Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 387 AIRFOIL (goe387-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 387 AIRFOIL (goe387-il)
Reynolds number: 50,000
Max Cl/Cd: 22.6 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe387-il-50000-n5.txt
Download as CSV file: xf-goe387-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 387 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.1551   0.11253   0.10588  -0.0584   0.9545   0.1462
  -8.750  -0.1610   0.11065   0.10403  -0.0652   0.9437   0.1494
  -8.500  -0.1604   0.10795   0.10135  -0.0695   0.9323   0.1498
  -8.250  -0.1401   0.10329   0.09669  -0.0731   0.9269   0.1504
  -7.750  -0.1115   0.08964   0.08283  -0.0818   0.9114   0.0902
  -7.250  -0.1122   0.07995   0.07307  -0.0919   0.8910   0.0811
  -7.000  -0.1083   0.07754   0.07066  -0.0918   0.8799   0.0799
  -6.750  -0.0959   0.07347   0.06652  -0.0954   0.8732   0.0780
  -6.500  -0.1014   0.07022   0.06321  -0.0960   0.8612   0.0763
  -6.250  -0.1085   0.06237   0.05489  -0.1020   0.8532   0.0722
  -6.000  -0.1107   0.06053   0.05297  -0.0998   0.8417   0.0718
  -5.750  -0.0946   0.05710   0.04932  -0.1014   0.8360   0.0711
  -5.500  -0.0937   0.05498   0.04702  -0.0995   0.8259   0.0706
  -5.250  -0.0797   0.05192   0.04365  -0.0999   0.8194   0.0701
  -5.000  -0.0663   0.04925   0.04063  -0.0995   0.8127   0.0700
  -4.750  -0.0586   0.04733   0.03839  -0.0977   0.8037   0.0702
  -4.500  -0.0353   0.04475   0.03533  -0.0981   0.7989   0.0709
  -4.250  -0.0265   0.04336   0.03358  -0.0957   0.7900   0.0715
  -4.000  -0.0057   0.04165   0.03145  -0.0950   0.7835   0.0720
  -3.750   0.0255   0.03974   0.02912  -0.0955   0.7794   0.0724
  -3.500   0.0342   0.03913   0.02834  -0.0927   0.7696   0.0727
  -3.250   0.0619   0.03789   0.02690  -0.0926   0.7640   0.0734
  -3.000   0.0976   0.03654   0.02534  -0.0935   0.7602   0.0744
  -2.750   0.1046   0.03641   0.02511  -0.0903   0.7496   0.0751
  -2.500   0.1371   0.03542   0.02396  -0.0907   0.7445   0.0772
  -2.250   0.1604   0.03491   0.02328  -0.0897   0.7375   0.0798
  -2.000   0.1817   0.03452   0.02276  -0.0884   0.7293   0.0824
  -1.750   0.2192   0.03364   0.02181  -0.0895   0.7248   0.0861
  -1.500   0.2338   0.03367   0.02176  -0.0874   0.7150   0.0888
  -1.250   0.2658   0.03309   0.02102  -0.0876   0.7089   0.0930
  -1.000   0.3063   0.03221   0.02007  -0.0889   0.7050   0.1013
  -0.750   0.3121   0.03262   0.02051  -0.0857   0.6932   0.1084
  -0.500   0.3498   0.03171   0.01972  -0.0867   0.6885   0.1384
  -0.250   0.3619   0.03167   0.02007  -0.0846   0.6784   0.2062
   0.000   0.3906   0.03066   0.01984  -0.0845   0.6723   0.4029
   0.250   0.4985   0.02853   0.01896  -0.0975   0.6693   1.0000
   0.500   0.5017   0.02938   0.01964  -0.0938   0.6580   1.0000
   0.750   0.5346   0.02923   0.01919  -0.0940   0.6524   1.0000
   1.000   0.5422   0.02998   0.01978  -0.0909   0.6420   1.0000
   1.250   0.5703   0.03002   0.01959  -0.0905   0.6356   1.0000
   1.500   0.5866   0.03051   0.01991  -0.0886   0.6271   1.0000
   1.750   0.6061   0.03086   0.02010  -0.0871   0.6188   1.0000
   2.000   0.6436   0.03057   0.01958  -0.0879   0.6144   1.0000
   2.250   0.6413   0.03179   0.02075  -0.0837   0.6023   1.0000
   2.500   0.6764   0.03157   0.02035  -0.0842   0.5974   1.0000
   2.750   0.6760   0.03279   0.02152  -0.0804   0.5860   1.0000
   3.000   0.7074   0.03272   0.02130  -0.0804   0.5806   1.0000
   3.250   0.7104   0.03388   0.02241  -0.0771   0.5704   1.0000
   3.500   0.7356   0.03408   0.02250  -0.0765   0.5641   1.0000
   4.000   0.7598   0.03561   0.02390  -0.0721   0.5478   1.0000
   4.250   0.7979   0.03530   0.02347  -0.0729   0.5439   1.0000
   4.500   0.7809   0.03750   0.02568  -0.0678   0.5321   1.0000
   4.750   0.8134   0.03739   0.02549  -0.0679   0.5279   1.0000
   5.250   0.8280   0.03981   0.02785  -0.0634   0.5120   1.0000
   5.500   0.8663   0.03934   0.02731  -0.0639   0.5091   1.0000
   6.000   0.8716   0.04267   0.03065  -0.0593   0.4933   1.0000
   6.500   0.8756   0.04661   0.03459  -0.0555   0.4776   1.0000
   6.750   0.9117   0.04603   0.03398  -0.0554   0.4753   1.0000
   7.250   0.9077   0.05104   0.03905  -0.0519   0.4593   1.0000
   7.750   0.9103   0.05573   0.04380  -0.0493   0.4431   1.0000
   8.250   0.9245   0.05909   0.04718  -0.0469   0.4268   1.0000
   8.500   0.9695   0.05675   0.04481  -0.0462   0.4246   1.0000
   9.000   0.9602   0.06275   0.05090  -0.0438   0.4058   1.0000
   9.500   0.9410   0.07075   0.05902  -0.0426   0.3864   1.0000
  10.000   0.9680   0.07321   0.06156  -0.0411   0.3748   1.0000
  11.000   0.9666   0.08541   0.07401  -0.0396   0.3413   1.0000
  11.250   0.9960   0.08444   0.07308  -0.0384   0.3390   1.0000
  11.750   0.9753   0.09386   0.08263  -0.0390   0.3198   1.0000
  12.250   0.9818   0.09948   0.08839  -0.0388   0.3057   1.0000
  12.750   0.9697   0.10832   0.09734  -0.0400   0.2901   1.0000
  13.250   0.9521   0.11850   0.10765  -0.0422   0.2758   1.0000
  13.500   0.9745   0.11856   0.10779  -0.0414   0.2733   1.0000
  14.000   0.9551   0.12929   0.11865  -0.0443   0.2599   1.0000
  14.250   0.9545   0.13316   0.12257  -0.0453   0.2547   1.0000
<< Back to GOE 387 AIRFOIL (goe387-il)

Polar data table (+)

Polar graphs


<< Back to GOE 387 AIRFOIL (goe387-il)