0.7798   0.7461
  -1.000   0.4090   0.02008   0.01411  -0.1203   0.7762   0.7495
  -0.750   0.4419   0.01994   0.01390  -0.1212   0.7738   0.7536
  -0.500   0.4811   0.01974   0.01358  -0.1238   0.7718   0.7582
  -0.250   0.4918   0.02006   0.01391  -0.1216   0.7644   0.7627
   0.000   0.5177   0.01999   0.01381  -0.1213   0.7605   0.7653
   0.250   0.5516   0.01981   0.01357  -0.1225   0.7579   0.7681
   0.500   0.5902   0.01960   0.01328  -0.1248   0.7557   0.7711
   0.750   0.6023   0.01996   0.01367  -0.1228   0.7483   0.7755
   1.000   0.6352   0.01983   0.01349  -0.1244   0.7444   0.7785
   1.250   0.6693   0.01960   0.01322  -0.1256   0.7416   0.7806
   1.500   0.7067   0.01939   0.01295  -0.1274   0.7393   0.7830
   1.750   0.7133   0.01981   0.01346  -0.1241   0.7312   0.7865
   2.000   0.7474   0.01964   0.01326  -0.1256   0.7274   0.7894
   2.250   0.7890   0.01937   0.01293  -0.1286   0.7247   0.7923
   2.500   0.8103   0.01951   0.01309  -0.1278   0.7189   0.7953
   2.750   0.8318   0.01952   0.01314  -0.1268   0.7131   0.7981
   3.000   0.8681   0.01925   0.01285  -0.1285   0.7097   0.8006
   3.250   0.9070   0.01902   0.01259  -0.1307   0.7066   0.8032
   3.500   0.9186   0.01933   0.01299  -0.1283   0.6983   0.8068
   3.750   0.9572   0.01908   0.01272  -0.1307   0.6942   0.8101
   4.000   0.9956   0.01876   0.01238  -0.1326   0.6910   0.8125
   4.250   1.0030   0.01904   0.01279  -0.1291   0.6822   0.8157
   4.500   1.0393   0.01873   0.01247  -0.1308   0.6776   0.8186
   4.750   1.0649   0.01871   0.01251  -0.1307   0.6711   0.8219
   5.000   1.0932   0.01863   0.01246  -0.1311   0.6642   0.8256
   5.250   1.1308   0.01826   0.01209  -0.1329   0.6597   0.8283
   5.500   1.1409   0.01842   0.01238  -0.1298   0.6507   0.8321
   5.750   1.1769   0.01811   0.01207  -0.1313   0.6449   0.8355
   6.000   1.1940   0.01821   0.01227  -0.1297   0.6359   0.8395
   6.250   1.2291   0.01792   0.01197  -0.1311   0.6290   0.8427
   6.500   1.2403   0.01798   0.01215  -0.1281   0.6192   0.8467
   6.750   1.2712   0.01775   0.01194  -0.1287   0.6112   0.8509
   7.000   1.2866   0.01781   0.01208  -0.1266   0.6004   0.8559
   7.250   1.3029   0.01781   0.01215  -0.1246   0.5899   0.8601
   7.500   1.3258   0.01771   0.01207  -0.1236   0.5796   0.8645
   7.750   1.3374   0.01778   0.01221  -0.1208   0.5674   0.8701
   8.000   1.3469   0.01794   0.01245  -0.1176   0.5546   0.8759
   8.250   1.3582   0.01812   0.01267  -0.1147   0.5412   0.8826
   8.500   1.3707   0.01835   0.01295  -0.1123   0.5268   0.8896
   8.750   1.3804   0.01863   0.01326  -0.1093   0.5118   0.8972
   9.000   1.3895   0.01901   0.01365  -0.1064   0.4957   0.9058
   9.250   1.3962   0.01947   0.01413  -0.1031   0.4788   0.9166
   9.500   1.3999   0.01997   0.01465  -0.0994   0.4618   0.9316
   9.750   1.4022   0.02045   0.01517  -0.0957   0.4442   0.9791
  10.250   1.4221   0.02261   0.01723  -0.0928   0.4026   1.0000
  10.500   1.4291   0.02390   0.01846  -0.0910   0.3820   1.0000
  10.750   1.4342   0.02532   0.01980  -0.0891   0.3619   1.0000
  11.000   1.4380   0.02686   0.02129  -0.0871   0.3421   1.0000
  11.250   1.4409   0.02849   0.02289  -0.0851   0.3225   1.0000
  11.500   1.4426   0.03026   0.02460  -0.0831   0.3035   1.0000
  11.750   1.4435   0.03214   0.02643  -0.0811   0.2852   1.0000
  12.000   1.4438   0.03415   0.02837  -0.0792   0.2674   1.0000
  12.250   1.4443   0.03622   0.03040  -0.0775   0.2500   1.0000
  12.500   1.4451   0.03836   0.03253  -0.0759   0.2330   1.0000
  12.750   1.4455   0.04061   0.03477  -0.0744   0.2167   1.0000
  13.000   1.4456   0.04298   0.03712  -0.0731   0.2009   1.0000
  13.250   1.4456   0.04544   0.03956  -0.0719   0.1859   1.0000
  13.500   1.4452   0.04803   0.04213  -0.0708   0.1712   1.0000
  13.750   1.4444   0.05075   0.04484  -0.0698   0.1571   1.0000
  14.000   1.4431   0.05361   0.04768  -0.0690   0.1436   1.0000
  14.250   1.4409   0.05665   0.05071  -0.0683   0.1307   1.0000
  14.500   1.4380   0.05987   0.05392  -0.0677   0.1186   1.0000
  14.750   1.4343   0.06329   0.05732  -0.0672   0.1071   1.0000
  15.000   1.4322   0.06663   0.06068  -0.0670   0.0959   1.0000
  15.250   1.4294   0.07015   0.06423  -0.0668   0.0857   1.0000
  15.500   1.4255   0.07388   0.06799  -0.0668   0.0768   1.0000
  15.750   1.4194   0.07799   0.07208  -0.0670   0.0695   1.0000
  16.000   1.4151   0.08196   0.07609  -0.0673   0.0626   1.0000
  16.250   1.4114   0.08595   0.08015  -0.0677   0.0565   1.0000
  16.500   1.4037   0.09051   0.08469  -0.0684   0.0519   1.0000
  16.750   1.4023   0.09436   0.08867  -0.0691   0.0469   1.0000
<< Back to EPPLER 587 AIRFOIL (e587-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 587 AIRFOIL (e587-il)