Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 564 AIRFOIL (goe564-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 564 AIRFOIL (goe564-il)
Reynolds number: 500,000
Max Cl/Cd: 77.57 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe564-il-500000-n5.txt
Download as CSV file: xf-goe564-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 564 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.6937   0.06821   0.06593  -0.0462   1.0000   0.0094
 -11.000  -0.8854   0.03237   0.02923  -0.0599   1.0000   0.0078
 -10.750  -0.8749   0.02880   0.02530  -0.0603   0.9988   0.0081
 -10.500  -0.8535   0.02603   0.02217  -0.0617   0.9965   0.0084
 -10.250  -0.8293   0.02393   0.01977  -0.0629   0.9946   0.0086
 -10.000  -0.8064   0.02210   0.01770  -0.0634   0.9921   0.0092
  -9.750  -0.7803   0.02087   0.01629  -0.0641   0.9899   0.0097
  -9.500  -0.7526   0.01988   0.01515  -0.0649   0.9880   0.0103
  -9.250  -0.7240   0.01896   0.01406  -0.0657   0.9865   0.0111
  -9.000  -0.6967   0.01804   0.01294  -0.0662   0.9847   0.0118
  -8.750  -0.6715   0.01713   0.01192  -0.0662   0.9817   0.0127
  -8.500  -0.6428   0.01663   0.01136  -0.0668   0.9796   0.0138
  -8.250  -0.6128   0.01610   0.01075  -0.0675   0.9778   0.0149
  -8.000  -0.5814   0.01559   0.01009  -0.0685   0.9763   0.0162
  -7.750  -0.5529   0.01512   0.00958  -0.0689   0.9734   0.0176
  -7.500  -0.5239   0.01491   0.00935  -0.0693   0.9699   0.0190
  -7.250  -0.4932   0.01455   0.00891  -0.0701   0.9676   0.0206
  -7.000  -0.4616   0.01417   0.00843  -0.0710   0.9658   0.0219
  -6.750  -0.4291   0.01384   0.00799  -0.0721   0.9645   0.0231
  -6.500  -0.4041   0.01320   0.00729  -0.0717   0.9602   0.0247
  -6.250  -0.3754   0.01284   0.00690  -0.0720   0.9569   0.0261
  -6.000  -0.3442   0.01251   0.00651  -0.0728   0.9544   0.0277
  -5.500  -0.2802   0.01182   0.00566  -0.0747   0.9500   0.0302
  -5.250  -0.2547   0.01156   0.00535  -0.0741   0.9445   0.0308
  -5.000  -0.2256   0.01103   0.00475  -0.0745   0.9407   0.0320
  -4.750  -0.1940   0.01055   0.00423  -0.0753   0.9377   0.0336
  -4.500  -0.1660   0.01022   0.00385  -0.0753   0.9324   0.0346
  -4.250  -0.1371   0.00992   0.00351  -0.0755   0.9271   0.0357
  -4.000  -0.1056   0.00964   0.00319  -0.0762   0.9231   0.0371
  -3.750  -0.0770   0.00941   0.00293  -0.0763   0.9177   0.0389
  -3.500  -0.0482   0.00921   0.00268  -0.0765   0.9120   0.0406
  -3.250  -0.0175   0.00898   0.00243  -0.0770   0.9075   0.0443
  -3.000   0.0097   0.00877   0.00225  -0.0768   0.9011   0.0515
  -2.750   0.0384   0.00854   0.00211  -0.0770   0.8951   0.0710
  -2.500   0.0670   0.00840   0.00200  -0.0771   0.8890   0.0853
  -2.250   0.0951   0.00830   0.00189  -0.0770   0.8821   0.0938
  -2.000   0.1234   0.00820   0.00179  -0.0770   0.8738   0.1022
  -1.750   0.1517   0.00812   0.00167  -0.0769   0.8621   0.1094
  -1.500   0.1789   0.00805   0.00157  -0.0766   0.8471   0.1175
  -1.250   0.2056   0.00799   0.00148  -0.0762   0.8322   0.1255
  -1.000   0.2320   0.00795   0.00140  -0.0757   0.8172   0.1354
  -0.750   0.2580   0.00788   0.00135  -0.0752   0.8024   0.1515
  -0.250   0.3071   0.00764   0.00124  -0.0737   0.7577   0.2325
   0.000   0.3290   0.00748   0.00119  -0.0724   0.7170   0.3204
   0.250   0.3479   0.00734   0.00117  -0.0705   0.6654   0.4401
   0.500   0.3670   0.00713   0.00122  -0.0687   0.6330   0.5779
   0.750   0.3855   0.00685   0.00129  -0.0667   0.6123   0.7189
   1.000   0.4048   0.00659   0.00140  -0.0645   0.5918   0.8582
   1.250   0.4739   0.00679   0.00158  -0.0734   0.5491   0.9755
   1.500   0.5211   0.00711   0.00167  -0.0778   0.4967   0.9975
   1.750   0.5473   0.00743   0.00176  -0.0776   0.4459   1.0000
   2.000   0.5668   0.00777   0.00187  -0.0759   0.3989   1.0000
   2.250   0.5868   0.00811   0.00201  -0.0743   0.3539   1.0000
   2.500   0.6071   0.00845   0.00216  -0.0728   0.3107   1.0000
