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=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 199 (L.F.G. 5406) AIRFOIL (goe199-il)
Reynolds number: 50,000
Max Cl/Cd: 40.36 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe199-il-50000-n5.txt
Download as CSV file: xf-goe199-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3233   0.10638   0.09977  -0.0234   1.0000   0.0700
  -7.750  -0.3253   0.10543   0.09893  -0.0268   1.0000   0.0716
  -7.500  -0.3238   0.10470   0.09828  -0.0321   1.0000   0.0722
  -7.250  -0.3149   0.09831   0.09197  -0.0268   1.0000   0.0745
  -7.000  -0.3104   0.09532   0.08905  -0.0263   1.0000   0.0769
  -6.750  -0.3079   0.09293   0.08674  -0.0267   1.0000   0.0794
  -6.500  -0.3057   0.09104   0.08489  -0.0283   1.0000   0.0823
  -6.250  -0.2998   0.09063   0.08446  -0.0339   1.0000   0.0844
  -6.000  -0.2975   0.08714   0.08106  -0.0328   1.0000   0.0856
  -5.750  -0.2977   0.08357   0.07758  -0.0289   1.0000   0.0880
  -5.500  -0.2941   0.08115   0.07519  -0.0281   1.0000   0.0921
  -5.250  -0.2721   0.08077   0.07462  -0.0368   1.0000   0.0989
  -5.000  -0.2761   0.07626   0.07028  -0.0315   1.0000   0.1013
  -4.750  -0.2697   0.07361   0.06765  -0.0305   1.0000   0.1064
  -4.500  -0.2367   0.07068   0.06453  -0.0381   0.9966   0.1152
  -4.250  -0.2100   0.06645   0.06027  -0.0405   0.9900   0.1233
  -4.000  -0.1723   0.06259   0.05626  -0.0467   0.9830   0.1351
  -3.750  -0.1360   0.05920   0.05274  -0.0520   0.9761   0.1541
  -3.500  -0.0993   0.05601   0.04938  -0.0574   0.9686   0.1771
  -3.250  -0.0620   0.05296   0.04616  -0.0622   0.9626   0.2065
  -3.000  -0.0358   0.04991   0.04310  -0.0639   0.9545   0.2388
  -2.250   0.1196   0.04143   0.03322  -0.0815   0.9347   0.1403
  -2.000   0.1679   0.03872   0.02986  -0.0847   0.9273   0.0947
  -1.750   0.2153   0.03704   0.02746  -0.0878   0.9206   0.0846
  -1.500   0.2533   0.03522   0.02538  -0.0901   0.9129   0.0842
  -1.250   0.2914   0.03342   0.02336  -0.0923   0.9052   0.0822
  -1.000   0.3275   0.03202   0.02169  -0.0938   0.8963   0.0797
  -0.750   0.3675   0.03070   0.02005  -0.0957   0.8884   0.0776
  -0.500   0.4010   0.02973   0.01879  -0.0964   0.8775   0.0763
  -0.250   0.4368   0.02877   0.01759  -0.0974   0.8674   0.0761
   0.000   0.4741   0.02785   0.01644  -0.0984   0.8568   0.0789
   0.250   0.5076   0.02712   0.01546  -0.0984   0.8428   0.0809
   0.500   0.5400   0.02636   0.01451  -0.0982   0.8278   0.0814
   0.750   0.5708   0.02566   0.01365  -0.0976   0.8120   0.0823
   1.000   0.6000   0.02505   0.01290  -0.0967   0.7955   0.0836
   1.250   0.6264   0.02459   0.01234  -0.0955   0.7772   0.0853
   1.500   0.6525   0.02416   0.01187  -0.0945   0.7583   0.0892
   1.750   0.6787   0.02383   0.01147  -0.0934   0.7393   0.0968
   2.000   0.7053   0.02351   0.01114  -0.0925   0.7203   0.1065
