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

GOE 199 (L.F.G. 5406) AIRFOIL (goe199-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 199 (L.F.G. 5406) AIRFOIL (goe199-il)
Reynolds number: 200,000
Max Cl/Cd: 71.96 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe199-il-200000.txt
Download as CSV file: xf-goe199-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 199 (L.F.G. 5406) AIRFOIL                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3203   0.08968   0.08643  -0.0232   1.0000   0.0273
  -7.000  -0.3165   0.08709   0.08390  -0.0239   1.0000   0.0279
  -6.750  -0.3183   0.08507   0.08194  -0.0235   1.0000   0.0284
  -6.500  -0.3284   0.08385   0.08079  -0.0211   1.0000   0.0286
  -6.250  -0.3376   0.08257   0.07956  -0.0192   0.9999   0.0290
  -6.000  -0.3006   0.07766   0.07457  -0.0284   0.9952   0.0304
  -5.750  -0.2296   0.07300   0.06960  -0.0490   0.9885   0.0318
  -5.500  -0.2121   0.06714   0.06380  -0.0505   0.9851   0.0323
  -5.250  -0.1891   0.06314   0.05980  -0.0527   0.9814   0.0333
  -5.000  -0.1568   0.05945   0.05605  -0.0576   0.9761   0.0353
  -4.750  -0.0949   0.05586   0.05211  -0.0691   0.9725   0.0390
  -4.500  -0.0555   0.05054   0.04665  -0.0754   0.9702   0.0400
  -4.250  -0.0314   0.04716   0.04328  -0.0773   0.9640   0.0412
  -4.000   0.0034   0.04416   0.04019  -0.0811   0.9595   0.0440
  -3.750   0.0599   0.04136   0.03686  -0.0876   0.9560   0.0495
  -3.500   0.0825   0.03794   0.03351  -0.0890   0.9484   0.0509
  -3.250   0.1135   0.03558   0.03108  -0.0911   0.9427   0.0538
  -3.000   0.1573   0.03472   0.02961  -0.0932   0.9354   0.0608
  -2.750   0.1795   0.03117   0.02621  -0.0942   0.9278   0.0631
  -2.500   0.2059   0.02962   0.02457  -0.0946   0.9192   0.0683
  -2.250   0.2375   0.02793   0.02254  -0.0949   0.9111   0.0752
  -2.000   0.2605   0.02635   0.02093  -0.0943   0.8992   0.0793
  -1.750   0.2870   0.02511   0.01936  -0.0933   0.8856   0.0889
  -1.500   0.3094   0.02384   0.01801  -0.0921   0.8706   0.0960
  -1.250   0.3341   0.02255   0.01655  -0.0912   0.8568   0.1076
  -1.000   0.3592   0.02159   0.01545  -0.0905   0.8443   0.1253
  -0.750   0.3840   0.02054   0.01427  -0.0901   0.8328   0.1516
  -0.500   0.4091   0.01954   0.01323  -0.0898   0.8197   0.1826
  -0.250   0.4344   0.01861   0.01221  -0.0895   0.8067   0.2237
   0.000   0.4600   0.01745   0.01105  -0.0891   0.7939   0.2576
   0.250   0.4880   0.01666   0.01013  -0.0885   0.7810   0.2785
   0.500   0.5311   0.01602   0.00837  -0.0859   0.7700   0.0840
   0.750   0.5589   0.01518   0.00738  -0.0852   0.7561   0.0781
   1.000   0.5869   0.01517   0.00715  -0.0844   0.7403   0.0732
   1.250   0.6140   0.01427   0.00625  -0.0839   0.7235   0.0704
   1.500   0.6411   0.01379   0.00574  -0.0834   0.7050   0.0685
   1.750   0.6681   0.01343   0.00534  -0.0829   0.6854   0.0676
   2.000   0.6950   0.01316   0.00502  -0.0824   0.6647   0.0678
   2.250   0.7219   0.01297   0.00479  -0.0820   0.6401   0.0692
   2.500   0.7486   0.01291   0.00462  -0.0816   0.6132   0.0727
   2.750   0.7750   0.01287   0.00443  -0.0811   0.5840   0.0793
   3.000   0.8012   0.01300   0.00438  -0.0806   0.5536   0.0872
   3.250   0.8275   0.01150   0.00441  -0.0804   0.5232   1.0000
   3.500   0.8531   0.01188   0.00451  -0.0799   0.4903   1.0000
   3.750   0.8783   0.01231   0.00468  -0.0795   0.4533   1.0000
   4.000   0.9031   0.01277   0.00491  -0.0790   0.4119   1.0000
   4.250   0.9275   0.01332   0.00518  -0.0785   0.3762   1.0000
   4.500   0.9520   0.01387   0.00552  -0.0781   0.3497   1.0000
   4.750   0.9766   0.01444   0.00590  -0.0778   0.3314   1.0000
