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

GOE 264 AIRFOIL (goe264-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 264 AIRFOIL (goe264-il)
Reynolds number: 200,000
Max Cl/Cd: 79.69 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe264-il-200000-n5.txt
Download as CSV file: xf-goe264-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 264 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3031   0.09415   0.09086  -0.0249   1.0000   0.0175
  -7.500  -0.3075   0.09216   0.08894  -0.0236   1.0000   0.0174
  -7.250  -0.3113   0.09006   0.08689  -0.0228   1.0000   0.0174
  -7.000  -0.2951   0.08589   0.08273  -0.0278   0.9954   0.0174
  -6.750  -0.2666   0.07941   0.07623  -0.0393   0.9881   0.0181
  -6.500  -0.2441   0.07780   0.07462  -0.0406   0.9846   0.0199
  -6.250  -0.2161   0.07364   0.07041  -0.0474   0.9781   0.0208
  -6.000  -0.1863   0.06884   0.06556  -0.0549   0.9707   0.0211
  -5.750  -0.1522   0.06391   0.06055  -0.0632   0.9643   0.0218
  -5.500  -0.1137   0.05794   0.05446  -0.0735   0.9546   0.0244
  -5.250  -0.0782   0.05292   0.04932  -0.0810   0.9460   0.0258
  -5.000  -0.0503   0.05072   0.04706  -0.0836   0.9376   0.0271
  -4.750  -0.0182   0.04762   0.04383  -0.0880   0.9270   0.0302
  -4.500   0.0327   0.04023   0.03607  -0.0984   0.9183   0.0332
  -4.250   0.0533   0.03994   0.03576  -0.0978   0.9071   0.0349
  -4.000   0.0829   0.03793   0.03362  -0.0998   0.8947   0.0382
  -3.750   0.1196   0.03416   0.02956  -0.1037   0.8826   0.0394
  -3.500   0.1624   0.02914   0.02406  -0.1083   0.8708   0.0426
  -3.250   0.1932   0.02640   0.02102  -0.1100   0.8567   0.0433
  -3.000   0.2225   0.02458   0.01897  -0.1109   0.8414   0.0441
  -2.750   0.2524   0.02280   0.01690  -0.1117   0.8249   0.0449
  -2.500   0.2818   0.02132   0.01514  -0.1122   0.8076   0.0461
  -2.250   0.3115   0.01998   0.01350  -0.1127   0.7886   0.0481
  -2.000   0.3432   0.01779   0.01082  -0.1136   0.7683   0.0483
  -1.750   0.3732   0.01631   0.00890  -0.1139   0.7462   0.0488
  -1.500   0.4020   0.01534   0.00755  -0.1139   0.7210   0.0498
  -1.250   0.4301   0.01468   0.00654  -0.1138   0.6952   0.0511
  -1.000   0.4579   0.01422   0.00580  -0.1136   0.6720   0.0524
  -0.750   0.4855   0.01392   0.00525  -0.1134   0.6516   0.0536
  -0.500   0.5132   0.01363   0.00475  -0.1133   0.6338   0.0547
  -0.250   0.5409   0.01328   0.00431  -0.1134   0.6180   0.0578
   0.000   0.5685   0.01319   0.00414  -0.1133   0.6032   0.0610
   0.250   0.5964   0.01305   0.00391  -0.1133   0.5892   0.0635
   0.500   0.6244   0.01294   0.00372  -0.1133   0.5756   0.0659
   0.750   0.6523   0.01287   0.00357  -0.1133   0.5621   0.0683
   1.000   0.6803   0.01279   0.00349  -0.1134   0.5485   0.0730
   1.250   0.7081   0.01279   0.00346  -0.1134   0.5353   0.0792
   1.500   0.7358   0.01281   0.00350  -0.1134   0.5221   0.0921
   1.750   0.7635   0.01284   0.00357  -0.1134   0.5087   0.1143
   2.000   0.7909   0.01292   0.00364  -0.1134   0.4955   0.1410
   2.250   0.8182   0.01300   0.00373  -0.1133   0.4829   0.1646
   2.500   0.8456   0.01308   0.00386  -0.1133   0.4711   0.1973
   2.750   0.8731   0.01304   0.00403  -0.1135   0.4598   0.2993
   3.250   0.9188   0.01221   0.00423  -0.1113   0.4393   1.0000
   3.500   0.9459   0.01244   0.00440  -0.1112   0.4286   1.0000
