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GOE 11K AIRFOIL (goe11k-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 11K AIRFOIL (goe11k-il)
Reynolds number: 100,000
Max Cl/Cd: 41.91 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe11k-il-100000.txt
Download as CSV file: xf-goe11k-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4018   0.10693   0.10218  -0.0302   1.0000   0.0857
  -9.250  -0.4183   0.10483   0.10016  -0.0314   1.0000   0.0862
  -9.000  -0.4340   0.10250   0.09790  -0.0322   1.0000   0.0864
  -8.750  -0.3988   0.09622   0.09156  -0.0268   1.0000   0.0918
  -8.500  -0.4026   0.09329   0.08867  -0.0260   1.0000   0.0947
  -8.250  -0.4118   0.09048   0.08592  -0.0257   1.0000   0.0972
  -8.000  -0.4299   0.08809   0.08360  -0.0259   1.0000   0.0991
  -7.750  -0.5125   0.09339   0.08863  -0.0235   1.0000   0.0919
  -7.500  -0.5236   0.09116   0.08647  -0.0218   1.0000   0.0935
  -7.250  -0.5324   0.08841   0.08377  -0.0216   1.0000   0.0960
  -7.000  -0.5462   0.08527   0.08064  -0.0248   1.0000   0.0993
  -6.750  -0.5626   0.08267   0.07777  -0.0305   1.0000   0.1007
  -6.500  -0.5515   0.07757   0.07294  -0.0246   1.0000   0.1030
  -6.250  -0.5467   0.07491   0.07030  -0.0220   1.0000   0.1065
  -6.000  -0.5507   0.07235   0.06723  -0.0284   1.0000   0.1150
  -5.750  -0.5427   0.06723   0.06240  -0.0247   1.0000   0.1171
  -5.500  -0.5351   0.06458   0.05978  -0.0222   1.0000   0.1216
  -5.250  -0.5280   0.06066   0.05560  -0.0241   1.0000   0.1314
  -5.000  -0.5180   0.05830   0.05296  -0.0244   1.0000   0.1448
  -4.750  -0.5087   0.05515   0.05002  -0.0213   1.0000   0.1503
  -4.500  -0.4972   0.05212   0.04683  -0.0209   1.0000   0.1633
  -4.250  -0.4844   0.04933   0.04389  -0.0202   1.0000   0.1774
  -4.000  -0.4710   0.04676   0.04122  -0.0191   1.0000   0.1931
  -3.750  -0.4576   0.04469   0.03897  -0.0181   1.0000   0.2198
  -3.500  -0.4110   0.03342   0.02563  -0.0200   1.0000   0.1029
  -3.250  -0.3894   0.03077   0.02241  -0.0183   1.0000   0.0987
  -3.000  -0.3694   0.02869   0.01991  -0.0168   1.0000   0.1028
  -2.750  -0.3488   0.02715   0.01819  -0.0154   1.0000   0.1062
  -2.500  -0.3272   0.02585   0.01653  -0.0140   1.0000   0.1103
  -2.250  -0.3053   0.02478   0.01506  -0.0126   1.0000   0.1162
  -2.000  -0.2836   0.02372   0.01394  -0.0115   1.0000   0.1207
  -1.750  -0.2619   0.02320   0.01320  -0.0103   1.0000   0.1281
  -1.500  -0.2395   0.02233   0.01227  -0.0093   1.0000   0.1328
  -1.250  -0.2175   0.02180   0.01170  -0.0082   1.0000   0.1388
  -1.000  -0.1956   0.02133   0.01119  -0.0071   1.0000   0.1468
  -0.750  -0.1741   0.02091   0.01082  -0.0061   1.0000   0.1537
  -0.500  -0.1528   0.02054   0.01048  -0.0049   1.0000   0.1615
  -0.250  -0.1314   0.02028   0.01025  -0.0038   1.0000   0.1736
   0.000  -0.0290   0.01804   0.01069  -0.0191   1.0000   1.0000
   0.250  -0.0082   0.01833   0.01073  -0.0180   1.0000   1.0000
   0.500   0.0126   0.01863   0.01086  -0.0171   1.0000   1.0000
   0.750   0.0333   0.01895   0.01105  -0.0161   1.0000   1.0000
   1.000   0.0540   0.01928   0.01129  -0.0153   1.0000   1.0000
   1.250   0.0746   0.01964   0.01157  -0.0144   1.0000   1.0000
   1.500   0.0952   0.02002   0.01187  -0.0136   1.0000   1.0000
   1.750   0.1157   0.02041   0.01222  -0.0128   1.0000   1.0000
   2.000   0.1360   0.02084   0.01262  -0.0120   1.0000   1.0000
   2.250   0.1563   0.02128   0.01305  -0.0113   1.0000   1.0000
   2.500   0.1764   0.02175   0.01352  -0.0106   1.0000   1.0000
   2.750   0.2187   0.02266   0.01447  -0.0145   0.9919   1.0000
   3.000   0.2643   0.02355   0.01543  -0.0189   0.9809   1.0000
   3.250   0.3076   0.02427   0.01625  -0.0227   0.9695   1.0000
   3.500   0.3565   0.02477   0.01687  -0.0274   0.9533   1.0000
   3.750   0.4805   0.02331   0.01575  -0.0432   0.9039   1.0000
   4.000   0.5641   0.02116   0.01394  -0.0508   0.8701   1.0000
   4.250   0.6260   0.01912   0.01223  -0.0543   0.8414   1.0000
   4.500   0.7208   0.01720   0.00765  -0.0595   0.2361   1.0000
   4.750   0.7334   0.01876   0.00861  -0.0566   0.1757   1.0000
   5.000   0.7505   0.01991   0.00957  -0.0545   0.1548   1.0000
   5.250   0.7712   0.02100   0.01057  -0.0530   0.1375   1.0000
   5.500   0.7966   0.02231   0.01180  -0.0526   0.1233   1.0000
   5.750   0.8260   0.02382   0.01328  -0.0529   0.1095   1.0000
   6.000   0.8584   0.02566   0.01509  -0.0537   0.0974   1.0000
   6.250   0.8963   0.02838   0.01774  -0.0558   0.0874   1.0000
   6.500   0.9187   0.02972   0.01949  -0.0541   0.0804   1.0000
   6.750   0.9470   0.03232   0.02217  -0.0543   0.0749   1.0000
   7.000   0.9690   0.03548   0.02578  -0.0528   0.0727   1.0000
   7.250   0.9834   0.03749   0.02835  -0.0496   0.0697   1.0000
   7.500   0.9968   0.04024   0.03156  -0.0465   0.0679   1.0000
   7.750   1.0070   0.04380   0.03560  -0.0430   0.0688   1.0000
   8.000   1.0147   0.04780   0.04000  -0.0395   0.0707   1.0000
   8.250   1.0229   0.05269   0.04510  -0.0369   0.0729   1.0000
   9.500   0.9932   0.08547   0.08005  -0.0183   0.1430   1.0000
   9.750   0.9505   0.08688   0.08170  -0.0127   0.1424   1.0000
  10.000   0.9160   0.08966   0.08457  -0.0089   0.1421   1.0000
  10.250   0.8828   0.09374   0.08871  -0.0080   0.1419   1.0000
  10.500   0.8532   0.09923   0.09422  -0.0102   0.1417   1.0000
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