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EPPLER 1230 AIRFOIL (e1230-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1230 AIRFOIL (e1230-il)
Reynolds number: 50,000
Max Cl/Cd: 16.72 at α=0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1230-il-50000.txt
Download as CSV file: xf-e1230-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1230 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2744   0.10889   0.10265  -0.0228   1.0000   0.3294
  -8.750  -0.2679   0.10636   0.10019  -0.0212   1.0000   0.3446
  -8.500  -0.3557   0.09090   0.08496  -0.0362   1.0000   0.1895
  -8.250  -0.5173   0.07912   0.07336  -0.0409   1.0000   0.1449
  -8.000  -0.5256   0.07630   0.07058  -0.0382   1.0000   0.1418
  -7.750  -0.5425   0.07369   0.06803  -0.0353   1.0000   0.1399
  -7.500  -0.6105   0.06830   0.06201  -0.0322   1.0000   0.1279
  -7.250  -0.6165   0.06541   0.05906  -0.0301   1.0000   0.1273
  -7.000  -0.6229   0.06252   0.05603  -0.0281   1.0000   0.1269
  -6.750  -0.6270   0.05956   0.05284  -0.0264   1.0000   0.1266
  -6.500  -0.6278   0.05661   0.04957  -0.0250   1.0000   0.1266
  -6.250  -0.6248   0.05372   0.04629  -0.0237   1.0000   0.1268
  -6.000  -0.5899   0.04954   0.04136  -0.0277   0.9903   0.1283
  -5.750  -0.5416   0.04670   0.03853  -0.0330   0.9775   0.1374
  -5.500  -0.5009   0.04355   0.03469  -0.0366   0.9636   0.1460
  -5.250  -0.4562   0.04142   0.03220  -0.0403   0.9493   0.1611
  -5.000  -0.4120   0.03929   0.03008  -0.0438   0.9356   0.1818
  -4.750  -0.3688   0.03763   0.02841  -0.0468   0.9214   0.2100
  -4.500  -0.3275   0.03607   0.02698  -0.0492   0.9071   0.2437
  -4.250  -0.2885   0.03471   0.02578  -0.0511   0.8929   0.2852
  -4.000  -0.2491   0.03323   0.02469  -0.0530   0.8795   0.3449
  -3.750  -0.2103   0.03156   0.02437  -0.0533   0.8681   0.4813
  -3.500  -0.1992   0.03296   0.02640  -0.0462   0.8525   0.6297
  -3.250  -0.1856   0.03398   0.02738  -0.0399   0.8380   0.7106
  -3.000  -0.1533   0.03462   0.02788  -0.0354   0.8268   0.7723
  -2.750  -0.1041   0.03539   0.02843  -0.0338   0.8147   0.8233
  -2.500   0.1904   0.03536   0.02730  -0.0688   0.8073   0.9137
  -2.250   0.3163   0.03356   0.02501  -0.0850   0.7922   0.9513
  -2.000   0.3850   0.03257   0.02373  -0.0931   0.7742   0.9737
  -1.750   0.4436   0.03182   0.02274  -0.0998   0.7567   0.9917
  -1.500   0.4822   0.03153   0.02228  -0.1032   0.7406   1.0000
  -1.250   0.4982   0.03172   0.02234  -0.1022   0.7274   1.0000
  -1.000   0.5163   0.03181   0.02229  -0.1013   0.7153   1.0000
  -0.750   0.5263   0.03246   0.02289  -0.0998   0.7019   1.0000
  -0.500   0.5434   0.03270   0.02299  -0.0987   0.6915   1.0000
  -0.250   0.5540   0.03332   0.02357  -0.0971   0.6796   1.0000
   0.000   0.5647   0.03401   0.02422  -0.0954   0.6693   1.0000
   0.250   0.5774   0.03454   0.02467  -0.0939   0.6592   1.0000
   0.500   0.5829   0.03552   0.02564  -0.0916   0.6494   1.0000
   0.750   0.5924   0.03622   0.02629  -0.0897   0.6402   1.0000
   1.000   0.5976   0.03721   0.02725  -0.0873   0.6319   1.0000
   1.250   0.5890   0.03880   0.02889  -0.0834   0.6228   1.0000
   1.500   0.6165   0.03864   0.02857  -0.0831   0.6168   1.0000
   1.750   0.5562   0.04235   0.03249  -0.0739   0.6072   1.0000
   2.000   0.5546   0.04322   0.03332  -0.0704   0.6006   1.0000
   2.250   0.5209   0.04518   0.03528  -0.0635   0.5949   1.0000
   2.500   0.3778   0.05196   0.04209  -0.0483   0.5887   1.0000
   2.750   0.3815   0.05408   0.04412  -0.0474   0.5830   1.0000
   3.000   0.4378   0.05339   0.04327  -0.0491   0.5772   1.0000
   3.250   0.3789   0.05954   0.04941  -0.0459   0.5730   1.0000
   3.500   0.3687   0.06305   0.05287  -0.0453   0.5712   1.0000
   3.750   0.3647   0.06628   0.05606  -0.0451   0.5706   1.0000
   4.000   0.3634   0.06941   0.05914  -0.0451   0.5708   1.0000
   4.250   0.3659   0.07253   0.06221  -0.0455   0.5728   1.0000
   4.500   0.3832   0.07551   0.06514  -0.0467   0.5766   1.0000
   4.750   0.2680   0.08502   0.07495  -0.0483   0.6979   1.0000
   5.000   0.2810   0.08733   0.07719  -0.0484   0.6905   1.0000
   5.250   0.2972   0.08912   0.07892  -0.0485   0.6788   1.0000
   5.500   0.3356   0.09349   0.08319  -0.0516   0.6745   1.0000
   5.750   0.3202   0.09316   0.08285  -0.0482   0.6615   1.0000
   6.000   0.3623   0.09755   0.08717  -0.0514   0.6556   1.0000
   6.250   0.3472   0.09744   0.08703  -0.0483   0.6427   1.0000
   6.500   0.3853   0.10151   0.09104  -0.0509   0.6366   1.0000
   6.750   0.3711   0.10185   0.09137  -0.0483   0.6262   1.0000
   7.000   0.3990   0.10497   0.09446  -0.0497   0.6189   1.0000
   7.250   0.3988   0.10669   0.09615  -0.0487   0.6109   1.0000
   7.500   0.4186   0.10909   0.09852  -0.0492   0.6009   1.0000
   7.750   0.4458   0.11320   0.10260  -0.0508   0.5963   1.0000
   8.000   0.4357   0.11332   0.10272  -0.0488   0.5844   1.0000
   8.250   0.4684   0.11741   0.10680  -0.0507   0.5791   1.0000
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