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EPPLER 583 AIRFOIL (e583-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 583 AIRFOIL (e583-il)
Reynolds number: 50,000
Max Cl/Cd: 4.76 at α=9.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e583-il-50000.txt
Download as CSV file: xf-e583-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 583 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4530   0.13038   0.12552  -0.0151   1.0000   0.2380
  -7.750  -0.4349   0.12663   0.12175  -0.0145   0.9980   0.2506
  -7.500  -0.4720   0.12613   0.12137  -0.0132   0.9984   0.2553
  -7.250  -0.4571   0.12273   0.11797  -0.0126   0.9960   0.2707
  -7.000  -0.4500   0.11980   0.11507  -0.0117   0.9942   0.2855
  -6.750  -0.4519   0.11735   0.11266  -0.0101   0.9933   0.2980
  -6.500  -0.4905   0.11687   0.11231  -0.0041   0.9984   0.3042
  -6.250  -0.4932   0.11451   0.11001  -0.0006   1.0000   0.3161
  -6.000  -0.5476   0.11402   0.10966   0.0029   1.0000   0.3234
  -5.750  -0.5147   0.11044   0.10607   0.0052   1.0000   0.3468
  -5.500  -0.5658   0.10975   0.10549   0.0096   1.0000   0.3573
  -3.500  -0.5395   0.06192   0.05434  -0.0237   1.0000   0.1247
  -3.250  -0.5226   0.05857   0.05099  -0.0233   1.0000   0.1192
  -3.000  -0.4962   0.05554   0.04698  -0.0234   1.0000   0.1082
  -2.750  -0.4757   0.05301   0.04419  -0.0232   1.0000   0.1061
  -2.500  -0.4537   0.05110   0.04184  -0.0228   1.0000   0.1062
  -2.250  -0.4314   0.04966   0.03993  -0.0224   1.0000   0.1078
  -2.000  -0.4091   0.04825   0.03814  -0.0218   1.0000   0.1101
  -1.750  -0.3880   0.04695   0.03675  -0.0212   1.0000   0.1132
  -1.500  -0.3668   0.04625   0.03587  -0.0205   1.0000   0.1210
  -1.250  -0.3459   0.04550   0.03507  -0.0195   1.0000   0.1305
  -1.000  -0.3256   0.04509   0.03465  -0.0181   1.0000   0.1465
  -0.750  -0.3054   0.04476   0.03434  -0.0166   1.0000   0.1741
  -0.500  -0.2296   0.04519   0.03832  -0.0146   1.0000   1.0000
  -0.250  -0.2231   0.04523   0.03798  -0.0125   1.0000   1.0000
   0.000  -0.2166   0.04528   0.03773  -0.0105   1.0000   1.0000
   0.250  -0.2099   0.04537   0.03755  -0.0086   1.0000   1.0000
   0.500  -0.2023   0.04553   0.03747  -0.0069   1.0000   1.0000
   0.750  -0.1932   0.04580   0.03749  -0.0055   1.0000   1.0000
   1.000  -0.1758   0.04664   0.03807  -0.0058   0.9975   1.0000
   1.250  -0.1484   0.04829   0.03942  -0.0080   0.9904   1.0000
   1.500  -0.1170   0.05058   0.04141  -0.0109   0.9829   1.0000
   1.750  -0.0904   0.05198   0.04258  -0.0130   0.9728   1.0000
   2.000  -0.0667   0.05323   0.04364  -0.0144   0.9626   1.0000
   2.250  -0.0425   0.05473   0.04495  -0.0160   0.9531   1.0000
   2.500  -0.0080   0.05748   0.04746  -0.0194   0.9446   1.0000
   2.750   0.0118   0.05827   0.04812  -0.0201   0.9322   1.0000
   3.000   0.0315   0.05943   0.04916  -0.0208   0.9212   1.0000
   3.250   0.0647   0.06229   0.05183  -0.0239   0.9133   1.0000
   3.500   0.0838   0.06316   0.05261  -0.0244   0.9004   1.0000
   3.750   0.1009   0.06430   0.05366  -0.0247   0.8891   1.0000
   4.000   0.1390   0.06779   0.05701  -0.0285   0.8813   1.0000
   4.250   0.1502   0.06805   0.05721  -0.0278   0.8679   1.0000
   4.500   0.1664   0.06940   0.05850  -0.0281   0.8575   1.0000
   4.750   0.2022   0.07260   0.06160  -0.0314   0.8485   1.0000
   5.000   0.2116   0.07308   0.06206  -0.0306   0.8357   1.0000
   5.250   0.2294   0.07486   0.06380  -0.0312   0.8261   1.0000
   5.500   0.2610   0.07763   0.06651  -0.0337   0.8157   1.0000
   5.750   0.2692   0.07842   0.06729  -0.0329   0.8036   1.0000
   6.000   0.2940   0.08110   0.06994  -0.0346   0.7958   1.0000
   6.250   0.3146   0.08288   0.07172  -0.0356   0.7833   1.0000
   6.500   0.3237   0.08412   0.07298  -0.0351   0.7720   1.0000
   6.750   0.3612   0.08810   0.07693  -0.0385   0.7644   1.0000
   7.000   0.3652   0.08855   0.07741  -0.0373   0.7511   1.0000
   7.250   0.3754   0.09024   0.07912  -0.0372   0.7409   1.0000
   7.500   0.4118   0.09401   0.08291  -0.0402   0.7319   1.0000
   7.750   0.4117   0.09459   0.08354  -0.0389   0.7196   1.0000
   8.000   0.4265   0.09695   0.08593  -0.0395   0.7109   1.0000
   8.250   0.4547   0.09993   0.08895  -0.0415   0.6998   1.0000
   8.500   0.4557   0.10110   0.09016  -0.0406   0.6885   1.0000
   8.750   0.4814   0.10456   0.09367  -0.0425   0.6810   1.0000
   9.000   0.4934   0.10625   0.09545  -0.0427   0.6683   1.0000
   9.250   0.4973   0.10811   0.09735  -0.0424   0.6587   1.0000
   9.500   0.5361   0.11256   0.10187  -0.0454   0.6497   1.0000
   9.750   0.5284   0.11307   0.10245  -0.0441   0.6379   1.0000
  10.000   0.5413   0.11592   0.10537  -0.0449   0.6303   1.0000
  10.250   0.5660   0.11897   0.10852  -0.0463   0.6187   1.0000
  10.500   0.5626   0.12052   0.11014  -0.0459   0.6088   1.0000
  10.750   0.5916   0.12464   0.11435  -0.0479   0.6010   1.0000
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