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

GOE 373 AIRFOIL (goe373-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 373 AIRFOIL (goe373-il)
Reynolds number: 50,000
Max Cl/Cd: 41.88 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe373-il-50000-n5.txt
Download as CSV file: xf-goe373-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 373 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3037   0.10468   0.09799  -0.0241   1.0000   0.0728
  -7.750  -0.3105   0.10432   0.09779  -0.0253   1.0000   0.0738
  -7.500  -0.3125   0.10402   0.09760  -0.0293   1.0000   0.0743
  -7.250  -0.2984   0.09674   0.09034  -0.0236   1.0000   0.0781
  -7.000  -0.2949   0.09426   0.08794  -0.0235   1.0000   0.0822
  -6.750  -0.2939   0.09274   0.08652  -0.0254   1.0000   0.0859
  -6.500  -0.2912   0.09282   0.08668  -0.0318   1.0000   0.0877
  -6.250  -0.2876   0.08800   0.08195  -0.0283   1.0000   0.0896
  -6.000  -0.2836   0.08473   0.07875  -0.0258   1.0000   0.0928
  -5.750  -0.2798   0.08246   0.07654  -0.0259   1.0000   0.0969
  -5.500  -0.2703   0.08186   0.07593  -0.0322   1.0000   0.1019
  -5.250  -0.2670   0.07858   0.07273  -0.0314   1.0000   0.1037
  -5.000  -0.2670   0.07549   0.06974  -0.0277   1.0000   0.1071
  -4.750  -0.2587   0.07365   0.06789  -0.0293   1.0000   0.1151
  -4.500  -0.2501   0.07108   0.06533  -0.0310   1.0000   0.1198
  -4.250  -0.2097   0.06759   0.06168  -0.0395   0.9909   0.1340
  -4.000  -0.1837   0.06352   0.05759  -0.0415   0.9823   0.1436
  -3.750  -0.1473   0.05979   0.05376  -0.0473   0.9728   0.1562
  -3.500  -0.1101   0.05642   0.05026  -0.0530   0.9630   0.1728
  -3.250  -0.0757   0.05327   0.04701  -0.0579   0.9521   0.1975
  -3.000  -0.0458   0.05027   0.04395  -0.0607   0.9417   0.2288
  -2.500   0.0086   0.04480   0.03849  -0.0638   0.9210   0.3324
  -2.000   0.1530   0.03887   0.03108  -0.0853   0.9033   0.1608
  -1.750   0.2012   0.03613   0.02779  -0.0887   0.8921   0.1099
  -1.500   0.2437   0.03407   0.02530  -0.0912   0.8808   0.0960
  -1.250   0.2870   0.03227   0.02298  -0.0934   0.8693   0.0852
  -1.000   0.3265   0.03044   0.02084  -0.0952   0.8570   0.0805
  -0.750   0.3653   0.02911   0.01901  -0.0962   0.8431   0.0757
  -0.500   0.4002   0.02798   0.01755  -0.0967   0.8283   0.0735
  -0.250   0.4336   0.02681   0.01614  -0.0972   0.8134   0.0722
   0.000   0.4651   0.02593   0.01502  -0.0972   0.7983   0.0721
   0.250   0.4955   0.02526   0.01409  -0.0969   0.7831   0.0732
   0.500   0.5260   0.02468   0.01324  -0.0966   0.7678   0.0740
   0.750   0.5563   0.02411   0.01246  -0.0963   0.7526   0.0740
   1.000   0.5851   0.02363   0.01182  -0.0958   0.7375   0.0740
   1.250   0.6128   0.02325   0.01131  -0.0951   0.7221   0.0743
   1.500   0.6399   0.02294   0.01090  -0.0945   0.7069   0.0749
   1.750   0.6667   0.02271   0.01059  -0.0939   0.6918   0.0761
   2.000   0.6933   0.02258   0.01038  -0.0933   0.6768   0.0779
   2.250   0.7199   0.02253   0.01027  -0.0928   0.6619   0.0812
