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

GOE 596 AIRFOIL (goe596-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 596 AIRFOIL (goe596-il)
Reynolds number: 50,000
Max Cl/Cd: 37 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe596-il-50000-n5.txt
Download as CSV file: xf-goe596-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 596 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3366   0.11225   0.10491  -0.0304   1.0000   0.1212
  -9.250  -0.3476   0.11115   0.10394  -0.0326   1.0000   0.1248
  -9.000  -0.3638   0.11037   0.10331  -0.0348   1.0000   0.1257
  -8.750  -0.3389   0.10432   0.09723  -0.0324   1.0000   0.1285
  -8.500  -0.3310   0.10125   0.09420  -0.0315   1.0000   0.1326
  -8.250  -0.3362   0.09934   0.09239  -0.0317   1.0000   0.1380
  -8.000  -0.3617   0.09900   0.09227  -0.0325   1.0000   0.1405
  -7.750  -0.3447   0.09453   0.08782  -0.0305   1.0000   0.1448
  -7.500  -0.3450   0.09229   0.08567  -0.0290   1.0000   0.1500
  -7.250  -0.3667   0.09135   0.08491  -0.0298   1.0000   0.1551
  -7.000  -0.3739   0.08872   0.08240  -0.0295   1.0000   0.1577
  -6.750  -0.3664   0.08582   0.07954  -0.0264   1.0000   0.1616
  -6.500  -0.3714   0.08375   0.07755  -0.0253   1.0000   0.1661
  -6.000  -0.3845   0.07918   0.07312  -0.0244   1.0000   0.1752
  -5.500  -0.3737   0.06702   0.06052  -0.0349   0.9997   0.0928
  -5.000  -0.3027   0.05501   0.04742  -0.0504   0.9797   0.0810
  -4.750  -0.2721   0.05135   0.04368  -0.0534   0.9712   0.0797
  -4.500  -0.2391   0.04771   0.03978  -0.0570   0.9621   0.0782
  -4.250  -0.2076   0.04436   0.03608  -0.0598   0.9517   0.0769
  -4.000  -0.1738   0.04129   0.03256  -0.0627   0.9421   0.0764
  -3.750  -0.1371   0.03860   0.02935  -0.0655   0.9330   0.0779
  -3.500  -0.1048   0.03638   0.02664  -0.0671   0.9221   0.0788
  -3.250  -0.0671   0.03424   0.02405  -0.0694   0.9139   0.0788
  -3.000  -0.0330   0.03248   0.02190  -0.0707   0.9034   0.0790
  -2.750   0.0020   0.03095   0.02000  -0.0720   0.8934   0.0795
  -2.500   0.0420   0.02961   0.01824  -0.0740   0.8855   0.0810
  -2.250   0.0727   0.02842   0.01702  -0.0747   0.8741   0.0843
  -2.000   0.1080   0.02744   0.01590  -0.0759   0.8646   0.0876
  -1.750   0.1444   0.02650   0.01477  -0.0770   0.8552   0.0901
  -1.500   0.1762   0.02582   0.01390  -0.0772   0.8438   0.0927
  -1.250   0.2116   0.02512   0.01304  -0.0781   0.8344   0.0962
  -1.000   0.2432   0.02451   0.01241  -0.0785   0.8238   0.1037
  -0.750   0.2715   0.02407   0.01191  -0.0784   0.8119   0.1134
  -0.500   0.3022   0.02357   0.01139  -0.0786   0.8014   0.1264
  -0.250   0.3330   0.02286   0.01097  -0.0789   0.7914   0.1668
   0.000   0.3865   0.02024   0.01063  -0.0832   0.7818   1.0000
   0.250   0.4153   0.02034   0.01038  -0.0829   0.7707   1.0000
   0.500   0.4435   0.02043   0.01021  -0.0826   0.7592   1.0000
   0.750   0.4679   0.02063   0.01020  -0.0817   0.7463   1.0000
   1.000   0.4933   0.02081   0.01020  -0.0810   0.7341   1.0000
   1.250   0.5203   0.02094   0.01016  -0.0805   0.7231   1.0000
   1.500   0.5476   0.02107   0.01014  -0.0800   0.7125   1.0000
   1.750   0.5714   0.02134   0.01031  -0.0791   0.7006   1.0000
   2.000   0.5970   0.02157   0.01043  -0.0785   0.6899   1.0000
   2.250   0.6256   0.02170   0.01045  -0.0781   0.6808   1.0000
   2.500   0.6487   0.02205   0.01075  -0.0773   0.6692   1.0000
   2.750   0.6738   0.02235   0.01099  -0.0766   0.6589   1.0000
   3.000   0.7021   0.02252   0.01110  -0.0763   0.6498   1.0000
