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

GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 100,000
Max Cl/Cd: 54.32 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe113-il-100000-n5.txt
Download as CSV file: xf-goe113-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4292   0.09697   0.09224  -0.0110   1.0000   0.0259
  -7.500  -0.4258   0.09373   0.08905  -0.0124   1.0000   0.0262
  -7.250  -0.4227   0.09050   0.08588  -0.0139   1.0000   0.0265
  -7.000  -0.4168   0.08697   0.08240  -0.0165   1.0000   0.0268
  -6.750  -0.4093   0.08334   0.07880  -0.0194   1.0000   0.0272
  -6.500  -0.4004   0.07966   0.07515  -0.0224   1.0000   0.0281
  -6.250  -0.3898   0.07591   0.07141  -0.0256   1.0000   0.0294
  -6.000  -0.3754   0.07202   0.06749  -0.0301   1.0000   0.0315
  -5.750  -0.3525   0.06808   0.06339  -0.0368   1.0000   0.0329
  -5.500  -0.3333   0.06431   0.05946  -0.0399   1.0000   0.0331
  -5.250  -0.3152   0.06053   0.05550  -0.0417   1.0000   0.0333
  -5.000  -0.2972   0.05672   0.05148  -0.0428   1.0000   0.0333
  -4.500  -0.2604   0.04909   0.04351  -0.0439   1.0000   0.0333
  -4.000  -0.2201   0.03967   0.03373  -0.0441   1.0000   0.0229
  -3.500  -0.1805   0.03361   0.02735  -0.0446   1.0000   0.0264
  -3.250  -0.1565   0.03108   0.02450  -0.0445   1.0000   0.0288
  -3.000  -0.1307   0.02827   0.02126  -0.0442   1.0000   0.0289
  -2.750  -0.0964   0.02545   0.01785  -0.0453   0.9973   0.0298
  -2.500  -0.0551   0.02301   0.01479  -0.0477   0.9912   0.0352
  -2.250  -0.0152   0.02097   0.01236  -0.0500   0.9845   0.0382
  -2.000   0.0242   0.01940   0.01039  -0.0521   0.9769   0.0453
  -1.750   0.0625   0.01797   0.00878  -0.0542   0.9684   0.0486
  -1.500   0.1002   0.01720   0.00783  -0.0561   0.9579   0.0554
  -1.250   0.1362   0.01626   0.00682  -0.0577   0.9462   0.0554
  -1.000   0.1710   0.01554   0.00601  -0.0590   0.9322   0.0555
  -0.750   0.2058   0.01499   0.00535  -0.0603   0.9170   0.0560
  -0.500   0.2410   0.01455   0.00479  -0.0616   0.9008   0.0572
  -0.250   0.2743   0.01419   0.00429  -0.0625   0.8819   0.0604
   0.000   0.3078   0.01392   0.00392  -0.0633   0.8629   0.0678
   0.250   0.3398   0.01122   0.00374  -0.0637   0.8457   1.0000
   0.500   0.3703   0.01127   0.00350  -0.0638   0.8228   1.0000
   0.750   0.3997   0.01135   0.00334  -0.0637   0.7995   1.0000
   1.000   0.4278   0.01147   0.00325  -0.0634   0.7756   1.0000
   1.250   0.4550   0.01161   0.00323  -0.0630   0.7525   1.0000
   1.500   0.4821   0.01178   0.00325  -0.0626   0.7311   1.0000
   1.750   0.5086   0.01196   0.00330  -0.0622   0.7104   1.0000
   2.000   0.5349   0.01215   0.00338  -0.0617   0.6894   1.0000
   2.250   0.5608   0.01235   0.00350  -0.0611   0.6670   1.0000
   2.500   0.5866   0.01256   0.00361  -0.0605   0.6450   1.0000
   2.750   0.6123   0.01277   0.00375  -0.0600   0.6228   1.0000
   3.000   0.6379   0.01300   0.00390  -0.0594   0.6000   1.0000
   3.250   0.6635   0.01322   0.00414  -0.0588   0.5766   1.0000
   3.500   0.6892   0.01346   0.00436  -0.0583   0.5560   1.0000
   3.750   0.7150   0.01369   0.00464  -0.0579   0.5348   1.0000
   4.000   0.7404   0.01396   0.00491  -0.0574   0.5116   1.0000
   4.250   0.7657   0.01423   0.00527  -0.0568   0.4842   1.0000
   4.500   0.7903   0.01455   0.00559  -0.0562   0.4455   1.0000
   4.750   0.8101   0.01529   0.00582  -0.0548   0.3391   1.0000
   5.000   0.8309   0.01623   0.00635  -0.0539   0.2681   1.0000
   5.250   0.8528   0.01711   0.00708  -0.0532   0.2233   1.0000
   5.500   0.8741   0.01810   0.00775  -0.0525   0.1537   1.0000
   5.750   0.8953   0.01919   0.00855  -0.0516   0.0680   1.0000
   6.000   0.9117   0.02113   0.01005  -0.0501   0.0244   1.0000
   6.250   0.9327   0.02240   0.01148  -0.0489   0.0201   1.0000
   6.500   0.9543   0.02352   0.01290  -0.0479   0.0190   1.0000
   6.750   0.9746   0.02480   0.01455  -0.0466   0.0182   1.0000
   7.000   0.9933   0.02622   0.01623  -0.0452   0.0169   1.0000
   7.250   1.0102   0.02778   0.01801  -0.0438   0.0153   1.0000
   7.500   1.0258   0.02951   0.01993  -0.0421   0.0145   1.0000
   7.750   1.0415   0.03142   0.02198  -0.0404   0.0141   1.0000
   8.000   1.0585   0.03351   0.02422  -0.0388   0.0139   1.0000
   8.250   1.0769   0.03584   0.02672  -0.0374   0.0137   1.0000
   8.500   1.0955   0.03838   0.02950  -0.0360   0.0137   1.0000
   8.750   1.1129   0.04115   0.03258  -0.0346   0.0137   1.0000
   9.000   1.1269   0.04413   0.03590  -0.0330   0.0136   1.0000
   9.250   1.1366   0.04731   0.03953  -0.0313   0.0133   1.0000
   9.500   1.1415   0.05085   0.04340  -0.0295   0.0128   1.0000
   9.750   1.1418   0.05461   0.04750  -0.0275   0.0125   1.0000
  10.000   1.1376   0.05847   0.05169  -0.0254   0.0123   1.0000
  10.250   1.1288   0.06199   0.05551  -0.0231   0.0122   1.0000
  10.500   1.1168   0.06526   0.05904  -0.0209   0.0122   1.0000
  10.750   1.1036   0.06878   0.06281  -0.0198   0.0122   1.0000
  11.000   1.0894   0.07277   0.06704  -0.0200   0.0122   1.0000
  11.250   1.0747   0.07734   0.07183  -0.0215   0.0123   1.0000
  11.500   1.0601   0.08249   0.07719  -0.0241   0.0124   1.0000
  11.750   1.0447   0.08845   0.08335  -0.0279   0.0125   1.0000
  12.000   1.0291   0.09525   0.09035  -0.0326   0.0127   1.0000
  12.250   1.0122   0.10325   0.09853  -0.0385   0.0130   1.0000
  12.500   0.9843   0.11671   0.11221  -0.0487   0.0144   1.0000
<< Back to GOE 113 AIRFOIL (goe113-il)

Polar data table (+)

Polar graphs


<< Back to GOE 113 AIRFOIL (goe113-il)