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GOE 758 AIRFOIL (goe758-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 758 AIRFOIL (goe758-il)
Reynolds number: 50,000
Max Cl/Cd: 25.8 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe758-il-50000-n5.txt
Download as CSV file: xf-goe758-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 758 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.1984   0.09440   0.08845  -0.0534   0.9722   0.0918
  -8.500  -0.3247   0.09516   0.08902  -0.0492   0.9890   0.0803
  -8.250  -0.3066   0.09070   0.08455  -0.0535   0.9776   0.0796
  -8.000  -0.2907   0.08622   0.08004  -0.0581   0.9651   0.0782
  -7.750  -0.2787   0.08114   0.07494  -0.0640   0.9511   0.0763
  -7.500  -0.2683   0.07438   0.06811  -0.0727   0.9365   0.0740
  -7.250  -0.2677   0.06573   0.05921  -0.0828   0.9212   0.0711
  -7.000  -0.2677   0.05791   0.05075  -0.0902   0.9063   0.0688
  -6.750  -0.2481   0.05345   0.04588  -0.0939   0.8967   0.0688
  -6.500  -0.2315   0.05011   0.04224  -0.0954   0.8844   0.0696
  -6.250  -0.2111   0.04769   0.03965  -0.0962   0.8732   0.0707
  -6.000  -0.1852   0.04466   0.03622  -0.0982   0.8653   0.0714
  -5.750  -0.1683   0.04219   0.03332  -0.0980   0.8531   0.0713
  -5.500  -0.1465   0.03992   0.03063  -0.0981   0.8431   0.0714
  -5.250  -0.1204   0.03784   0.02815  -0.0985   0.8346   0.0719
  -5.000  -0.0990   0.03629   0.02627  -0.0980   0.8244   0.0724
  -4.750  -0.0698   0.03470   0.02436  -0.0984   0.8173   0.0737
  -4.500  -0.0485   0.03369   0.02311  -0.0975   0.8071   0.0756
  -4.250  -0.0180   0.03246   0.02155  -0.0979   0.8006   0.0783
  -4.000   0.0040   0.03166   0.02048  -0.0969   0.7910   0.0800
  -3.750   0.0339   0.03059   0.01927  -0.0970   0.7847   0.0817
  -3.500   0.0556   0.02997   0.01859  -0.0960   0.7757   0.0837
  -3.250   0.0836   0.02924   0.01774  -0.0958   0.7692   0.0866
  -3.000   0.1071   0.02878   0.01713  -0.0950   0.7612   0.0901
  -2.750   0.1339   0.02825   0.01647  -0.0947   0.7541   0.0959
  -2.500   0.1601   0.02777   0.01591  -0.0945   0.7476   0.1053
  -2.250   0.1828   0.02736   0.01548  -0.0937   0.7397   0.1178
  -2.000   0.2131   0.02641   0.01468  -0.0941   0.7347   0.1553
  -1.750   0.2298   0.02577   0.01475  -0.0930   0.7259   0.2984
  -1.500   0.2511   0.02497   0.01483  -0.0914   0.7205   0.5122
  -1.250   0.2645   0.02479   0.01514  -0.0880   0.7128   0.6572
  -1.000   0.3405   0.02386   0.01482  -0.0942   0.7076   0.9544
  -0.750   0.3846   0.02379   0.01432  -0.0975   0.7025   1.0000
  -0.500   0.3983   0.02427   0.01457  -0.0957   0.6935   1.0000
  -0.250   0.4263   0.02431   0.01429  -0.0957   0.6881   1.0000
   0.000   0.4426   0.02482   0.01461  -0.0942   0.6795   1.0000
   0.250   0.4698   0.02494   0.01448  -0.0939   0.6736   1.0000
   0.500   0.4899   0.02538   0.01473  -0.0929   0.6659   1.0000
