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GOE 57 AIRFOIL (goe57-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 57 AIRFOIL (goe57-il)
Reynolds number: 50,000
Max Cl/Cd: 43.5 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe57-il-50000-n5.txt
Download as CSV file: xf-goe57-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 57 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3230   0.11238   0.10564  -0.0266   1.0000   0.0737
  -8.000  -0.3284   0.11169   0.10509  -0.0277   1.0000   0.0741
  -7.750  -0.3296   0.11066   0.10417  -0.0304   1.0000   0.0743
  -7.500  -0.3164   0.10306   0.09659  -0.0257   1.0000   0.0762
  -7.250  -0.3105   0.09957   0.09314  -0.0241   1.0000   0.0794
  -7.000  -0.3093   0.09735   0.09101  -0.0239   1.0000   0.0830
  -6.750  -0.3106   0.09592   0.08967  -0.0252   1.0000   0.0864
  -6.500  -0.3110   0.09555   0.08939  -0.0295   1.0000   0.0882
  -6.000  -0.3060   0.08870   0.08268  -0.0268   1.0000   0.0908
  -5.750  -0.3031   0.08589   0.07991  -0.0253   1.0000   0.0934
  -5.500  -0.2986   0.08348   0.07753  -0.0255   1.0000   0.0966
  -5.250  -0.2869   0.08176   0.07578  -0.0304   1.0000   0.1016
  -5.000  -0.2761   0.07878   0.07278  -0.0330   1.0000   0.1039
  -4.750  -0.2743   0.07560   0.06967  -0.0294   1.0000   0.1083
  -4.250  -0.2434   0.07010   0.06408  -0.0352   1.0000   0.1195
  -4.000  -0.2317   0.06734   0.06131  -0.0352   1.0000   0.1243
  -3.750  -0.1929   0.06387   0.05765  -0.0433   0.9961   0.1334
  -3.500  -0.1477   0.06069   0.05422  -0.0516   0.9899   0.1468
  -3.250  -0.1175   0.05711   0.05058  -0.0549   0.9842   0.1630
  -3.000  -0.0845   0.05390   0.04729  -0.0587   0.9779   0.1799
  -2.750  -0.0478   0.05097   0.04423  -0.0632   0.9713   0.1966
  -2.250   0.0722   0.04329   0.03534  -0.0801   0.9593   0.0994
  -2.000   0.1233   0.04035   0.03182  -0.0856   0.9533   0.0875
  -1.750   0.1623   0.03822   0.02944  -0.0891   0.9455   0.0904
  -1.500   0.2044   0.03635   0.02733  -0.0928   0.9386   0.0932
  -1.250   0.2490   0.03462   0.02492  -0.0959   0.9295   0.0884
  -1.000   0.2949   0.03273   0.02277  -0.0998   0.9227   0.0872
  -0.750   0.3330   0.03132   0.02106  -0.1019   0.9119   0.0865
  -0.500   0.3740   0.03000   0.01944  -0.1042   0.9025   0.0862
  -0.250   0.4185   0.02874   0.01786  -0.1070   0.8943   0.0865
   0.000   0.4546   0.02781   0.01666  -0.1082   0.8830   0.0875
   0.250   0.4915   0.02692   0.01555  -0.1093   0.8722   0.0897
   0.500   0.5293   0.02618   0.01472  -0.1106   0.8612   0.0984
   0.750   0.5664   0.02537   0.01378  -0.1115   0.8491   0.1061
   1.000   0.5995   0.02463   0.01296  -0.1115   0.8345   0.1116
   1.250   0.6297   0.02400   0.01226  -0.1110   0.8175   0.1185
   1.500   0.6585   0.02345   0.01169  -0.1103   0.7988   0.1310
   1.750   0.6881   0.02285   0.01121  -0.1097   0.7792   0.1597
   2.000   0.7181   0.02203   0.01067  -0.1092   0.7591   0.2248