   2.750   0.6279   0.00879   0.00232  -0.0714   0.2738   1.0000
   3.000   0.6496   0.00906   0.00247  -0.0702   0.2487   1.0000
   3.250   0.6717   0.00933   0.00264  -0.0690   0.2290   1.0000
   3.500   0.6941   0.00957   0.00281  -0.0679   0.2129   1.0000
   4.000   0.7397   0.01001   0.00315  -0.0658   0.1882   1.0000
   4.250   0.7631   0.01020   0.00332  -0.0649   0.1809   1.0000
   4.500   0.7864   0.01040   0.00351  -0.0640   0.1735   1.0000
   4.750   0.8098   0.01060   0.00370  -0.0631   0.1660   1.0000
   5.000   0.8331   0.01082   0.00391  -0.0622   0.1578   1.0000
   5.250   0.8564   0.01104   0.00413  -0.0612   0.1472   1.0000
   5.500   0.8789   0.01133   0.00435  -0.0602   0.1297   1.0000
   5.750   0.8995   0.01179   0.00464  -0.0590   0.0998   1.0000
   6.000   0.9210   0.01218   0.00496  -0.0578   0.0875   1.0000
   6.250   0.9430   0.01253   0.00531  -0.0567   0.0804   1.0000
   6.500   0.9656   0.01283   0.00563  -0.0558   0.0745   1.0000
   6.750   0.9874   0.01320   0.00598  -0.0547   0.0635   1.0000
   7.000   1.0069   0.01378   0.00640  -0.0533   0.0389   1.0000
   7.250   1.0274   0.01429   0.00688  -0.0520   0.0308   1.0000
   7.500   1.0479   0.01479   0.00740  -0.0507   0.0261   1.0000
   7.750   1.0693   0.01519   0.00786  -0.0496   0.0229   1.0000
   8.000   1.0898   0.01567   0.00836  -0.0484   0.0193   1.0000
   8.250   1.1100   0.01619   0.00891  -0.0471   0.0168   1.0000
   8.500   1.1303   0.01666   0.00943  -0.0458   0.0151   1.0000
   8.750   1.1496   0.01722   0.01001  -0.0444   0.0135   1.0000
   9.000   1.1673   0.01792   0.01077  -0.0428   0.0122   1.0000
   9.250   1.1865   0.01843   0.01136  -0.0414   0.0113   1.0000
   9.500   1.2046   0.01901   0.01201  -0.0399   0.0106   1.0000
   9.750   1.2218   0.01964   0.01270  -0.0382   0.0098   1.0000
  10.000   1.2371   0.02034   0.01345  -0.0363   0.0093   1.0000
  10.250   1.2480   0.02125   0.01445  -0.0336   0.0087   1.0000
  10.500   1.2605   0.02200   0.01529  -0.0312   0.0084   1.0000
  10.750   1.2726   0.02278   0.01619  -0.0288   0.0081   1.0000
  11.000   1.2842   0.02359   0.01710  -0.0265   0.0078   1.0000
  11.250   1.2958   0.02440   0.01800  -0.0242   0.0074   1.0000
  11.500   1.3060   0.02531   0.01900  -0.0219   0.0071   1.0000
  11.750   1.3176   0.02613   0.01989  -0.0199   0.0068   1.0000
  12.000   1.3247   0.02727   0.02112  -0.0175   0.0065   1.0000
  12.250   1.3287   0.02866   0.02260  -0.0149   0.0063   1.0000
  12.500   1.3303   0.03026   0.02432  -0.0122   0.0061   1.0000
  12.750   1.3346   0.03172   0.02592  -0.0101   0.0060   1.0000
  13.000   1.3372   0.03338   0.02770  -0.0081   0.0059   1.0000
  13.250   1.3390   0.03519   0.02966  -0.0062   0.0058   1.0000
  13.500   1.3399   0.03718   0.03180  -0.0046   0.0056   1.0000
  13.750   1.3385   0.03947   0.03424  -0.0032   0.0056   1.0000
  14.000   1.3352   0.04211   0.03703  -0.0022   0.0055   1.0000
  14.250   1.3323   0.04488   0.03994  -0.0015   0.0054   1.0000
  14.500   1.3269   0.04813   0.04334  -0.0013   0.0053   1.0000
  14.750   1.3203   0.05177   0.04713  -0.0017   0.0052   1.0000
  15.000   1.3143   0.05563   0.05113  -0.0026   0.0051   1.0000
  15.250   1.3018   0.06065   0.05632  -0.0041   0.0051   1.0000
  15.500   1.2908   0.06579   0.06161  -0.0061   0.0051   1.0000
  15.750   1.2775   0.07166   0.06765  -0.0088   0.0050   1.0000
  16.000   1.2656   0.07769   0.07382  -0.0119   0.0049   1.0000
  16.250   1.2465   0.08532   0.08162  -0.0158   0.0050   1.0000
  16.500   1.2298   0.09288   0.08933  -0.0199   0.0049   1.0000
  16.750   1.2114   0.10098   0.09757  -0.0243   0.0049   1.0000
  17.000   1.1897   0.10988   0.10663  -0.0291   0.0050   1.0000
  17.250   1.1709   0.11840   0.11527  -0.0338   0.0050   1.0000
<< Back to GOE 564 AIRFOIL (goe564-il)

Polar data table (+)

Polar graphs


<< Back to GOE 564 AIRFOIL (goe564-il)