   3.750   0.9728   0.01269   0.00459  -0.1110   0.4188   1.0000
   4.000   0.9995   0.01296   0.00482  -0.1109   0.4093   1.0000
   4.250   1.0265   0.01320   0.00506  -0.1108   0.4005   1.0000
   4.500   1.0530   0.01349   0.00531  -0.1107   0.3929   1.0000
   4.750   1.0799   0.01373   0.00559  -0.1105   0.3854   1.0000
   5.000   1.1062   0.01404   0.00590  -0.1104   0.3782   1.0000
   5.250   1.1328   0.01429   0.00622  -0.1102   0.3704   1.0000
   5.500   1.1588   0.01462   0.00655  -0.1100   0.3641   1.0000
   5.750   1.1852   0.01488   0.00692  -0.1098   0.3561   1.0000
   6.000   1.2105   0.01519   0.00724  -0.1095   0.3411   1.0000
   6.250   1.2349   0.01555   0.00756  -0.1091   0.3214   1.0000
   6.500   1.2591   0.01593   0.00792  -0.1086   0.2963   1.0000
   6.750   1.2827   0.01640   0.00834  -0.1081   0.2685   1.0000
   7.000   1.3031   0.01725   0.00895  -0.1072   0.2096   1.0000
   7.250   1.3073   0.02029   0.01095  -0.1047   0.0540   1.0000
   7.500   1.3238   0.02162   0.01221  -0.1032   0.0302   1.0000
   7.750   1.3425   0.02259   0.01326  -0.1019   0.0232   1.0000
   8.000   1.3591   0.02371   0.01449  -0.1004   0.0196   1.0000
   8.250   1.3736   0.02496   0.01591  -0.0985   0.0178   1.0000
   8.500   1.3887   0.02603   0.01718  -0.0968   0.0163   1.0000
   8.750   1.4016   0.02720   0.01848  -0.0948   0.0149   1.0000
   9.000   1.4108   0.02855   0.01998  -0.0924   0.0140   1.0000
   9.250   1.4159   0.03002   0.02157  -0.0894   0.0133   1.0000
   9.500   1.4162   0.03168   0.02335  -0.0858   0.0128   1.0000
   9.750   1.4148   0.03367   0.02547  -0.0826   0.0124   1.0000
  10.000   1.4110   0.03607   0.02799  -0.0796   0.0120   1.0000
  10.250   1.4043   0.03901   0.03105  -0.0767   0.0115   1.0000
  10.500   1.4066   0.04125   0.03341  -0.0749   0.0111   1.0000
  10.750   1.4111   0.04336   0.03566  -0.0734   0.0107   1.0000
  11.000   1.4145   0.04567   0.03811  -0.0721   0.0102   1.0000
  11.250   1.4169   0.04820   0.04077  -0.0707   0.0099   1.0000
  11.500   1.4192   0.05083   0.04358  -0.0694   0.0096   1.0000
  11.750   1.4217   0.05351   0.04640  -0.0682   0.0094   1.0000
  12.000   1.4240   0.05628   0.04931  -0.0670   0.0091   1.0000
  12.250   1.4258   0.05917   0.05236  -0.0661   0.0089   1.0000
  12.500   1.4269   0.06221   0.05556  -0.0652   0.0088   1.0000
  12.750   1.4269   0.06543   0.05895  -0.0646   0.0086   1.0000
  13.000   1.4255   0.06889   0.06257  -0.0641   0.0085   1.0000
  13.250   1.4229   0.07259   0.06646  -0.0639   0.0083   1.0000
  13.500   1.4187   0.07651   0.07055  -0.0640   0.0082   1.0000
  13.750   1.4134   0.08063   0.07483  -0.0644   0.0081   1.0000
  14.000   1.4070   0.08500   0.07936  -0.0651   0.0080   1.0000
  14.250   1.3992   0.08970   0.08422  -0.0660   0.0078   1.0000
  14.500   1.3893   0.09487   0.08956  -0.0671   0.0077   1.0000
  14.750   1.3767   0.10072   0.09559  -0.0688   0.0076   1.0000
  15.000   1.3648   0.10672   0.10183  -0.0715   0.0075   1.0000
  15.250   1.3523   0.11306   0.10839  -0.0745   0.0075   1.0000
  15.500   1.3393   0.11979   0.11534  -0.0780   0.0075   1.0000
  15.750   1.3261   0.12685   0.12260  -0.0818   0.0074   1.0000
  16.000   1.3128   0.13427   0.13022  -0.0861   0.0074   1.0000
<< Back to GOE 264 AIRFOIL (goe264-il)

Polar data table (+)

Polar graphs


<< Back to GOE 264 AIRFOIL (goe264-il)