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E71 (5.15%) (e71-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: E71 (5.15%) (e71-il)
Reynolds number: 500,000
Max Cl/Cd: 162.29 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e71-il-500000.txt
Download as CSV file: xf-e71-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E71  (5.15%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3892   0.10571   0.10357  -0.0184   1.0000   0.0090
  -7.750  -0.3912   0.10350   0.10138  -0.0174   1.0000   0.0090
  -7.500  -0.3949   0.10148   0.09939  -0.0161   1.0000   0.0090
  -7.000  -0.3902   0.09307   0.09104  -0.0189   0.9969   0.0095
  -6.750  -0.3722   0.08930   0.08726  -0.0228   0.9946   0.0097
  -6.500  -0.3551   0.08564   0.08360  -0.0263   0.9912   0.0101
  -6.250  -0.3332   0.08160   0.07956  -0.0318   0.9875   0.0107
  -6.000  -0.3059   0.07716   0.07511  -0.0391   0.9848   0.0112
  -5.750  -0.2738   0.07239   0.07032  -0.0481   0.9830   0.0127
  -5.500  -0.2453   0.06795   0.06584  -0.0563   0.9763   0.0146
  -5.250  -0.2014   0.06315   0.06100  -0.0675   0.9736   0.0153
  -5.000  -0.1566   0.05745   0.05523  -0.0790   0.9719   0.0155
  -2.500   0.2834   0.01415   0.00898  -0.1455   0.9605   0.0266
  -2.250   0.3218   0.01262   0.00728  -0.1476   0.9600   0.0228
  -2.000   0.3604   0.01150   0.00600  -0.1497   0.9594   0.0221
  -1.750   0.3995   0.01077   0.00519  -0.1520   0.9585   0.0232
  -1.500   0.4316   0.01032   0.00469  -0.1529   0.9546   0.0245
  -1.250   0.4655   0.00991   0.00425  -0.1541   0.9508   0.0257
  -1.000   0.5031   0.00952   0.00375  -0.1559   0.9484   0.0271
  -0.750   0.5417   0.00916   0.00342  -0.1581   0.9467   0.0389
  -0.500   0.5830   0.00837   0.00324  -0.1613   0.9456   0.2593
  -0.250   0.6144   0.00774   0.00323  -0.1624   0.9399   0.4895
   0.000   0.6430   0.00669   0.00323  -0.1622   0.9365   0.8787
   0.250   0.6777   0.00629   0.00287  -0.1632   0.9333   1.0000
   0.500   0.7082   0.00614   0.00269  -0.1636   0.9258   1.0000
   0.750   0.7454   0.00592   0.00244  -0.1653   0.9200   1.0000
   1.000   0.7770   0.00581   0.00232  -0.1659   0.9102   1.0000
   1.250   0.8114   0.00570   0.00219  -0.1671   0.9000   1.0000
   1.500   0.8455   0.00562   0.00208  -0.1682   0.8862   1.0000
   1.750   0.8781   0.00560   0.00202  -0.1690   0.8688   1.0000
   2.000   0.9082   0.00566   0.00204  -0.1692   0.8474   1.0000
   2.250   0.9364   0.00577   0.00207  -0.1690   0.8208   1.0000
   2.500   0.9627   0.00596   0.00214  -0.1684   0.7880   1.0000
   2.750   0.9873   0.00622   0.00226  -0.1675   0.7490   1.0000
   3.000   1.0094   0.00662   0.00242  -0.1660   0.6933   1.0000
   3.250   1.0247   0.00749   0.00274  -0.1632   0.5798   1.0000
   3.500   1.0401   0.00856   0.00317  -0.1608   0.4543   1.0000
   3.750   1.0601   0.00941   0.00359  -0.1594   0.3640   1.0000
   4.000   1.0788   0.01052   0.00410  -0.1580   0.2512   1.0000
   4.250   1.0980   0.01167   0.00467  -0.1567   0.1463   1.0000
   4.500   1.1173   0.01291   0.00536  -0.1554   0.0550   1.0000
   4.750   1.1380   0.01406   0.00624  -0.1539   0.0113   1.0000
   5.000   1.1621   0.01469   0.00704  -0.1530   0.0098   1.0000
   5.250   1.1854   0.01542   0.00790  -0.1519   0.0093   1.0000
   5.500   1.2077   0.01633   0.00894  -0.1506   0.0091   1.0000
   5.750   1.2287   0.01744   0.01018  -0.1491   0.0090   1.0000
   6.000   1.2486   0.01879   0.01169  -0.1474   0.0091   1.0000
   6.250   1.2681   0.02047   0.01353  -0.1455   0.0094   1.0000
   6.500   1.2887   0.02243   0.01569  -0.1439   0.0095   1.0000
   6.750   1.3112   0.02372   0.01711  -0.1428   0.0089   1.0000
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