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EPPLER 580 AIRFOIL (e580-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 580 AIRFOIL (e580-il)
Reynolds number: 50,000
Max Cl/Cd: 7.51 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e580-il-50000.txt
Download as CSV file: xf-e580-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 580 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  -0.3084   0.12221   0.11602  -0.0254   1.0000   0.3115
  -8.750  -0.3119   0.12014   0.11403  -0.0235   1.0000   0.3191
  -8.500  -0.3441   0.12131   0.11538  -0.0206   1.0000   0.3304
  -8.250  -0.3172   0.11658   0.11064  -0.0191   1.0000   0.3379
  -7.750  -0.3345   0.11408   0.10833  -0.0142   1.0000   0.3583
  -7.500  -0.3395   0.11219   0.10653  -0.0120   1.0000   0.3652
  -7.250  -0.3655   0.11245   0.10694  -0.0087   1.0000   0.3771
  -6.750  -0.3646   0.10794   0.10254  -0.0049   1.0000   0.3890
  -6.000  -0.5898   0.07004   0.06428  -0.0467   1.0000   0.1595
  -5.750  -0.5749   0.06352   0.05722  -0.0512   1.0000   0.1415
  -5.500  -0.5504   0.05754   0.05021  -0.0571   1.0000   0.1276
  -5.250  -0.5330   0.05405   0.04670  -0.0574   1.0000   0.1243
  -5.000  -0.5092   0.05053   0.04272  -0.0595   1.0000   0.1196
  -4.750  -0.4799   0.04744   0.03883  -0.0619   1.0000   0.1148
  -4.500  -0.4534   0.04576   0.03655  -0.0629   1.0000   0.1129
  -4.250  -0.4304   0.04409   0.03453  -0.0632   1.0000   0.1127
  -4.000  -0.4094   0.04232   0.03269  -0.0631   1.0000   0.1145
  -3.750  -0.3896   0.04119   0.03156  -0.0626   1.0000   0.1184
  -3.500  -0.3695   0.04040   0.03066  -0.0619   1.0000   0.1222
  -3.250  -0.3498   0.03982   0.02991  -0.0607   1.0000   0.1261
  -3.000  -0.3327   0.03927   0.02938  -0.0589   1.0000   0.1299
  -2.750  -0.2836   0.03960   0.02968  -0.0619   0.9868   0.1430
  -2.500  -0.2370   0.03893   0.02943  -0.0663   0.9756   0.1787
  -2.250  -0.2221   0.04039   0.03345  -0.0595   0.9678   0.6134
  -2.000  -0.2302   0.04268   0.03572  -0.0486   0.9620   0.6655
  -1.750  -0.2392   0.04423   0.03730  -0.0381   0.9558   0.7100
  -1.500  -0.2513   0.04504   0.03813  -0.0280   0.9515   0.7522
  -1.250  -0.2656   0.04552   0.03863  -0.0170   0.9478   0.7970
  -1.000  -0.2798   0.04537   0.03848  -0.0066   0.9439   0.8353
  -0.750  -0.2887   0.04490   0.03795   0.0012   0.9415   0.8701
  -0.500  -0.2846   0.04454   0.03746   0.0057   0.9380   0.8987
  -0.250  -0.2550   0.04476   0.03738   0.0038   0.9284   0.9110
   0.000  -0.2299   0.04495   0.03734   0.0017   0.9220   0.9182
   0.250  -0.1992   0.04529   0.03744  -0.0010   0.9125   0.9249
   0.500  -0.1619   0.04600   0.03790  -0.0050   0.9024   0.9314
   0.750  -0.1404   0.04624   0.03799  -0.0066   0.8971   0.9372
   1.000  -0.1029   0.04702   0.03858  -0.0106   0.8862   0.9436
   1.250  -0.0766   0.04759   0.03903  -0.0132   0.8802   0.9495
   1.500  -0.0391   0.04849   0.03975  -0.0173   0.8693   0.9560
   1.750   0.0080   0.04966   0.04077  -0.0229   0.8573   0.9625
   2.000   0.0310   0.05041   0.04145  -0.0253   0.8513   0.9695
   2.250   0.0748   0.05159   0.04252  -0.0306   0.8389   0.9770
   2.500   0.1046   0.05264   0.04353  -0.0341   0.8323   0.9880
   2.750   0.1291   0.05363   0.04444  -0.0365   0.8241   1.0000
   3.000   0.1434   0.05478   0.04554  -0.0378   0.8214   1.0000
   3.250   0.1490   0.05648   0.04725  -0.0390   0.8334   1.0000
   3.500   0.1797   0.05835   0.04906  -0.0429   0.8297   1.0000
   3.750   0.2095   0.06049   0.05115  -0.0467   0.8274   1.0000
   4.000   0.2382   0.06165   0.05227  -0.0494   0.8099   1.0000
   4.250   0.2724   0.06318   0.05375  -0.0529   0.7939   1.0000
   4.500   0.3065   0.06474   0.05527  -0.0563   0.7778   1.0000
   4.750   0.3647   0.06586   0.05632  -0.0614   0.7492   1.0000
   5.000   0.3770   0.06810   0.05858  -0.0633   0.7477   1.0000
   5.250   0.3942   0.07063   0.06114  -0.0659   0.7478   1.0000
   5.500   0.4280   0.07242   0.06293  -0.0691   0.7326   1.0000
   5.750   0.4499   0.07540   0.06595  -0.0725   0.7341   1.0000
   6.000   0.4783   0.07715   0.06772  -0.0749   0.7180   1.0000
   6.250   0.5314   0.07788   0.06844  -0.0784   0.6885   1.0000
   6.500   0.5374   0.08082   0.07145  -0.0799   0.6864   1.0000
   6.750   0.5561   0.08425   0.07494  -0.0829   0.6884   1.0000
   7.000   0.5795   0.08607   0.07679  -0.0846   0.6713   1.0000
   7.250   0.6083   0.08820   0.07897  -0.0870   0.6576   1.0000
   7.500   0.6231   0.09173   0.08258  -0.0894   0.6567   1.0000
   7.750   0.6616   0.09200   0.08288  -0.0904   0.6247   1.0000
   8.000   0.6662   0.09580   0.08675  -0.0921   0.6246   1.0000
   8.250   0.7190   0.09578   0.08680  -0.0937   0.5946   1.0000
   8.500   0.7111   0.09995   0.09103  -0.0946   0.5930   1.0000
   8.750   0.7142   0.10403   0.09519  -0.0962   0.5930   1.0000
   9.000   0.7343   0.10649   0.09773  -0.0975   0.5802   1.0000
   9.250   0.7804   0.10613   0.09742  -0.0977   0.5485   1.0000
   9.500   0.7700   0.11055   0.10191  -0.0985   0.5451   1.0000
   9.750   0.7701   0.11502   0.10648  -0.1000   0.5452   1.0000
  10.000   0.7849   0.11791   0.10946  -0.1011   0.5345   1.0000
  10.250   0.7958   0.12223   0.11388  -0.1029   0.5331   1.0000
  10.500   0.8451   0.11956   0.11128  -0.1008   0.4873   1.0000
  10.750   0.8764   0.12069   0.11253  -0.1009   0.4703   1.0000
  11.000   0.8520   0.12653   0.11843  -0.1022   0.4671   1.0000
  11.250   0.8454   0.13130   0.12327  -0.1037   0.4639   1.0000
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