   3.250   0.7247   0.02295   0.01153  -0.0753   0.6385   1.0000
   3.500   0.7497   0.02330   0.01186  -0.0747   0.6284   1.0000
   3.750   0.7769   0.02356   0.01211  -0.0742   0.6189   1.0000
   4.000   0.7989   0.02407   0.01266  -0.0733   0.6077   1.0000
   4.250   0.8239   0.02446   0.01306  -0.0726   0.5977   1.0000
   4.500   0.8495   0.02483   0.01346  -0.0720   0.5876   1.0000
   4.750   0.8708   0.02541   0.01412  -0.0710   0.5764   1.0000
   5.000   0.8955   0.02583   0.01459  -0.0703   0.5663   1.0000
   5.250   0.9198   0.02628   0.01510  -0.0695   0.5558   1.0000
   5.500   0.9402   0.02689   0.01585  -0.0683   0.5441   1.0000
   5.750   0.9629   0.02733   0.01636  -0.0673   0.5322   1.0000
   6.000   0.9872   0.02763   0.01671  -0.0663   0.5201   1.0000
   6.250   1.0092   0.02801   0.01720  -0.0651   0.5072   1.0000
   6.500   1.0281   0.02858   0.01791  -0.0637   0.4940   1.0000
   6.750   1.0477   0.02917   0.01865  -0.0624   0.4819   1.0000
   7.000   1.0695   0.02965   0.01927  -0.0612   0.4708   1.0000
   7.250   1.0926   0.03005   0.01981  -0.0602   0.4599   1.0000
   7.500   1.1102   0.03052   0.02044  -0.0585   0.4455   1.0000
   7.750   1.1277   0.03068   0.02068  -0.0564   0.4282   1.0000
   8.000   1.1387   0.03103   0.02116  -0.0537   0.4072   1.0000
   8.250   1.1522   0.03114   0.02133  -0.0511   0.3855   1.0000
   8.500   1.1628   0.03160   0.02184  -0.0484   0.3641   1.0000
   8.750   1.1718   0.03217   0.02245  -0.0456   0.3418   1.0000
   9.000   1.1775   0.03291   0.02319  -0.0426   0.3171   1.0000
   9.250   1.1799   0.03384   0.02410  -0.0393   0.2925   1.0000
   9.500   1.1808   0.03509   0.02547  -0.0362   0.2686   1.0000
   9.750   1.1798   0.03657   0.02699  -0.0333   0.2413   1.0000
  10.000   1.1758   0.03846   0.02880  -0.0307   0.2100   1.0000
  10.250   1.1679   0.04090   0.03103  -0.0283   0.1803   1.0000
  10.500   1.1587   0.04384   0.03378  -0.0266   0.1559   1.0000
  10.750   1.1503   0.04704   0.03687  -0.0253   0.1360   1.0000
  11.000   1.1439   0.05028   0.04005  -0.0245   0.1218   1.0000
  11.250   1.1397   0.05347   0.04324  -0.0238   0.1119   1.0000
  11.500   1.1354   0.05677   0.04648  -0.0234   0.1049   1.0000
  11.750   1.1344   0.05982   0.04959  -0.0230   0.0984   1.0000
  12.000   1.1321   0.06303   0.05273  -0.0228   0.0936   1.0000
  12.250   1.1341   0.06588   0.05574  -0.0224   0.0884   1.0000
  12.500   1.1357   0.06874   0.05864  -0.0221   0.0844   1.0000
  12.750   1.1399   0.07132   0.06122  -0.0216   0.0811   1.0000
  13.000   1.1452   0.07398   0.06411  -0.0212   0.0773   1.0000
  13.250   1.1498   0.07668   0.06689  -0.0209   0.0737   1.0000
  13.500   1.1607   0.07852   0.06865  -0.0198   0.0704   1.0000
  13.750   1.1653   0.08168   0.07212  -0.0197   0.0681   1.0000
  14.000   1.1670   0.08521   0.07595  -0.0199   0.0656   1.0000
  14.250   1.1681   0.08878   0.07972  -0.0204   0.0633   1.0000
  14.500   1.1730   0.09181   0.08285  -0.0204   0.0610   1.0000
  14.750   1.1834   0.09446   0.08554  -0.0198   0.0589   1.0000
  15.000   1.1705   0.10031   0.09172  -0.0220   0.0583   1.0000
  15.250   1.1550   0.10684   0.09855  -0.0250   0.0578   1.0000
  15.500   1.1365   0.11428   0.10626  -0.0288   0.0575   1.0000
  15.750   1.1153   0.12277   0.11499  -0.0336   0.0575   1.0000
  16.000   1.0906   0.13272   0.12516  -0.0395   0.0576   1.0000
  16.250   1.0641   0.14417   0.13676  -0.0465   0.0581   1.0000
<< Back to GOE 596 AIRFOIL (goe596-il)

Polar data table (+)

Polar graphs


<< Back to GOE 596 AIRFOIL (goe596-il)