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GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 565 AIRFOIL (goe565-il)
Reynolds number: 50,000
Max Cl/Cd: 37.39 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe565-il-50000.txt
Download as CSV file: xf-goe565-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 565 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4081   0.10230   0.09511  -0.0170   1.0000   0.2575
  -8.250  -0.4083   0.09955   0.09244  -0.0167   1.0000   0.2675
  -8.000  -0.4270   0.09919   0.09223  -0.0165   1.0000   0.2776
  -7.750  -0.4132   0.09507   0.08811  -0.0151   1.0000   0.2913
  -7.500  -0.4069   0.09176   0.08484  -0.0137   1.0000   0.3049
  -7.250  -0.4063   0.08917   0.08232  -0.0121   1.0000   0.3198
  -7.000  -0.4104   0.08702   0.08027  -0.0100   1.0000   0.3357
  -6.750  -0.4164   0.08500   0.07836  -0.0076   1.0000   0.3516
  -6.500  -0.4252   0.08314   0.07663  -0.0051   1.0000   0.3675
  -6.250  -0.4245   0.08028   0.07385  -0.0028   1.0000   0.3846
  -6.000  -0.4115   0.07675   0.07033  -0.0003   1.0000   0.4032
  -5.750  -0.4116   0.07446   0.06811   0.0027   1.0000   0.4247
  -5.500  -0.4070   0.07178   0.06549   0.0061   1.0000   0.4507
  -5.250  -0.4061   0.06944   0.06320   0.0098   1.0000   0.4792
  -4.750  -0.3966   0.04628   0.03863  -0.0397   1.0000   0.1905
  -4.500  -0.3772   0.04243   0.03452  -0.0401   1.0000   0.1839
  -4.250  -0.3520   0.03761   0.02884  -0.0422   1.0000   0.1730
  -4.000  -0.3292   0.03470   0.02536  -0.0424   1.0000   0.1740
  -3.750  -0.3068   0.03229   0.02262  -0.0421   1.0000   0.1768
  -3.500  -0.2827   0.03013   0.02004  -0.0417   1.0000   0.1778
  -3.250  -0.2590   0.02845   0.01810  -0.0411   1.0000   0.1815
  -3.000  -0.2344   0.02712   0.01631  -0.0406   1.0000   0.1896
  -2.750  -0.2111   0.02587   0.01504  -0.0398   1.0000   0.1966
  -2.500  -0.1852   0.02478   0.01360  -0.0393   1.0000   0.2042
  -2.250  -0.1619   0.02389   0.01270  -0.0384   1.0000   0.2165
  -2.000  -0.1388   0.02318   0.01204  -0.0375   1.0000   0.2365
  -1.750  -0.1130   0.02243   0.01133  -0.0370   1.0000   0.2658
  -1.500  -0.0863   0.02104   0.01052  -0.0366   1.0000   0.3587
  -1.250  -0.0576   0.01813   0.00976  -0.0344   1.0000   1.0000
  -1.000  -0.0356   0.01838   0.00953  -0.0339   1.0000   1.0000
  -0.750  -0.0145   0.01865   0.00946  -0.0333   1.0000   1.0000
  -0.500   0.0062   0.01897   0.00947  -0.0327   1.0000   1.0000
  -0.250   0.0266   0.01931   0.00958  -0.0321   1.0000   1.0000
   0.000   0.0467   0.01969   0.00976  -0.0316   1.0000   1.0000
   0.250   0.0665   0.02011   0.01001  -0.0310   1.0000   1.0000
   0.500   0.0860   0.02058   0.01032  -0.0305   1.0000   1.0000
   0.750   0.1051   0.02108   0.01070  -0.0300   1.0000   1.0000
   1.000   0.1238   0.02165   0.01116  -0.0295   1.0000   1.0000
   1.250   0.1420   0.02226   0.01170  -0.0291   1.0000   1.0000
   1.500   0.1598   0.02294   0.01231  -0.0288   1.0000   1.0000
   1.750   0.1770   0.02369   0.01301  -0.0285   1.0000   1.0000
   2.000   0.1937   0.02451   0.01380  -0.0283   1.0000   1.0000
   2.250   0.2487   0.02612   0.01541  -0.0352   0.9826   1.0000
   2.500   0.3023   0.02743   0.01676  -0.0416   0.9604   1.0000
   2.750   0.3561   0.02864   0.01803  -0.0475   0.9393   1.0000
   3.000   0.4034   0.02956   0.01907  -0.0520   0.9163   1.0000
   3.250   0.4527   0.03033   0.01997  -0.0564   0.8932   1.0000
   3.500   0.5126   0.03061   0.02045  -0.0615   0.8669   1.0000
   3.750   0.5688   0.03023   0.02029  -0.0649   0.8357   1.0000
   4.000   0.6301   0.02927   0.01959  -0.0681   0.8061   1.0000
   4.250   0.6882   0.02820   0.01884  -0.0706   0.7808   1.0000
   4.500   0.7342   0.02726   0.01815  -0.0710   0.7527   1.0000
   4.750   0.7810   0.02609   0.01726  -0.0710   0.7223   1.0000
   5.000   0.8177   0.02531   0.01665  -0.0695   0.6853   1.0000
   5.250   0.8534   0.02460   0.01603  -0.0677   0.6432   1.0000
   5.500   0.8825   0.02436   0.01581  -0.0653   0.5950   1.0000
   5.750   0.9105   0.02436   0.01557  -0.0628   0.5434   1.0000
   6.000   0.9296   0.02486   0.01586  -0.0597   0.4892   1.0000
   6.250   0.9463   0.02541   0.01612  -0.0565   0.4375   1.0000
   6.500   0.9591   0.02603   0.01658  -0.0532   0.3873   1.0000
   6.750   0.9723   0.02690   0.01718  -0.0501   0.3381   1.0000
   7.000   0.9825   0.02830   0.01827  -0.0468   0.2793   1.0000
   7.250   0.9921   0.03061   0.02009  -0.0434   0.2118   1.0000
   7.500   1.0126   0.03323   0.02234  -0.0419   0.1734   1.0000
   7.750   1.0382   0.03596   0.02512  -0.0410   0.1532   1.0000
   8.000   1.0599   0.03842   0.02772  -0.0399   0.1390   1.0000
   8.250   1.0818   0.04140   0.03094  -0.0388   0.1305   1.0000
   8.500   1.1016   0.04406   0.03380  -0.0375   0.1227   1.0000
   8.750   1.1143   0.04730   0.03748  -0.0355   0.1172   1.0000
   9.000   1.1233   0.05097   0.04168  -0.0333   0.1142   1.0000
   9.250   1.1294   0.05489   0.04605  -0.0311   0.1125   1.0000
   9.500   1.1346   0.05865   0.05017  -0.0290   0.1105   1.0000
   9.750   1.1462   0.06253   0.05408  -0.0279   0.1071   1.0000
  10.000   1.1378   0.06691   0.05890  -0.0253   0.1067   1.0000
  10.250   1.1263   0.07146   0.06379  -0.0230   0.1068   1.0000
  10.500   1.1158   0.07623   0.06880  -0.0213   0.1073   1.0000
  10.750   1.0284   0.08318   0.07623  -0.0190   0.1161   1.0000
  11.000   1.0039   0.08943   0.08255  -0.0204   0.1182   1.0000
  11.250   0.9879   0.09590   0.08905  -0.0225   0.1197   1.0000
  11.500   0.9789   0.10239   0.09556  -0.0245   0.1206   1.0000
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