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GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 622 AIRFOIL (goe622-il)
Reynolds number: 500,000
Max Cl/Cd: 88.55 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe622-il-500000.txt
Download as CSV file: xf-goe622-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 622 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.3964   0.11231   0.11011  -0.0120   1.0000   0.0192
 -10.500  -0.3963   0.10837   0.10617  -0.0132   1.0000   0.0200
  -9.000  -0.5065   0.09458   0.09236  -0.0126   1.0000   0.0214
  -7.250  -0.5318   0.04026   0.03759  -0.0463   1.0000   0.0152
  -7.000  -0.5386   0.03097   0.02767  -0.0461   1.0000   0.0144
  -6.750  -0.5332   0.02526   0.02132  -0.0443   1.0000   0.0142
  -6.500  -0.5195   0.02261   0.01826  -0.0425   1.0000   0.0148
  -6.250  -0.5031   0.02122   0.01659  -0.0407   1.0000   0.0153
  -6.000  -0.4788   0.01936   0.01437  -0.0405   0.9991   0.0157
  -5.750  -0.4490   0.01561   0.01011  -0.0417   0.9973   0.0167
  -5.500  -0.4148   0.01453   0.00890  -0.0434   0.9953   0.0178
  -5.250  -0.3806   0.01370   0.00795  -0.0448   0.9931   0.0192
  -5.000  -0.3466   0.01306   0.00720  -0.0462   0.9901   0.0211
  -4.750  -0.3129   0.01192   0.00593  -0.0476   0.9874   0.0237
  -4.500  -0.2775   0.01131   0.00529  -0.0493   0.9850   0.0269
  -4.250  -0.2416   0.01093   0.00482  -0.0510   0.9827   0.0303
  -4.000  -0.2096   0.01019   0.00408  -0.0520   0.9780   0.0383
  -3.750  -0.1744   0.00966   0.00353  -0.0536   0.9746   0.0473
  -3.500  -0.1388   0.00926   0.00313  -0.0553   0.9717   0.0594
  -3.250  -0.1066   0.00894   0.00288  -0.0562   0.9662   0.0794
  -3.000  -0.0738   0.00871   0.00273  -0.0572   0.9603   0.1045
  -2.750  -0.0411   0.00850   0.00253  -0.0582   0.9545   0.1217
  -2.500  -0.0109   0.00829   0.00233  -0.0587   0.9466   0.1347
  -2.250   0.0190   0.00808   0.00215  -0.0590   0.9385   0.1498
  -2.000   0.0481   0.00784   0.00199  -0.0592   0.9299   0.1733
  -1.750   0.0751   0.00754   0.00185  -0.0590   0.9196   0.2244
  -1.500   0.1012   0.00703   0.00175  -0.0587   0.9096   0.3534
  -1.250   0.1247   0.00624   0.00170  -0.0581   0.8992   0.5790
  -1.000   0.1441   0.00553   0.00173  -0.0559   0.8872   0.7994
  -0.750   0.1655   0.00530   0.00175  -0.0535   0.8734   0.9209
  -0.500   0.2050   0.00529   0.00168  -0.0556   0.8571   0.9757
  -0.250   0.2477   0.00530   0.00158  -0.0586   0.8355   0.9979
   0.000   0.2754   0.00534   0.00149  -0.0584   0.8147   1.0000
   0.250   0.3009   0.00538   0.00144  -0.0578   0.7941   1.0000
   0.500   0.3265   0.00545   0.00139  -0.0571   0.7735   1.0000
   0.750   0.3522   0.00552   0.00136  -0.0565   0.7500   1.0000
   1.000   0.3777   0.00563   0.00135  -0.0559   0.7233   1.0000
   1.250   0.4032   0.00576   0.00134  -0.0553   0.6954   1.0000
   1.500   0.4288   0.00592   0.00136  -0.0548   0.6694   1.0000
   1.750   0.4544   0.00609   0.00141  -0.0542   0.6445   1.0000
   2.000   0.4805   0.00625   0.00148  -0.0538   0.6225   1.0000
   2.250   0.5065   0.00643   0.00155  -0.0534   0.6021   1.0000
   2.500   0.5329   0.00658   0.00164  -0.0531   0.5827   1.0000
   2.750   0.5589   0.00677   0.00173  -0.0527   0.5608   1.0000
   3.000   0.5852   0.00694   0.00185  -0.0523   0.5385   1.0000
   3.250   0.6116   0.00712   0.00196  -0.0520   0.5183   1.0000
   3.500   0.6380   0.00729   0.00208  -0.0517   0.4950   1.0000
   3.750   0.6641   0.00750   0.00222  -0.0514   0.4663   1.0000
   4.000   0.6893   0.00780   0.00236  -0.0509   0.4201   1.0000
   4.250   0.7133   0.00828   0.00256  -0.0503   0.3557   1.0000
   4.500   0.7370   0.00885   0.00285  -0.0498   0.2944   1.0000
   4.750   0.7608   0.00943   0.00318  -0.0493   0.2399   1.0000
   5.000   0.7855   0.00988   0.00347  -0.0489   0.2025   1.0000
   5.250   0.8099   0.01039   0.00379  -0.0484   0.1582   1.0000
   5.500   0.8340   0.01093   0.00416  -0.0480   0.1268   1.0000
   5.750   0.8589   0.01137   0.00452  -0.0476   0.1062   1.0000
   6.000   0.8814   0.01214   0.00500  -0.0469   0.0541   1.0000
   6.250   0.9033   0.01305   0.00570  -0.0460   0.0233   1.0000
   6.500   0.9276   0.01361   0.00634  -0.0453   0.0192   1.0000
   6.750   0.9513   0.01422   0.00704  -0.0446   0.0175   1.0000
   7.000   0.9740   0.01498   0.00788  -0.0438   0.0160   1.0000
   7.250   0.9943   0.01608   0.00912  -0.0425   0.0149   1.0000
   7.500   1.0129   0.01739   0.01055  -0.0410   0.0142   1.0000
   7.750   1.0348   0.01817   0.01142  -0.0401   0.0137   1.0000
   8.000   1.0562   0.01902   0.01236  -0.0392   0.0129   1.0000
   8.250   1.0761   0.02009   0.01352  -0.0380   0.0124   1.0000
   8.500   1.0953   0.02130   0.01484  -0.0367   0.0119   1.0000
   8.750   1.1140   0.02261   0.01627  -0.0354   0.0116   1.0000
   9.000   1.1323   0.02405   0.01783  -0.0341   0.0112   1.0000
   9.250   1.1498   0.02567   0.01957  -0.0328   0.0109   1.0000
   9.500   1.1663   0.02755   0.02161  -0.0313   0.0107   1.0000
   9.750   1.1816   0.02973   0.02398  -0.0297   0.0106   1.0000
  10.000   1.1946   0.03233   0.02689  -0.0278   0.0108   1.0000
  10.250   1.1903   0.03859   0.03392  -0.0237   0.0127   1.0000
  10.750   1.1487   0.05334   0.04983  -0.0147   0.0197   1.0000
  11.000   1.1391   0.05528   0.05197  -0.0117   0.0190   1.0000
  11.250   1.1237   0.05849   0.05538  -0.0099   0.0186   1.0000
  11.500   1.1059   0.06250   0.05957  -0.0095   0.0184   1.0000
  11.750   1.0849   0.06764   0.06489  -0.0109   0.0185   1.0000
  12.000   1.0623   0.07394   0.07136  -0.0141   0.0187   1.0000
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