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GOE 451 AIRFOIL (modified line 6) (goe451-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 451 AIRFOIL (modified line 6) (goe451-il)
Reynolds number: 50,000
Max Cl/Cd: 36.5 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe451-il-50000.txt
Download as CSV file: xf-goe451-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 451 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4351   0.10936   0.10235  -0.0218   1.0000   0.0809
  -8.000  -0.4388   0.10882   0.10194  -0.0231   1.0000   0.0822
  -7.750  -0.4382   0.10842   0.10164  -0.0267   1.0000   0.0828
  -7.500  -0.4274   0.10062   0.09391  -0.0230   1.0000   0.0863
  -7.250  -0.4220   0.09747   0.09084  -0.0233   1.0000   0.0904
  -7.000  -0.4176   0.09542   0.08888  -0.0258   1.0000   0.0943
  -6.750  -0.4093   0.09519   0.08871  -0.0326   1.0000   0.0962
  -6.500  -0.4044   0.08883   0.08246  -0.0281   1.0000   0.1003
  -6.250  -0.3959   0.08594   0.07962  -0.0296   1.0000   0.1066
  -6.000  -0.3815   0.08411   0.07783  -0.0364   1.0000   0.1106
  -5.750  -0.3769   0.07941   0.07322  -0.0327   1.0000   0.1173
  -5.500  -0.3596   0.07697   0.07064  -0.0386   1.0000   0.1245
  -5.250  -0.3473   0.07365   0.06735  -0.0396   1.0000   0.1354
  -5.000  -0.3381   0.06968   0.06347  -0.0389   1.0000   0.1427
  -4.750  -0.3220   0.06633   0.06014  -0.0414   1.0000   0.1548
  -4.500  -0.3061   0.06301   0.05683  -0.0431   1.0000   0.1692
  -4.250  -0.2824   0.06009   0.05385  -0.0478   1.0000   0.1914
  -4.000  -0.2725   0.05637   0.05022  -0.0462   1.0000   0.2123
  -3.750  -0.2558   0.05319   0.04706  -0.0474   1.0000   0.2474
  -2.000   0.0016   0.02986   0.02165  -0.0762   1.0000   0.1854
  -1.750   0.0394   0.02717   0.01818  -0.0780   1.0000   0.1726
  -1.500   0.0747   0.02476   0.01511  -0.0790   1.0000   0.1591
  -1.250   0.1069   0.02296   0.01257  -0.0792   1.0000   0.1498
  -1.000   0.1347   0.02150   0.01068  -0.0789   1.0000   0.1486
  -0.750   0.1616   0.02046   0.00924  -0.0783   1.0000   0.1508
  -0.500   0.1866   0.01957   0.00821  -0.0778   1.0000   0.1602
  -0.250   0.2107   0.01890   0.00754  -0.0775   1.0000   0.1909
   0.000   0.2376   0.01668   0.00708  -0.0780   1.0000   0.5733
   0.250   0.2527   0.01652   0.00666  -0.0757   1.0000   1.0000
   0.500   0.2726   0.01686   0.00673  -0.0750   1.0000   1.0000
   0.750   0.3016   0.01728   0.00695  -0.0764   0.9948   1.0000
   1.000   0.3751   0.01737   0.00692  -0.0865   0.9601   1.0000
   1.250   0.4669   0.01583   0.00518  -0.0970   0.8742   1.0000
   1.500   0.5022   0.01559   0.00481  -0.0982   0.8301   1.0000
   1.750   0.5371   0.01545   0.00436  -0.0989   0.7631   1.0000
   2.000   0.5638   0.01562   0.00421  -0.0983   0.6932   1.0000
   2.250   0.5855   0.01604   0.00418  -0.0970   0.5849   1.0000
   2.500   0.5886   0.01931   0.00483  -0.0934   0.0947   1.0000
   2.750   0.6091   0.02033   0.00607  -0.0928   0.0796   1.0000
   3.000   0.6277   0.02187   0.00782  -0.0922   0.0666   1.0000
   3.250   0.6425   0.02392   0.00999  -0.0912   0.0547   1.0000
   3.500   0.6586   0.02614   0.01219  -0.0902   0.0421   1.0000
   3.750   0.7001   0.03062   0.01580  -0.0929   0.0307   1.0000
   4.000   0.7213   0.03112   0.01666  -0.0922   0.0246   1.0000
   4.250   0.7484   0.03291   0.01847  -0.0925   0.0219   1.0000
   4.500   0.7716   0.03384   0.01967  -0.0919   0.0191   1.0000
   4.750   0.8000   0.03760   0.02336  -0.0928   0.0167   1.0000
   5.000   0.8256   0.04130   0.02736  -0.0930   0.0158   1.0000
   5.250   0.8448   0.04033   0.02701  -0.0912   0.0151   1.0000
   5.500   0.8665   0.04143   0.02849  -0.0903   0.0142   1.0000
   5.750   0.8883   0.04323   0.03065  -0.0896   0.0135   1.0000
   6.000   0.9091   0.04541   0.03323  -0.0889   0.0127   1.0000
   6.250   0.9286   0.04821   0.03648  -0.0882   0.0119   1.0000
   6.500   0.9465   0.05235   0.04105  -0.0878   0.0113   1.0000
   6.750   0.9646   0.06055   0.04944  -0.0888   0.0108   1.0000
   8.500   1.0582   0.07036   0.06198  -0.0802   0.0091   1.0000
   8.750   1.0665   0.07361   0.06558  -0.0789   0.0090   1.0000
   9.000   1.0730   0.07704   0.06933  -0.0776   0.0088   1.0000
   9.250   1.0775   0.08060   0.07319  -0.0763   0.0087   1.0000
   9.500   1.0799   0.08426   0.07713  -0.0750   0.0086   1.0000
   9.750   1.0801   0.08797   0.08111  -0.0738   0.0085   1.0000
  10.000   1.0780   0.09172   0.08509  -0.0727   0.0085   1.0000
  10.250   1.0735   0.09546   0.08904  -0.0716   0.0085   1.0000
  10.500   1.0665   0.09917   0.09293  -0.0707   0.0085   1.0000
  10.750   1.0561   0.10270   0.09660  -0.0696   0.0085   1.0000
  11.000   1.0440   0.10631   0.10034  -0.0690   0.0085   1.0000
  11.250   1.0319   0.11030   0.10443  -0.0695   0.0086   1.0000
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