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EPPLER E852 AIRFOIL (e852-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E852 AIRFOIL (e852-il)
Reynolds number: 100,000
Max Cl/Cd: 58.64 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e852-il-100000.txt
Download as CSV file: xf-e852-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E852 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4178   0.10293   0.09807  -0.0405   1.0000   0.0908
  -9.000  -0.4352   0.10015   0.09543  -0.0436   1.0000   0.0942
  -8.750  -0.4591   0.09755   0.09298  -0.0467   1.0000   0.0949
  -8.500  -0.4326   0.09323   0.08863  -0.0416   1.0000   0.0990
  -8.250  -0.4349   0.09056   0.08603  -0.0406   1.0000   0.1027
  -8.000  -0.4519   0.08810   0.08370  -0.0407   1.0000   0.1060
  -7.750  -0.4812   0.08624   0.08200  -0.0399   1.0000   0.1072
  -7.500  -0.5114   0.08306   0.07895  -0.0425   1.0000   0.1077
  -7.250  -0.5428   0.08005   0.07587  -0.0451   1.0000   0.1083
  -7.000  -0.5155   0.07796   0.07397  -0.0356   1.0000   0.1129
  -6.750  -0.5259   0.07502   0.07108  -0.0355   1.0000   0.1161
  -6.500  -0.5609   0.07016   0.06588  -0.0441   1.0000   0.1222
  -5.750  -0.5438   0.06075   0.05665  -0.0387   1.0000   0.1411
  -5.500  -0.5387   0.05724   0.05304  -0.0396   1.0000   0.1534
  -5.250  -0.5298   0.05406   0.04978  -0.0399   1.0000   0.1676
  -5.000  -0.4723   0.03735   0.03076  -0.0498   1.0000   0.0501
  -4.750  -0.4432   0.03482   0.02723  -0.0492   1.0000   0.0429
  -4.500  -0.4183   0.03094   0.02296  -0.0495   1.0000   0.0418
  -4.250  -0.3930   0.02847   0.02007  -0.0492   1.0000   0.0417
  -4.000  -0.3687   0.02619   0.01758  -0.0492   1.0000   0.0470
  -3.750  -0.3444   0.02485   0.01604  -0.0485   1.0000   0.0533
  -3.500  -0.3191   0.02291   0.01407  -0.0479   1.0000   0.0603
  -3.250  -0.2946   0.02179   0.01294  -0.0475   1.0000   0.0824
  -3.000  -0.2670   0.01996   0.01151  -0.0481   1.0000   0.1395
  -2.750  -0.2424   0.01809   0.01165  -0.0479   1.0000   0.5657
  -2.500  -0.2167   0.01869   0.01214  -0.0470   0.9963   0.6288
  -2.250  -0.1895   0.01925   0.01263  -0.0464   0.9920   0.6736
  -2.000  -0.1643   0.01960   0.01294  -0.0457   0.9872   0.7053
  -1.750  -0.1374   0.02005   0.01331  -0.0452   0.9829   0.7399
  -1.500  -0.1163   0.02029   0.01356  -0.0435   0.9782   0.7703
  -1.250  -0.0917   0.02044   0.01363  -0.0431   0.9733   0.7946
  -1.000  -0.0612   0.02069   0.01380  -0.0439   0.9693   0.8134
  -0.750  -0.0378   0.02071   0.01369  -0.0437   0.9640   0.8328
  -0.500  -0.0103   0.02083   0.01377  -0.0440   0.9592   0.8536
  -0.250   0.0181   0.02098   0.01390  -0.0446   0.9546   0.8791
   0.000   0.0426   0.02095   0.01389  -0.0445   0.9490   0.9148
   0.250   0.0952   0.02105   0.01394  -0.0507   0.9445   1.0000
   0.500   0.1325   0.02138   0.01412  -0.0545   0.9391   1.0000
   0.750   0.1718   0.02180   0.01440  -0.0583   0.9337   1.0000
   1.000   0.2123   0.02230   0.01479  -0.0620   0.9290   1.0000
   1.250   0.2408   0.02268   0.01510  -0.0633   0.9221   1.0000
   1.500   0.2828   0.02320   0.01554  -0.0668   0.9174   1.0000
   1.750   0.3065   0.02360   0.01591  -0.0670   0.9094   1.0000
   2.000   0.3486   0.02406   0.01636  -0.0703   0.9043   1.0000
   2.250   0.3718   0.02449   0.01682  -0.0702   0.8955   1.0000
   2.500   0.4118   0.02490   0.01727  -0.0729   0.8894   1.0000
   2.750   0.4402   0.02528   0.01769  -0.0737   0.8803   1.0000
   3.000   0.4705   0.02567   0.01814  -0.0746   0.8713   1.0000
   3.250   0.5163   0.02584   0.01842  -0.0779   0.8646   1.0000
   3.500   0.5455   0.02609   0.01884  -0.0783   0.8534   1.0000
   3.750   0.5794   0.02623   0.01911  -0.0794   0.8422   1.0000
   4.000   0.6194   0.02608   0.01915  -0.0811   0.8306   1.0000
   4.250   0.6699   0.02528   0.01859  -0.0836   0.8171   1.0000
   4.500   0.7569   0.02228   0.01609  -0.0899   0.8006   1.0000
   4.750   0.8176   0.01994   0.01410  -0.0920   0.7800   1.0000
   5.000   0.8670   0.01771   0.01216  -0.0920   0.7480   1.0000
   5.250   0.8964   0.01635   0.01098  -0.0892   0.7003   1.0000
   5.500   0.9118   0.01555   0.01000  -0.0840   0.5791   1.0000
   5.750   0.8993   0.01797   0.01033  -0.0758   0.2693   1.0000
   6.000   0.8929   0.02093   0.01206  -0.0704   0.1415   1.0000
   6.250   0.9006   0.02306   0.01383  -0.0670   0.0961   1.0000
   6.500   0.9121   0.02545   0.01603  -0.0645   0.0659   1.0000
   6.750   0.9337   0.02766   0.01826  -0.0630   0.0457   1.0000
   7.000   0.9682   0.03105   0.02170  -0.0635   0.0375   1.0000
   7.250   1.0016   0.03424   0.02518  -0.0636   0.0340   1.0000
   7.500   1.0252   0.03754   0.02861  -0.0633   0.0302   1.0000
   7.750   1.0427   0.04103   0.03260  -0.0615   0.0282   1.0000
   8.000   1.0575   0.04394   0.03603  -0.0591   0.0271   1.0000
   8.250   1.0686   0.04789   0.04044  -0.0566   0.0275   1.0000
   8.500   1.0749   0.05222   0.04517  -0.0540   0.0281   1.0000
   8.750   1.0816   0.05841   0.05161  -0.0524   0.0295   1.0000
   9.000   1.0767   0.06065   0.05473  -0.0465   0.0336   1.0000
   9.250   1.0662   0.06565   0.06012  -0.0428   0.0365   1.0000
   9.500   1.0581   0.07016   0.06482  -0.0402   0.0386   1.0000
   9.750   0.9597   0.08600   0.08198  -0.0325   0.1167   1.0000
  10.000   0.9288   0.09033   0.08643  -0.0310   0.1167   1.0000
  10.250   0.9065   0.09585   0.09203  -0.0313   0.1168   1.0000
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