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GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 598 AIRFOIL (goe598-il)
Reynolds number: 50,000
Max Cl/Cd: 31.67 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe598-il-50000.txt
Download as CSV file: xf-goe598-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 598 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  -0.4792   0.10106   0.09464   0.0065   1.0000   0.2541
  -8.500  -0.4932   0.09853   0.09221   0.0054   1.0000   0.2641
  -8.250  -0.6013   0.10423   0.09759   0.0159   1.0000   0.2525
  -8.000  -0.5990   0.10073   0.09415   0.0157   1.0000   0.2661
  -7.750  -0.5907   0.09687   0.09031   0.0164   1.0000   0.2820
  -7.250  -0.5964   0.09113   0.08473   0.0160   1.0000   0.3181
  -7.000  -0.5768   0.08654   0.08012   0.0186   1.0000   0.3400
  -6.750  -0.5716   0.08323   0.07685   0.0201   1.0000   0.3655
  -6.500  -0.5709   0.08070   0.07439   0.0220   1.0000   0.4002
  -6.250  -0.5559   0.07728   0.07098   0.0260   1.0000   0.4405
  -6.000  -0.5425   0.07362   0.06733   0.0287   1.0000   0.4739
  -5.750  -0.5321   0.05147   0.04391  -0.0282   1.0000   0.1557
  -5.500  -0.5085   0.04571   0.03758  -0.0301   1.0000   0.1354
  -5.250  -0.4840   0.04101   0.03184  -0.0313   1.0000   0.1263
  -5.000  -0.4604   0.03710   0.02755  -0.0311   1.0000   0.1251
  -4.750  -0.4352   0.03351   0.02348  -0.0307   1.0000   0.1230
  -4.500  -0.4085   0.03043   0.01983  -0.0301   1.0000   0.1236
  -4.250  -0.3832   0.02805   0.01713  -0.0295   1.0000   0.1342
  -4.000  -0.3560   0.02575   0.01443  -0.0287   1.0000   0.1426
  -3.750  -0.3286   0.02375   0.01220  -0.0277   1.0000   0.1542
  -3.500  -0.3028   0.02203   0.01056  -0.0267   1.0000   0.1814
  -3.250  -0.2765   0.01991   0.00872  -0.0257   1.0000   0.2356
  -3.000  -0.2678   0.01599   0.00753  -0.0208   1.0000   0.6262
  -2.750  -0.1762   0.01528   0.00658  -0.0272   1.0000   1.0000
  -2.500  -0.1597   0.01503   0.00597  -0.0260   1.0000   1.0000
  -2.250  -0.1421   0.01485   0.00548  -0.0248   1.0000   1.0000
  -2.000  -0.1236   0.01474   0.00508  -0.0236   1.0000   1.0000
  -1.750  -0.1042   0.01466   0.00476  -0.0225   1.0000   1.0000
  -1.500  -0.0842   0.01462   0.00445  -0.0214   1.0000   1.0000
  -1.250  -0.0637   0.01461   0.00424  -0.0204   1.0000   1.0000
  -1.000  -0.0429   0.01463   0.00410  -0.0195   1.0000   1.0000
  -0.750  -0.0218   0.01467   0.00401  -0.0187   1.0000   1.0000
  -0.500  -0.0007   0.01475   0.00397  -0.0178   1.0000   1.0000
  -0.250   0.0204   0.01485   0.00396  -0.0171   1.0000   1.0000
   0.000   0.0415   0.01499   0.00402  -0.0164   1.0000   1.0000
   0.250   0.0627   0.01516   0.00414  -0.0157   1.0000   1.0000
   0.500   0.0839   0.01537   0.00432  -0.0151   1.0000   1.0000
   0.750   0.1050   0.01562   0.00456  -0.0146   1.0000   1.0000
   1.000   0.1261   0.01591   0.00486  -0.0142   1.0000   1.0000
   1.250   0.1471   0.01625   0.00521  -0.0138   1.0000   1.0000
   1.500   0.1680   0.01662   0.00562  -0.0136   1.0000   1.0000
   1.750   0.1889   0.01703   0.00609  -0.0134   1.0000   1.0000
   2.000   0.2097   0.01749   0.00663  -0.0133   1.0000   1.0000
   2.250   0.2304   0.01800   0.00727  -0.0132   1.0000   1.0000
   2.500   0.2510   0.01855   0.00795  -0.0133   1.0000   1.0000
   2.750   0.2714   0.01916   0.00870  -0.0134   1.0000   1.0000
   3.000   0.2916   0.01982   0.00953  -0.0136   1.0000   1.0000
   3.250   0.3115   0.02056   0.01045  -0.0138   1.0000   1.0000
   3.500   0.3310   0.02137   0.01155  -0.0142   1.0000   1.0000
   3.750   0.4162   0.02253   0.01341  -0.0265   0.9657   1.0000
   4.000   0.5659   0.01787   0.01027  -0.0310   0.7366   1.0000
   4.250   0.5537   0.02141   0.01026  -0.0188   0.1971   1.0000
   4.500   0.5757   0.02422   0.01258  -0.0172   0.1410   1.0000
   4.750   0.6026   0.02650   0.01464  -0.0160   0.1186   1.0000
   5.000   0.6314   0.02873   0.01689  -0.0152   0.1054   1.0000
   5.250   0.6614   0.03124   0.01983  -0.0141   0.1015   1.0000
   5.500   0.6896   0.03419   0.02324  -0.0131   0.1003   1.0000
   5.750   0.7154   0.03737   0.02702  -0.0121   0.0993   1.0000
   6.000   0.7384   0.04063   0.03078  -0.0111   0.0971   1.0000
   6.250   0.7591   0.04452   0.03527  -0.0101   0.0986   1.0000
   6.500   0.7785   0.04914   0.04028  -0.0094   0.1035   1.0000
   6.750   0.7941   0.05439   0.04636  -0.0088   0.1152   1.0000
   7.000   0.8056   0.06101   0.05379  -0.0097   0.1378   1.0000
   7.250   0.7421   0.05594   0.04963  -0.0061   0.1518   1.0000
   8.000   0.7330   0.09977   0.09336  -0.0561   0.3724   1.0000
   8.250   0.7542   0.10444   0.09805  -0.0538   0.3501   1.0000
   8.500   0.7425   0.10709   0.10062  -0.0538   0.3297   1.0000
   8.750   0.7531   0.11120   0.10478  -0.0525   0.3119   1.0000
   9.000   0.6191   0.10750   0.10130  -0.0376   0.3230   1.0000
   9.250   0.6161   0.11119   0.10495  -0.0372   0.3088   1.0000
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