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GOE 599 AIRFOIL (goe599-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 599 AIRFOIL (goe599-il)
Reynolds number: 100,000
Max Cl/Cd: 48.51 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe599-il-100000.txt
Download as CSV file: xf-goe599-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 599 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5330   0.09315   0.08813  -0.0244   1.0000   0.1428
  -8.750  -0.5368   0.08953   0.08456  -0.0258   1.0000   0.1484
  -8.500  -0.5948   0.08374   0.07894  -0.0363   1.0000   0.1518
  -8.250  -0.5549   0.08149   0.07668  -0.0283   1.0000   0.1565
  -8.000  -0.6100   0.07640   0.07157  -0.0372   1.0000   0.1660
  -7.750  -0.6448   0.05400   0.04817  -0.0437   1.0000   0.0805
  -7.500  -0.6262   0.05451   0.04909  -0.0421   1.0000   0.0994
  -7.250  -0.6428   0.04317   0.03602  -0.0399   1.0000   0.0694
  -7.000  -0.6300   0.03888   0.03149  -0.0386   1.0000   0.0682
  -6.750  -0.6159   0.03531   0.02753  -0.0370   1.0000   0.0676
  -6.500  -0.5993   0.03229   0.02403  -0.0354   1.0000   0.0678
  -6.250  -0.5811   0.02977   0.02103  -0.0339   1.0000   0.0701
  -6.000  -0.5610   0.02779   0.01897  -0.0329   1.0000   0.0735
  -5.750  -0.5391   0.02597   0.01689  -0.0317   1.0000   0.0762
  -5.500  -0.5165   0.02460   0.01517  -0.0304   1.0000   0.0809
  -5.250  -0.4944   0.02305   0.01366  -0.0296   1.0000   0.0870
  -5.000  -0.4710   0.02193   0.01235  -0.0285   1.0000   0.0937
  -4.750  -0.4486   0.02067   0.01116  -0.0277   1.0000   0.1043
  -4.500  -0.4258   0.01939   0.00995  -0.0268   1.0000   0.1171
  -4.250  -0.4036   0.01805   0.00882  -0.0260   1.0000   0.1435
  -4.000  -0.3806   0.01636   0.00791  -0.0256   1.0000   0.2514
  -3.750  -0.3582   0.01545   0.00762  -0.0251   1.0000   0.3771
  -3.500  -0.3359   0.01501   0.00750  -0.0243   1.0000   0.4659
  -3.250  -0.3144   0.01476   0.00755  -0.0230   1.0000   0.5518
  -3.000  -0.2938   0.01456   0.00762  -0.0214   1.0000   0.6298
  -2.750  -0.2743   0.01441   0.00766  -0.0193   1.0000   0.6978
  -2.500  -0.2555   0.01431   0.00772  -0.0169   1.0000   0.7580
  -2.250  -0.2384   0.01427   0.00781  -0.0140   1.0000   0.8205
  -2.000  -0.2195   0.01430   0.00794  -0.0113   1.0000   0.8927
  -1.750  -0.1692   0.01445   0.00797  -0.0158   1.0000   0.9812
  -1.500  -0.1450   0.01445   0.00777  -0.0168   1.0000   1.0000
  -1.250  -0.1258   0.01458   0.00771  -0.0165   1.0000   1.0000
  -1.000  -0.1052   0.01478   0.00776  -0.0162   1.0000   1.0000
  -0.750  -0.0840   0.01503   0.00786  -0.0160   1.0000   1.0000
  -0.500  -0.0625   0.01532   0.00804  -0.0158   1.0000   1.0000
  -0.250  -0.0410   0.01566   0.00827  -0.0157   1.0000   1.0000
   0.000   0.0024   0.01612   0.00863  -0.0197   0.9941   1.0000
   0.250   0.0537   0.01655   0.00898  -0.0251   0.9842   1.0000
   0.500   0.1008   0.01687   0.00925  -0.0296   0.9734   1.0000
   0.750   0.1452   0.01714   0.00950  -0.0335   0.9626   1.0000
   1.000   0.1906   0.01738   0.00974  -0.0374   0.9528   1.0000
   1.250   0.2368   0.01755   0.00993  -0.0414   0.9431   1.0000
   1.500   0.2760   0.01766   0.01007  -0.0439   0.9311   1.0000
   1.750   0.3163   0.01772   0.01020  -0.0465   0.9194   1.0000
   2.000   0.3602   0.01765   0.01022  -0.0496   0.9075   1.0000
   2.250   0.4124   0.01716   0.00984  -0.0535   0.8925   1.0000
   2.500   0.4637   0.01633   0.00915  -0.0565   0.8754   1.0000
   2.750   0.5102   0.01549   0.00843  -0.0584   0.8577   1.0000
   3.000   0.5450   0.01492   0.00794  -0.0581   0.8347   1.0000
   3.250   0.5786   0.01437   0.00747  -0.0575   0.8073   1.0000
   3.500   0.6063   0.01411   0.00724  -0.0561   0.7767   1.0000
   3.750   0.6314   0.01392   0.00704  -0.0542   0.7387   1.0000
   4.000   0.6540   0.01387   0.00700  -0.0521   0.6945   1.0000
   4.250   0.6748   0.01391   0.00692  -0.0495   0.6309   1.0000
   4.500   0.6897   0.01445   0.00682  -0.0457   0.4972   1.0000
   4.750   0.6889   0.01690   0.00744  -0.0405   0.2034   1.0000
   5.000   0.7028   0.01859   0.00859  -0.0383   0.1563   1.0000
   5.250   0.7216   0.01969   0.00963  -0.0366   0.1423   1.0000
   5.500   0.7422   0.02071   0.01065  -0.0352   0.1334   1.0000
   5.750   0.7633   0.02195   0.01174  -0.0341   0.1251   1.0000
   6.000   0.7880   0.02294   0.01281  -0.0333   0.1187   1.0000
   6.250   0.8141   0.02439   0.01414  -0.0330   0.1129   1.0000
   6.500   0.8401   0.02564   0.01554  -0.0324   0.1063   1.0000
   6.750   0.8665   0.02727   0.01710  -0.0323   0.0994   1.0000
   7.000   0.8918   0.02888   0.01899  -0.0315   0.0929   1.0000
   7.250   0.9175   0.03133   0.02136  -0.0316   0.0855   1.0000
   7.500   0.9391   0.03324   0.02381  -0.0299   0.0811   1.0000
   7.750   0.9610   0.03527   0.02598  -0.0290   0.0751   1.0000
   8.000   0.9803   0.03955   0.03048  -0.0281   0.0721   1.0000
   8.250   0.9946   0.04237   0.03391  -0.0256   0.0709   1.0000
   8.500   1.0049   0.04544   0.03759  -0.0228   0.0690   1.0000
   8.750   1.0125   0.04915   0.04180  -0.0202   0.0682   1.0000
   9.000   1.0183   0.05375   0.04681  -0.0178   0.0697   1.0000
   9.250   1.0065   0.06041   0.05432  -0.0134   0.0805   1.0000
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