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

GOE 512 AIRFOIL (goe512-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 512 AIRFOIL (goe512-il)
Reynolds number: 50,000
Max Cl/Cd: 19.48 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe512-il-50000.txt
Download as CSV file: xf-goe512-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 512 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3378   0.12266   0.11669  -0.0216   1.0000   0.2195
  -8.500  -0.3021   0.11587   0.10987  -0.0193   1.0000   0.2275
  -8.250  -0.3216   0.11529   0.10942  -0.0178   1.0000   0.2333
  -8.000  -0.3596   0.11626   0.11057  -0.0159   1.0000   0.2350
  -7.750  -0.3306   0.11029   0.10460  -0.0141   1.0000   0.2404
  -7.500  -0.3400   0.10857   0.10297  -0.0119   1.0000   0.2464
  -7.250  -0.3737   0.10872   0.10328  -0.0094   1.0000   0.2500
  -7.000  -0.4183   0.10951   0.10425  -0.0059   1.0000   0.2511
  -6.750  -0.3866   0.10363   0.09837  -0.0046   1.0000   0.2572
  -6.500  -0.4054   0.10236   0.09720  -0.0015   1.0000   0.2620
  -6.250  -0.4398   0.10208   0.09706   0.0010   1.0000   0.2659
  -6.000  -0.4689   0.10105   0.09611   0.0022   1.0000   0.2682
  -5.750  -0.4532   0.09694   0.09205   0.0056   1.0000   0.2751
  -5.500  -0.4770   0.09585   0.09101   0.0067   1.0000   0.2818
  -5.250  -0.4846   0.09298   0.08820   0.0085   1.0000   0.2861
  -5.000  -0.4875   0.09058   0.08584   0.0107   1.0000   0.2941
  -4.750  -0.5002   0.08832   0.08359   0.0113   1.0000   0.3010
  -4.500  -0.5005   0.08569   0.08100   0.0135   1.0000   0.3091
  -4.250  -0.5070   0.08313   0.07844   0.0141   1.0000   0.3171
  -4.000  -0.4854   0.07974   0.07494   0.0099   0.9908   0.3316
  -3.750  -0.4548   0.07576   0.07088   0.0062   0.9766   0.3495
  -3.500  -0.3917   0.06553   0.05937  -0.0139   0.9637   0.2005
  -3.250  -0.3584   0.05859   0.05155  -0.0183   0.9527   0.1714
  -3.000  -0.3248   0.05521   0.04765  -0.0206   0.9396   0.1715
  -2.750  -0.2900   0.05199   0.04385  -0.0226   0.9269   0.1716
  -2.500  -0.2600   0.05027   0.04186  -0.0236   0.9127   0.1773
  -2.250  -0.2295   0.04871   0.03994  -0.0243   0.8987   0.1830
  -2.000  -0.1984   0.04734   0.03817  -0.0250   0.8846   0.1921
  -1.750  -0.1669   0.04640   0.03686  -0.0257   0.8709   0.2032
  -1.500  -0.1354   0.04605   0.03634  -0.0263   0.8574   0.2186
  -1.250  -0.1017   0.04574   0.03588  -0.0273   0.8444   0.2376
  -1.000  -0.0588   0.04525   0.03520  -0.0294   0.8329   0.2699
  -0.750  -0.0349   0.04462   0.03450  -0.0290   0.8194   0.3091
  -0.500   0.0212   0.04295   0.03289  -0.0346   0.8073   0.4029
  -0.250   0.2537   0.03991   0.03195  -0.0720   0.7972   1.0000
   0.000   0.2983   0.04020   0.03180  -0.0748   0.7852   1.0000
   0.250   0.3198   0.04089   0.03222  -0.0741   0.7708   1.0000
   0.500   0.3416   0.04158   0.03266  -0.0735   0.7571   1.0000
   0.750   0.3690   0.04208   0.03291  -0.0734   0.7444   1.0000
   1.000   0.4324   0.04120   0.03170  -0.0776   0.7349   1.0000
   1.250   0.4479   0.04184   0.03219  -0.0756   0.7203   1.0000
   1.500   0.4646   0.04245   0.03264  -0.0738   0.7061   1.0000
   1.750   0.4859   0.04290   0.03295  -0.0725   0.6929   1.0000
   2.000   0.5699   0.04044   0.03023  -0.0783   0.6860   1.0000
   2.250   0.5741   0.04154   0.03125  -0.0748   0.6715   1.0000
   2.500   0.5797   0.04273   0.03235  -0.0717   0.6579   1.0000
   2.750   0.6947   0.03918   0.02859  -0.0820   0.6523   1.0000
   3.000   0.6731   0.04132   0.03069  -0.0749   0.6377   1.0000
   3.250   0.6629   0.04317   0.03250  -0.0697   0.6236   1.0000
   3.500   0.6800   0.04391   0.03316  -0.0679   0.6119   1.0000
   3.750   0.7703   0.04138   0.03052  -0.0750   0.6032   1.0000
   4.000   0.7308   0.04452   0.03364  -0.0660   0.5889   1.0000
   4.250   0.7195   0.04673   0.03583  -0.0611   0.5757   1.0000
   4.500   0.8361   0.04292   0.03189  -0.0711   0.5685   1.0000
   4.750   0.7568   0.04828   0.03728  -0.0580   0.5543   1.0000
   5.000   0.7026   0.05398   0.04299  -0.0508   0.5399   1.0000
   5.250   0.8070   0.04927   0.03819  -0.0564   0.5353   1.0000
   5.500   0.7191   0.05749   0.04643  -0.0471   0.5202   1.0000
   5.750   0.6581   0.06548   0.05444  -0.0431   0.5079   1.0000
   6.000   0.7064   0.06412   0.05304  -0.0430   0.5022   1.0000
   6.250   0.6353   0.07358   0.06252  -0.0402   0.4955   1.0000
   6.500   0.6182   0.07823   0.06717  -0.0391   0.4916   1.0000
   6.750   0.6548   0.07838   0.06729  -0.0387   0.4855   1.0000
   7.000   0.6346   0.08347   0.07237  -0.0378   0.4829   1.0000
   7.250   0.6226   0.08810   0.07701  -0.0374   0.4831   1.0000
   8.000   0.5020   0.11170   0.10088  -0.0422   0.6004   1.0000
   8.250   0.4914   0.11265   0.10183  -0.0401   0.5910   1.0000
   8.500   0.5198   0.11620   0.10537  -0.0413   0.5851   1.0000
   8.750   0.5086   0.11738   0.10654  -0.0394   0.5765   1.0000
   9.000   0.5352   0.12063   0.10980  -0.0402   0.5686   1.0000
<< Back to GOE 512 AIRFOIL (goe512-il)

Polar data table (+)

Polar graphs


<< Back to GOE 512 AIRFOIL (goe512-il)