   2.250   0.7318   0.02322   0.01086  -0.0915   0.7009   0.1213
   2.500   0.7557   0.02147   0.01075  -0.0904   0.6787   1.0000
   2.750   0.7813   0.02153   0.01051  -0.0891   0.6567   1.0000
   3.000   0.8058   0.02168   0.01044  -0.0880   0.6315   1.0000
   3.250   0.8307   0.02181   0.01035  -0.0868   0.6065   1.0000
   3.750   0.8798   0.02222   0.01034  -0.0845   0.5531   1.0000
   4.000   0.9039   0.02257   0.01048  -0.0835   0.5261   1.0000
   4.250   0.9276   0.02302   0.01072  -0.0824   0.4997   1.0000
   4.500   0.9508   0.02356   0.01112  -0.0815   0.4736   1.0000
   4.750   0.9739   0.02415   0.01159  -0.0807   0.4490   1.0000
   5.000   0.9971   0.02477   0.01209  -0.0799   0.4278   1.0000
   5.500   1.0441   0.02610   0.01327  -0.0785   0.3941   1.0000
   5.750   1.0678   0.02680   0.01394  -0.0779   0.3804   1.0000
   6.000   1.0918   0.02754   0.01469  -0.0775   0.3681   1.0000
   6.250   1.1160   0.02829   0.01549  -0.0770   0.3577   1.0000
   6.500   1.1406   0.02904   0.01622  -0.0766   0.3494   1.0000
   6.750   1.1650   0.02987   0.01720  -0.0762   0.3410   1.0000
   7.000   1.1904   0.03068   0.01808  -0.0760   0.3350   1.0000
   7.250   1.2150   0.03160   0.01918  -0.0757   0.3293   1.0000
   7.500   1.2397   0.03254   0.02031  -0.0754   0.3241   1.0000
   7.750   1.2652   0.03347   0.02137  -0.0752   0.3200   1.0000
   8.000   1.2894   0.03454   0.02271  -0.0750   0.3160   1.0000
   8.250   1.3119   0.03571   0.02419  -0.0745   0.3116   1.0000
   8.500   1.3352   0.03684   0.02559  -0.0741   0.3077   1.0000
   8.750   1.3600   0.03782   0.02672  -0.0738   0.3035   1.0000
   9.000   1.3755   0.03901   0.02835  -0.0725   0.2950   1.0000
   9.250   1.3913   0.03960   0.02911  -0.0709   0.2826   1.0000
   9.500   1.4035   0.04009   0.02976  -0.0688   0.2680   1.0000
   9.750   1.4115   0.04064   0.03051  -0.0663   0.2524   1.0000
  10.000   1.4144   0.04132   0.03139  -0.0632   0.2356   1.0000
  10.250   1.4137   0.04224   0.03252  -0.0599   0.2188   1.0000
  10.500   1.4106   0.04346   0.03402  -0.0564   0.2028   1.0000
  10.750   1.4023   0.04502   0.03580  -0.0527   0.1855   1.0000
  11.000   1.3933   0.04726   0.03833  -0.0499   0.1587   1.0000
  11.250   1.3804   0.05033   0.04127  -0.0478   0.1023   1.0000
  11.500   1.3609   0.05476   0.04529  -0.0466   0.0889   1.0000
  11.750   1.3424   0.05979   0.05019  -0.0462   0.0819   1.0000
  12.000   1.3265   0.06497   0.05549  -0.0463   0.0760   1.0000
  12.250   1.3093   0.07064   0.06128  -0.0471   0.0725   1.0000
  12.500   1.2912   0.07675   0.06749  -0.0485   0.0700   1.0000
  12.750   1.2766   0.08263   0.07352  -0.0500   0.0672   1.0000
  13.000   1.2633   0.08848   0.07956  -0.0516   0.0646   1.0000
  13.250   1.2505   0.09428   0.08549  -0.0532   0.0623   1.0000
  13.500   1.2388   0.09991   0.09121  -0.0548   0.0602   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)