   2.250   0.7393   0.02046   0.01028  -0.1069   0.7358   1.0000
   2.500   0.7653   0.02039   0.00994  -0.1055   0.7100   1.0000
   2.750   0.7918   0.02033   0.00967  -0.1043   0.6832   1.0000
   3.000   0.8188   0.02033   0.00950  -0.1032   0.6579   1.0000
   3.250   0.8444   0.02048   0.00951  -0.1022   0.6327   1.0000
   3.500   0.8700   0.02068   0.00957  -0.1011   0.6072   1.0000
   3.750   0.8947   0.02095   0.00973  -0.1000   0.5798   1.0000
   4.000   0.9185   0.02129   0.00995  -0.0989   0.5502   1.0000
   4.250   0.9415   0.02169   0.01024  -0.0976   0.5190   1.0000
   4.500   0.9640   0.02216   0.01057  -0.0964   0.4873   1.0000
   4.750   0.9860   0.02270   0.01099  -0.0951   0.4566   1.0000
   5.000   1.0076   0.02333   0.01147  -0.0938   0.4281   1.0000
   5.250   1.0291   0.02404   0.01204  -0.0927   0.4026   1.0000
   5.500   1.0509   0.02478   0.01272  -0.0916   0.3803   1.0000
   5.750   1.0731   0.02557   0.01348  -0.0907   0.3622   1.0000
   6.000   1.0961   0.02636   0.01428  -0.0899   0.3471   1.0000
   6.250   1.1198   0.02716   0.01514  -0.0893   0.3342   1.0000
   6.500   1.1443   0.02798   0.01605  -0.0887   0.3239   1.0000
   6.750   1.1688   0.02880   0.01695  -0.0882   0.3155   1.0000
   7.000   1.1886   0.02954   0.01776  -0.0870   0.3006   1.0000
   7.250   1.2030   0.03023   0.01853  -0.0852   0.2788   1.0000
   7.500   1.2201   0.03097   0.01936  -0.0838   0.2631   1.0000
   7.750   1.2386   0.03174   0.02028  -0.0825   0.2511   1.0000
   8.000   1.2560   0.03254   0.02130  -0.0811   0.2389   1.0000
   8.250   1.2733   0.03337   0.02230  -0.0796   0.2285   1.0000
   8.500   1.2855   0.03428   0.02336  -0.0777   0.2098   1.0000
   8.750   1.2971   0.03530   0.02452  -0.0757   0.1926   1.0000
   9.000   1.3036   0.03663   0.02586  -0.0734   0.1656   1.0000
   9.250   1.3056   0.03838   0.02736  -0.0709   0.0799   1.0000
   9.500   1.3038   0.04115   0.02986  -0.0682   0.0553   1.0000
   9.750   1.3046   0.04377   0.03254  -0.0658   0.0458   1.0000
  10.000   1.3047   0.04639   0.03532  -0.0637   0.0411   1.0000
  10.250   1.3011   0.04941   0.03854  -0.0619   0.0383   1.0000
  10.500   1.2975   0.05247   0.04179  -0.0605   0.0362   1.0000
  10.750   1.2921   0.05588   0.04542  -0.0595   0.0347   1.0000
  11.000   1.2845   0.05972   0.04948  -0.0591   0.0336   1.0000
  11.250   1.2754   0.06401   0.05398  -0.0592   0.0329   1.0000
  11.500   1.2651   0.06874   0.05892  -0.0599   0.0323   1.0000
  11.750   1.2540   0.07386   0.06425  -0.0610   0.0320   1.0000
  12.000   1.2429   0.07926   0.06985  -0.0625   0.0317   1.0000
  12.250   1.2319   0.08478   0.07555  -0.0642   0.0315   1.0000
  12.500   1.2218   0.09026   0.08118  -0.0660   0.0312   1.0000
  12.750   1.2128   0.09554   0.08660  -0.0675   0.0310   1.0000
  13.000   1.2054   0.10051   0.09168  -0.0689   0.0307   1.0000
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