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GOE 509 AIRFOIL (goe509-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 509 AIRFOIL (goe509-il)
Reynolds number: 50,000
Max Cl/Cd: 35.32 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe509-il-50000-n5.txt
Download as CSV file: xf-goe509-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 509 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3428   0.10921   0.10314  -0.0308   1.0000   0.1114
  -8.000  -0.3622   0.10832   0.10240  -0.0288   1.0000   0.1116
  -7.750  -0.3815   0.10726   0.10145  -0.0272   1.0000   0.1118
  -7.500  -0.3976   0.10587   0.10015  -0.0260   1.0000   0.1120
  -6.750  -0.3825   0.09150   0.08573  -0.0300   0.9911   0.0797
  -6.500  -0.3633   0.08878   0.08308  -0.0288   0.9863   0.0907
  -6.000  -0.3313   0.07532   0.06920  -0.0430   0.9663   0.0656
  -5.750  -0.3129   0.07127   0.06506  -0.0452   0.9582   0.0646
  -5.500  -0.2914   0.06666   0.06025  -0.0485   0.9500   0.0639
  -5.250  -0.2744   0.06208   0.05542  -0.0508   0.9397   0.0646
  -5.000  -0.2545   0.05687   0.04976  -0.0533   0.9303   0.0657
  -4.750  -0.2317   0.05226   0.04465  -0.0550   0.9215   0.0660
  -4.500  -0.2141   0.04857   0.04057  -0.0548   0.9109   0.0667
  -4.250  -0.1897   0.04688   0.03893  -0.0552   0.9011   0.0703
  -4.000  -0.1603   0.04383   0.03542  -0.0563   0.8925   0.0738
  -3.750  -0.1388   0.04091   0.03198  -0.0555   0.8814   0.0753
  -3.500  -0.1134   0.03835   0.02871  -0.0550   0.8718   0.0793
  -3.250  -0.0808   0.03609   0.02607  -0.0559   0.8649   0.0830
  -3.000  -0.0568   0.03463   0.02434  -0.0552   0.8553   0.0859
  -2.750  -0.0180   0.03302   0.02228  -0.0568   0.8496   0.0927
  -2.250   0.0499   0.03037   0.01911  -0.0583   0.8332   0.1029
  -2.000   0.0874   0.02946   0.01784  -0.0595   0.8240   0.1117
  -1.750   0.1447   0.02811   0.01639  -0.0644   0.8194   0.1234
  -1.500   0.1782   0.02743   0.01567  -0.0652   0.8086   0.1385
  -1.250   0.2231   0.02646   0.01478  -0.0677   0.8017   0.1640
  -1.000   0.2531   0.02588   0.01453  -0.0677   0.7905   0.2145
  -0.750   0.2818   0.02548   0.01427  -0.0674   0.7792   0.3077
  -0.500   0.4354   0.02268   0.01320  -0.0917   0.7758   1.0000
  -0.250   0.4654   0.02265   0.01284  -0.0916   0.7623   1.0000
   0.000   0.4955   0.02260   0.01251  -0.0915   0.7488   1.0000
   0.250   0.5254   0.02256   0.01221  -0.0914   0.7351   1.0000
   0.500   0.5537   0.02256   0.01197  -0.0910   0.7208   1.0000
   0.750   0.5778   0.02265   0.01187  -0.0899   0.7050   1.0000
   1.000   0.6015   0.02277   0.01182  -0.0888   0.6894   1.0000
   1.250   0.6247   0.02292   0.01181  -0.0876   0.6741   1.0000
   1.500   0.6475   0.02311   0.01184  -0.0864   0.6592   1.0000
   1.750   0.6700   0.02332   0.01192  -0.0852   0.6448   1.0000
   2.000   0.6926   0.02356   0.01203  -0.0841   0.6312   1.0000