   0.750   0.5145   0.02563   0.01479  -0.0923   0.6591   1.0000
   1.000   0.5406   0.02587   0.01483  -0.0919   0.6530   1.0000
   1.250   0.5597   0.02639   0.01524  -0.0908   0.6447   1.0000
   1.500   0.5910   0.02642   0.01508  -0.0910   0.6397   1.0000
   1.750   0.6051   0.02720   0.01579  -0.0892   0.6305   1.0000
   2.000   0.6340   0.02735   0.01579  -0.0892   0.6248   1.0000
   2.250   0.6514   0.02803   0.01641  -0.0878   0.6167   1.0000
   2.500   0.6764   0.02836   0.01664  -0.0873   0.6099   1.0000
   2.750   0.7002   0.02878   0.01697  -0.0867   0.6033   1.0000
   3.000   0.7174   0.02948   0.01766  -0.0853   0.5950   1.0000
   3.250   0.7501   0.02950   0.01756  -0.0856   0.5901   1.0000
   3.500   0.7576   0.03066   0.01876  -0.0832   0.5801   1.0000
   3.750   0.7878   0.03078   0.01881  -0.0833   0.5747   1.0000
   4.000   0.7965   0.03192   0.01999  -0.0811   0.5652   1.0000
   4.250   0.8241   0.03214   0.02016  -0.0808   0.5593   1.0000
   4.500   0.8341   0.03324   0.02130  -0.0788   0.5504   1.0000
   4.750   0.8577   0.03365   0.02171  -0.0781   0.5439   1.0000
   5.000   0.8706   0.03462   0.02270  -0.0764   0.5359   1.0000
   5.250   0.8889   0.03529   0.02339  -0.0752   0.5286   1.0000
   5.500   0.9090   0.03589   0.02403  -0.0743   0.5221   1.0000
   5.750   0.9161   0.03716   0.02535  -0.0720   0.5134   1.0000
   6.000   0.9520   0.03690   0.02509  -0.0725   0.5094   1.0000
   6.250   0.9384   0.03931   0.02758  -0.0685   0.4982   1.0000
   6.500   0.9735   0.03907   0.02739  -0.0689   0.4943   1.0000
   6.750   0.9519   0.04193   0.03032  -0.0643   0.4830   1.0000
   7.000   0.9868   0.04166   0.03008  -0.0646   0.4792   1.0000
   7.250   0.9612   0.04519   0.03367  -0.0604   0.4676   1.0000
   7.500   0.9941   0.04490   0.03346  -0.0603   0.4640   1.0000
   8.000   0.9977   0.04896   0.03765  -0.0569   0.4487   1.0000
   8.500   0.9984   0.05386   0.04268  -0.0544   0.4331   1.0000
   9.000   0.9970   0.05956   0.04853  -0.0528   0.4176   1.0000
  10.000   0.9876   0.07328   0.06258  -0.0514   0.3871   1.0000
  10.500   0.9802   0.08081   0.07027  -0.0514   0.3723   1.0000
  10.750   1.0061   0.08088   0.07046  -0.0506   0.3700   1.0000
  11.000   0.9706   0.08892   0.07854  -0.0519   0.3584   1.0000
  11.250   0.9933   0.08941   0.07919  -0.0512   0.3554   1.0000
  11.500   0.9669   0.09640   0.08623  -0.0525   0.3455   1.0000
  11.750   0.9809   0.09813   0.08809  -0.0522   0.3414   1.0000
  12.000   1.0064   0.09824   0.08836  -0.0514   0.3388   1.0000
  12.250   0.9758   0.10587   0.09604  -0.0532   0.3280   1.0000
  12.500   1.0006   0.10588   0.09621  -0.0523   0.3242   1.0000
  13.500   1.0370   0.11251   0.10339  -0.0510   0.2928   1.0000
  14.000   1.0314   0.11974   0.11085  -0.0524   0.2747   1.0000
  14.500   1.0282   0.12681   0.11819  -0.0540   0.2560   1.0000
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