   5.000   1.0015   0.01497   0.00633  -0.0774   0.3178   1.0000
   5.250   1.0271   0.01543   0.00674  -0.0771   0.3069   1.0000
   5.500   1.0523   0.01595   0.00720  -0.0768   0.2987   1.0000
   5.750   1.0779   0.01641   0.00764  -0.0766   0.2916   1.0000
   6.000   1.1032   0.01698   0.00816  -0.0763   0.2858   1.0000
   6.250   1.1289   0.01738   0.00862  -0.0761   0.2795   1.0000
   6.500   1.1539   0.01790   0.00907  -0.0758   0.2731   1.0000
   6.750   1.1792   0.01826   0.00955  -0.0755   0.2663   1.0000
   7.000   1.2038   0.01868   0.00996  -0.0752   0.2594   1.0000
   7.250   1.2284   0.01907   0.01044  -0.0749   0.2527   1.0000
   7.500   1.2527   0.01942   0.01084  -0.0745   0.2459   1.0000
   7.750   1.2767   0.01985   0.01136  -0.0741   0.2393   1.0000
   8.000   1.2999   0.02011   0.01170  -0.0735   0.2303   1.0000
   8.250   1.3227   0.02036   0.01208  -0.0730   0.2199   1.0000
   8.500   1.3449   0.02057   0.01241  -0.0723   0.2070   1.0000
   8.750   1.3661   0.02077   0.01267  -0.0715   0.1836   1.0000
   9.000   1.3732   0.02284   0.01401  -0.0691   0.0681   1.0000
   9.250   1.3838   0.02466   0.01578  -0.0667   0.0497   1.0000
   9.500   1.3969   0.02598   0.01724  -0.0647   0.0442   1.0000
   9.750   1.4049   0.02752   0.01889  -0.0621   0.0406   1.0000
  10.000   1.4066   0.02927   0.02077  -0.0586   0.0383   1.0000
  10.250   1.4097   0.03067   0.02233  -0.0553   0.0365   1.0000
  10.500   1.4097   0.03236   0.02414  -0.0519   0.0350   1.0000
  10.750   1.4089   0.03432   0.02619  -0.0490   0.0338   1.0000
  11.000   1.4075   0.03653   0.02849  -0.0464   0.0328   1.0000
  11.250   1.4053   0.03901   0.03105  -0.0440   0.0319   1.0000
  11.500   1.4028   0.04183   0.03392  -0.0417   0.0310   1.0000
  11.750   1.4055   0.04449   0.03663  -0.0396   0.0301   1.0000
  12.000   1.4114   0.04650   0.03881  -0.0383   0.0293   1.0000
  12.250   1.4168   0.04870   0.04117  -0.0370   0.0283   1.0000
  12.500   1.4230   0.05101   0.04361  -0.0356   0.0276   1.0000
  12.750   1.4296   0.05342   0.04616  -0.0342   0.0270   1.0000
  13.000   1.4361   0.05594   0.04883  -0.0329   0.0265   1.0000
  13.250   1.4420   0.05863   0.05167  -0.0317   0.0261   1.0000
  13.500   1.4464   0.06154   0.05476  -0.0305   0.0258   1.0000
  13.750   1.4490   0.06465   0.05802  -0.0296   0.0254   1.0000
  14.000   1.4480   0.06800   0.06156  -0.0289   0.0250   1.0000
  14.250   1.4463   0.07154   0.06523  -0.0284   0.0245   1.0000
  14.500   1.4445   0.07550   0.06933  -0.0279   0.0240   1.0000
  14.750   1.4383   0.08018   0.07420  -0.0277   0.0237   1.0000
  15.000   1.4268   0.08509   0.07933  -0.0280   0.0236   1.0000
  15.250   1.4133   0.08997   0.08443  -0.0291   0.0237   1.0000
  15.500   1.3981   0.09514   0.08983  -0.0308   0.0238   1.0000
  15.750   1.3813   0.10078   0.09570  -0.0332   0.0239   1.0000
  16.000   1.3632   0.10712   0.10228  -0.0364   0.0241   1.0000
  16.250   1.3423   0.11455   0.10998  -0.0408   0.0245   1.0000
  16.500   1.3087   0.12531   0.12108  -0.0483   0.0251   1.0000
  16.750   1.2742   0.13769   0.13378  -0.0571   0.0258   1.0000
  17.000   1.2425   0.15063   0.14694  -0.0663   0.0267   1.0000
  17.250   1.2101   0.16494   0.16143  -0.0762   0.0280   1.0000
  17.500   1.1849   0.17795   0.17453  -0.0846   0.0293   1.0000
  17.750   1.1816   0.18378   0.18036  -0.0865   0.0304   1.0000
  18.000   0.8468   0.18285   0.17967  -0.0655   0.0385   1.0000
<< Back to GOE 199 (L.F.G. 5406) AIRFOIL (goe199-il)

Polar data table (+)

Polar graphs


<< Back to GOE 199 (L.F.G. 5406) AIRFOIL (goe199-il)