   2.250   0.7163   0.02379   0.01213  -0.0831   0.6189   1.0000
   2.500   0.7427   0.02396   0.01216  -0.0826   0.6082   1.0000
   2.750   0.7620   0.02436   0.01250  -0.0811   0.5959   1.0000
   3.000   0.7836   0.02472   0.01278  -0.0799   0.5850   1.0000
   3.250   0.8102   0.02493   0.01287  -0.0795   0.5761   1.0000
   3.500   0.8280   0.02543   0.01337  -0.0778   0.5655   1.0000
   3.750   0.8510   0.02577   0.01366  -0.0768   0.5564   1.0000
   4.000   0.8730   0.02613   0.01396  -0.0757   0.5469   1.0000
   4.250   0.8916   0.02656   0.01437  -0.0740   0.5362   1.0000
   4.500   0.9133   0.02682   0.01457  -0.0727   0.5252   1.0000
   4.750   0.9353   0.02705   0.01472  -0.0714   0.5138   1.0000
   5.000   0.9513   0.02750   0.01515  -0.0693   0.5017   1.0000
   5.250   0.9702   0.02789   0.01553  -0.0677   0.4907   1.0000
   5.500   0.9946   0.02817   0.01572  -0.0669   0.4811   1.0000
   5.750   1.0100   0.02880   0.01642  -0.0649   0.4715   1.0000
   6.000   1.0327   0.02924   0.01687  -0.0639   0.4631   1.0000
   6.250   1.0502   0.02985   0.01753  -0.0623   0.4540   1.0000
   6.500   1.0703   0.03041   0.01812  -0.0610   0.4454   1.0000
   6.750   1.0894   0.03101   0.01877  -0.0596   0.4365   1.0000
   7.000   1.1063   0.03171   0.01957  -0.0580   0.4278   1.0000
   7.250   1.1265   0.03233   0.02024  -0.0568   0.4192   1.0000
   7.500   1.1401   0.03316   0.02120  -0.0547   0.4105   1.0000
   7.750   1.1608   0.03378   0.02187  -0.0536   0.4017   1.0000
   8.000   1.1692   0.03476   0.02304  -0.0508   0.3925   1.0000
   8.250   1.1917   0.03532   0.02362  -0.0499   0.3840   1.0000
   8.500   1.1934   0.03650   0.02501  -0.0463   0.3744   1.0000
   8.750   1.2152   0.03705   0.02559  -0.0453   0.3663   1.0000
   9.000   1.2130   0.03842   0.02721  -0.0413   0.3571   1.0000
   9.250   1.2341   0.03903   0.02787  -0.0403   0.3499   1.0000
   9.500   1.2262   0.04065   0.02972  -0.0357   0.3419   1.0000
   9.750   1.2472   0.04128   0.03042  -0.0348   0.3355   1.0000
  10.000   1.2359   0.04322   0.03258  -0.0302   0.3285   1.0000
  10.250   1.2450   0.04431   0.03378  -0.0280   0.3221   1.0000
  10.500   1.2445   0.04587   0.03551  -0.0250   0.3156   1.0000
  10.750   1.2377   0.04786   0.03766  -0.0218   0.3089   1.0000
  11.000   1.2525   0.04847   0.03834  -0.0203   0.3026   1.0000
  11.250   1.2254   0.05232   0.04240  -0.0167   0.2967   1.0000
  11.500   1.2520   0.05174   0.04187  -0.0157   0.2897   1.0000
  11.750   1.2147   0.05725   0.04758  -0.0130   0.2852   1.0000
  12.000   1.1688   0.06466   0.05513  -0.0118   0.2803   1.0000
  12.250   1.2162   0.06127   0.05182  -0.0104   0.2734   1.0000
  12.500   1.0234   0.09308   0.08350  -0.0185   0.2658   1.0000
  12.750   1.0544   0.09076   0.08135  -0.0163   0.2622   1.0000
  13.250   1.0170   0.10419   0.09487  -0.0197   0.2485   1.0000
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