   2.500   0.7462   0.02257   0.01021  -0.0922   0.6472   0.0878
   2.750   0.7725   0.02258   0.01023  -0.0917   0.6328   0.0976
   3.000   0.7989   0.02260   0.01036  -0.0911   0.6188   0.1173
   3.250   0.8253   0.02134   0.01056  -0.0908   0.6053   1.0000
   3.500   0.8511   0.02169   0.01072  -0.0900   0.5921   1.0000
   3.750   0.8770   0.02205   0.01093  -0.0894   0.5794   1.0000
   4.000   0.9031   0.02242   0.01118  -0.0887   0.5674   1.0000
   4.250   0.9281   0.02286   0.01156  -0.0880   0.5544   1.0000
   4.500   0.9526   0.02332   0.01198  -0.0873   0.5409   1.0000
   4.750   0.9767   0.02376   0.01238  -0.0864   0.5266   1.0000
   5.000   1.0004   0.02418   0.01279  -0.0855   0.5112   1.0000
   5.250   1.0234   0.02460   0.01319  -0.0844   0.4953   1.0000
   5.500   1.0462   0.02505   0.01365  -0.0834   0.4800   1.0000
   5.750   1.0691   0.02558   0.01427  -0.0825   0.4665   1.0000
   6.000   1.0919   0.02612   0.01489  -0.0816   0.4536   1.0000
   6.250   1.1148   0.02665   0.01551  -0.0807   0.4410   1.0000
   6.500   1.1375   0.02716   0.01611  -0.0797   0.4283   1.0000
   6.750   1.1595   0.02770   0.01679  -0.0787   0.4151   1.0000
   7.000   1.1806   0.02831   0.01756  -0.0776   0.4013   1.0000
   7.250   1.2012   0.02889   0.01827  -0.0764   0.3867   1.0000
   7.500   1.2193   0.02933   0.01877  -0.0747   0.3673   1.0000
   7.750   1.2353   0.02987   0.01944  -0.0729   0.3458   1.0000
   8.000   1.2510   0.03040   0.01996  -0.0710   0.3245   1.0000
   8.250   1.2629   0.03113   0.02077  -0.0688   0.2992   1.0000
   8.500   1.2761   0.03192   0.02161  -0.0669   0.2783   1.0000
   8.750   1.2904   0.03284   0.02273  -0.0653   0.2610   1.0000
   9.000   1.3028   0.03383   0.02389  -0.0634   0.2419   1.0000
   9.250   1.3130   0.03490   0.02511  -0.0615   0.2230   1.0000
   9.500   1.3193   0.03617   0.02639  -0.0592   0.2008   1.0000
   9.750   1.3221   0.03767   0.02784  -0.0567   0.1817   1.0000
  10.000   1.3241   0.03935   0.02956  -0.0543   0.1668   1.0000
  10.250   1.3235   0.04136   0.03162  -0.0521   0.1508   1.0000
  10.500   1.3200   0.04383   0.03415  -0.0503   0.1267   1.0000
  10.750   1.3144   0.04675   0.03707  -0.0491   0.0968   1.0000
  11.000   1.3040   0.05042   0.04061  -0.0484   0.0737   1.0000
  11.250   1.2936   0.05446   0.04466  -0.0482   0.0617   1.0000
  11.500   1.2823   0.05892   0.04913  -0.0486   0.0540   1.0000
  11.750   1.2736   0.06335   0.05368  -0.0492   0.0478   1.0000
  12.000   1.2632   0.06822   0.05865  -0.0503   0.0442   1.0000
  12.250   1.2540   0.07312   0.06369  -0.0515   0.0417   1.0000
  12.500   1.2455   0.07807   0.06881  -0.0528   0.0398   1.0000
  12.750   1.2368   0.08318   0.07408  -0.0543   0.0384   1.0000
  13.000   1.2278   0.08842   0.07947  -0.0560   0.0372   1.0000
  13.250   1.2191   0.09372   0.08490  -0.0577   0.0362   1.0000
  13.500   1.2110   0.09903   0.09033  -0.0596   0.0354   1.0000
<< Back to GOE 373 AIRFOIL (goe373-il)

Polar data table (+)

Polar graphs


<< Back to GOE 373 AIRFOIL (goe373-il)