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EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il)
Reynolds number: 100,000
Max Cl/Cd: 33.52 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ea61012-il-100000.txt
Download as CSV file: xf-ea61012-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER EA 6(-1)-012 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5380   0.10183   0.09679  -0.0038   1.0000   0.2132
 -10.000  -0.4958   0.09904   0.09394  -0.0011   1.0000   0.2161
  -9.750  -0.8636   0.06041   0.05427  -0.0239   1.0000   0.1171
  -9.500  -0.8649   0.05602   0.04968  -0.0230   1.0000   0.1156
  -9.250  -0.8669   0.05177   0.04515  -0.0217   1.0000   0.1142
  -9.000  -0.8654   0.04772   0.04073  -0.0203   1.0000   0.1128
  -8.750  -0.8599   0.04388   0.03650  -0.0189   1.0000   0.1113
  -8.500  -0.8493   0.04076   0.03300  -0.0175   1.0000   0.1111
  -8.250  -0.8345   0.03830   0.03024  -0.0164   1.0000   0.1122
  -8.000  -0.8178   0.03604   0.02766  -0.0153   1.0000   0.1134
  -7.750  -0.7992   0.03388   0.02519  -0.0142   1.0000   0.1143
  -7.500  -0.7791   0.03193   0.02297  -0.0133   1.0000   0.1152
  -7.250  -0.7586   0.03039   0.02113  -0.0123   1.0000   0.1171
  -7.000  -0.7382   0.02919   0.01962  -0.0111   1.0000   0.1190
  -6.750  -0.7153   0.02720   0.01766  -0.0107   1.0000   0.1211
  -6.500  -0.6940   0.02590   0.01638  -0.0097   1.0000   0.1233
  -6.250  -0.6759   0.02489   0.01535  -0.0083   1.0000   0.1258
  -6.000  -0.6619   0.02410   0.01451  -0.0061   1.0000   0.1289
  -5.750  -0.6498   0.02345   0.01376  -0.0037   1.0000   0.1321
  -5.500  -0.6359   0.02261   0.01290  -0.0015   1.0000   0.1352
  -5.250  -0.6214   0.02179   0.01221   0.0004   1.0000   0.1388
  -5.000  -0.6069   0.02117   0.01161   0.0024   1.0000   0.1431
  -4.750  -0.5917   0.02066   0.01104   0.0043   1.0000   0.1486
  -4.500  -0.5777   0.01994   0.01051   0.0062   1.0000   0.1558
  -4.250  -0.5625   0.01943   0.01003   0.0080   1.0000   0.1652
  -4.000  -0.5484   0.01880   0.00953   0.0099   1.0000   0.1784
  -3.750  -0.5272   0.01785   0.00891   0.0104   0.9973   0.2122
  -3.500  -0.5159   0.01539   0.00865   0.0123   0.9913   0.5128
  -3.250  -0.4895   0.01593   0.00954   0.0137   0.9847   0.6546
  -3.000  -0.4595   0.01664   0.01026   0.0143   0.9788   0.6978
  -2.750  -0.4319   0.01726   0.01083   0.0151   0.9729   0.7278
  -2.500  -0.4002   0.01788   0.01138   0.0151   0.9684   0.7542
  -2.250  -0.3783   0.01825   0.01172   0.0169   0.9621   0.7748
  -2.000  -0.3511   0.01878   0.01227   0.0184   0.9568   0.7986
  -1.750  -0.3249   0.01937   0.01288   0.0205   0.9528   0.8267
  -1.500  -0.3075   0.01975   0.01327   0.0239   0.9469   0.8507
  -1.250  -0.2742   0.02025   0.01376   0.0245   0.9424   0.8736
  -1.000  -0.2189   0.02075   0.01420   0.0205   0.9386   0.8896
  -0.750  -0.1686   0.02093   0.01433   0.0165   0.9312   0.8972
  -0.500  -0.1177   0.02090   0.01424   0.0120   0.9238   0.9048
  -0.250  -0.0393   0.02100   0.01430   0.0026   0.9211   0.9076
   0.000   0.0003   0.02106   0.01437  -0.0001   0.9139   0.9139
   0.250   0.0399   0.02100   0.01430  -0.0027   0.9076   0.9211
   0.500   0.1175   0.02091   0.01425  -0.0119   0.9049   0.9239
   0.750   0.1687   0.02093   0.01433  -0.0165   0.8972   0.9312
   1.000   0.2188   0.02074   0.01420  -0.0204   0.8895   0.9385
   1.250   0.2741   0.02025   0.01376  -0.0244   0.8736   0.9423
   1.500   0.3077   0.01975   0.01327  -0.0239   0.8507   0.9470
   1.750   0.3248   0.01936   0.01286  -0.0204   0.8263   0.9528
   2.000   0.3512   0.01877   0.01226  -0.0184   0.7986   0.9568
   2.250   0.3783   0.01824   0.01171  -0.0168   0.7747   0.9621
   2.500   0.4004   0.01788   0.01138  -0.0151   0.7544   0.9684
   2.750   0.4320   0.01727   0.01084  -0.0151   0.7280   0.9730
   3.000   0.4595   0.01664   0.01025  -0.0143   0.6977   0.9788
   3.250   0.4895   0.01593   0.00954  -0.0137   0.6539   0.9847
   3.500   0.5158   0.01539   0.00863  -0.0122   0.5097   0.9913
   3.750   0.5273   0.01784   0.00891  -0.0104   0.2124   0.9973
   4.000   0.5483   0.01880   0.00953  -0.0099   0.1786   1.0000
   4.250   0.5625   0.01942   0.01002  -0.0080   0.1653   1.0000
   4.500   0.5777   0.01993   0.01050  -0.0062   0.1557   1.0000
   4.750   0.5916   0.02067   0.01104  -0.0043   0.1487   1.0000
   5.000   0.6068   0.02117   0.01161  -0.0024   0.1431   1.0000
   5.250   0.6214   0.02179   0.01221  -0.0004   0.1388   1.0000
   5.500   0.6358   0.02261   0.01290   0.0015   0.1351   1.0000
   5.750   0.6497   0.02344   0.01375   0.0037   0.1321   1.0000
   6.000   0.6619   0.02410   0.01450   0.0061   0.1289   1.0000
   6.250   0.6758   0.02488   0.01535   0.0083   0.1258   1.0000
   6.500   0.6939   0.02590   0.01638   0.0098   0.1233   1.0000
   6.750   0.7153   0.02719   0.01765   0.0107   0.1212   1.0000
   7.000   0.7381   0.02918   0.01961   0.0112   0.1189   1.0000
   7.250   0.7586   0.03038   0.02112   0.0123   0.1171   1.0000
   7.500   0.7792   0.03194   0.02297   0.0133   0.1152   1.0000
   7.750   0.7992   0.03387   0.02519   0.0142   0.1143   1.0000
   8.000   0.8178   0.03604   0.02767   0.0153   0.1134   1.0000
   8.250   0.8345   0.03830   0.03024   0.0164   0.1122   1.0000
   8.500   0.8495   0.04072   0.03294   0.0175   0.1109   1.0000
   8.750   0.8599   0.04389   0.03651   0.0189   0.1113   1.0000
   9.000   0.8655   0.04771   0.04073   0.0203   0.1127   1.0000
   9.250   0.8669   0.05179   0.04517   0.0217   0.1142   1.0000
   9.500   0.8651   0.05603   0.04969   0.0230   0.1156   1.0000
   9.750   0.8635   0.06042   0.05428   0.0239   0.1170   1.0000
  10.000   0.6146   0.09997   0.09456   0.0000   0.2148   1.0000
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