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GOE 11K AIRFOIL (goe11k-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 11K AIRFOIL (goe11k-il)
Reynolds number: 200,000
Max Cl/Cd: 72.42 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe11k-il-200000.txt
Download as CSV file: xf-goe11k-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 11K 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5211   0.10237   0.09895  -0.0316   1.0000   0.0458
  -8.500  -0.5360   0.09999   0.09663  -0.0307   1.0000   0.0458
  -8.250  -0.5463   0.09693   0.09359  -0.0309   1.0000   0.0459
  -8.000  -0.5556   0.09087   0.08759  -0.0283   1.0000   0.0466
  -7.750  -0.5562   0.08794   0.08468  -0.0254   1.0000   0.0473
  -7.500  -0.5589   0.08513   0.08189  -0.0239   1.0000   0.0477
  -7.250  -0.5614   0.08233   0.07909  -0.0227   1.0000   0.0483
  -7.000  -0.5634   0.07941   0.07617  -0.0218   1.0000   0.0492
  -6.750  -0.5648   0.07619   0.07294  -0.0216   1.0000   0.0500
  -6.500  -0.5648   0.07273   0.06944  -0.0218   1.0000   0.0511
  -6.250  -0.5627   0.06899   0.06565  -0.0223   1.0000   0.0527
  -6.000  -0.5579   0.06489   0.06145  -0.0233   1.0000   0.0545
  -5.750  -0.5441   0.06042   0.05643  -0.0265   1.0000   0.0582
  -5.500  -0.5352   0.05688   0.05259  -0.0257   1.0000   0.0585
  -5.250  -0.5338   0.05023   0.04608  -0.0252   1.0000   0.0605
  -5.000  -0.5224   0.04840   0.04426  -0.0238   1.0000   0.0626
  -4.750  -0.5041   0.04551   0.04119  -0.0239   0.9991   0.0654
  -4.500  -0.4746   0.04047   0.03535  -0.0261   0.9957   0.0737
  -4.250  -0.4500   0.03754   0.03238  -0.0271   0.9931   0.0758
  -4.000  -0.4250   0.03557   0.03025  -0.0274   0.9901   0.0811
  -3.750  -0.3994   0.03442   0.02864  -0.0279   0.9871   0.1022
  -3.500  -0.3736   0.03192   0.02612  -0.0284   0.9854   0.1058
  -3.000  -0.3191   0.02227   0.01448  -0.0240   0.9814   0.0677
  -2.750  -0.2913   0.02153   0.01341  -0.0238   0.9783   0.0735
  -2.500  -0.2628   0.02065   0.01251  -0.0242   0.9760   0.0788
  -2.250  -0.2317   0.02045   0.01217  -0.0250   0.9741   0.0850
  -2.000  -0.2053   0.01978   0.01143  -0.0248   0.9717   0.0891
  -1.750  -0.1818   0.01940   0.01107  -0.0242   0.9681   0.0942
  -1.500  -0.1538   0.01918   0.01080  -0.0244   0.9652   0.0986
  -1.250  -0.1243   0.01888   0.01050  -0.0248   0.9629   0.1026
  -1.000  -0.0937   0.01881   0.01050  -0.0256   0.9611   0.1096
  -0.750  -0.0736   0.01863   0.01028  -0.0242   0.9572   0.1142
  -0.500  -0.0489   0.01842   0.01016  -0.0238   0.9536   0.1215
  -0.250  -0.0186   0.01836   0.01016  -0.0245   0.9506   0.1349
   0.000   0.0153   0.01775   0.01032  -0.0260   0.9480   0.3041
   0.250   0.1339   0.01625   0.01073  -0.0463   0.9527   1.0000
   0.500   0.1711   0.01647   0.01086  -0.0484   0.9494   1.0000
   0.750   0.1981   0.01661   0.01094  -0.0485   0.9442   1.0000
   1.000   0.2294   0.01672   0.01101  -0.0494   0.9392   1.0000
   1.250   0.2683   0.01688   0.01115  -0.0518   0.9361   1.0000
   1.500   0.2927   0.01698   0.01124  -0.0513   0.9297   1.0000
   1.750   0.3371   0.01687   0.01113  -0.0545   0.9238   1.0000
   2.000   0.3845   0.01634   0.01065  -0.0578   0.9127   1.0000
   2.250   0.4386   0.01569   0.01005  -0.0624   0.9044   1.0000
   2.500   0.5111   0.01425   0.00872  -0.0699   0.8944   1.0000
   2.750   0.5798   0.01244   0.00704  -0.0763   0.8805   1.0000
   3.000   0.6356   0.01129   0.00601  -0.0807   0.8642   1.0000
   3.250   0.6715   0.01065   0.00546  -0.0813   0.8375   1.0000
   3.500   0.7394   0.01021   0.00414  -0.0875   0.5862   1.0000
   3.750   0.7194   0.01186   0.00464  -0.0773   0.3939   1.0000
   4.000   0.7090   0.01367   0.00527  -0.0699   0.1797   1.0000
   4.250   0.7200   0.01463   0.00587  -0.0666   0.1312   1.0000
   4.500   0.7353   0.01528   0.00644  -0.0641   0.1159   1.0000
   4.750   0.7527   0.01582   0.00697  -0.0619   0.1054   1.0000
   5.000   0.7696   0.01650   0.00762  -0.0597   0.0966   1.0000
   5.250   0.7864   0.01739   0.00842  -0.0577   0.0878   1.0000
   5.500   0.8066   0.01796   0.00908  -0.0561   0.0795   1.0000
   5.750   0.8281   0.01914   0.01022  -0.0550   0.0716   1.0000
   6.000   0.8498   0.01992   0.01101  -0.0538   0.0638   1.0000
   6.250   0.8774   0.02138   0.01254  -0.0539   0.0568   1.0000
   6.500   0.9020   0.02241   0.01362  -0.0532   0.0514   1.0000
   6.750   0.9313   0.02430   0.01562  -0.0537   0.0464   1.0000
   7.000   0.9556   0.02549   0.01702  -0.0528   0.0430   1.0000
   7.250   0.9795   0.02693   0.01862  -0.0520   0.0404   1.0000
   7.500   1.0028   0.02900   0.02073  -0.0516   0.0378   1.0000
   7.750   1.0204   0.03106   0.02321  -0.0495   0.0358   1.0000
   8.000   1.0366   0.03364   0.02622  -0.0470   0.0348   1.0000
   8.250   1.0481   0.03679   0.02982  -0.0438   0.0343   1.0000
   8.500   1.0537   0.04039   0.03387  -0.0398   0.0343   1.0000
   8.750   1.0552   0.04420   0.03808  -0.0355   0.0349   1.0000
   9.000   1.0522   0.04806   0.04231  -0.0308   0.0352   1.0000
   9.250   1.0474   0.05165   0.04622  -0.0261   0.0353   1.0000
   9.500   1.0390   0.05548   0.05033  -0.0215   0.0357   1.0000
   9.750   1.0284   0.05942   0.05448  -0.0171   0.0361   1.0000
  10.000   0.9514   0.07144   0.06740  -0.0050   0.0554   1.0000
  10.250   0.9289   0.07505   0.07113  -0.0016   0.0555   1.0000
  10.500   0.9056   0.07925   0.07542   0.0001   0.0559   1.0000
  10.750   0.8828   0.08427   0.08052   0.0003   0.0570   1.0000
  11.000   0.8576   0.09044   0.08678  -0.0021   0.0572   1.0000
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