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GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 6K AIRFOIL (goe06k-il)
Reynolds number: 500,000
Max Cl/Cd: 117.88 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe06k-il-500000.txt
Download as CSV file: xf-goe06k-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 6K 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4693   0.11067   0.10843  -0.0240   1.0000   0.0143
  -9.250  -0.4731   0.10815   0.10594  -0.0237   1.0000   0.0145
  -9.000  -0.4784   0.10566   0.10349  -0.0235   1.0000   0.0146
  -8.750  -0.4702   0.10172   0.09955  -0.0276   0.9986   0.0148
  -8.500  -0.4593   0.09745   0.09529  -0.0321   0.9966   0.0149
  -8.250  -0.4482   0.09301   0.09084  -0.0371   0.9947   0.0149
  -8.000  -0.4419   0.08884   0.08669  -0.0414   0.9910   0.0150
  -7.750  -0.4334   0.08437   0.08221  -0.0471   0.9871   0.0150
  -7.500  -0.4126   0.07818   0.07598  -0.0559   0.9846   0.0150
  -7.250  -0.4023   0.07345   0.07121  -0.0600   0.9793   0.0151
  -7.000  -0.3855   0.06779   0.06547  -0.0655   0.9756   0.0151
  -6.750  -0.3766   0.05867   0.05623  -0.0723   0.9729   0.0156
  -6.500  -0.3688   0.05500   0.05249  -0.0729   0.9681   0.0160
  -6.250  -0.3537   0.05170   0.04911  -0.0742   0.9643   0.0166
  -6.000  -0.3322   0.04787   0.04514  -0.0767   0.9621   0.0171
  -5.750  -0.3079   0.04400   0.04110  -0.0792   0.9605   0.0180
  -5.500  -0.2948   0.04079   0.03772  -0.0782   0.9555   0.0191
  -5.250  -0.2643   0.03885   0.03543  -0.0777   0.9521   0.0227
  -5.000  -0.2406   0.03543   0.03169  -0.0781   0.9498   0.0228
  -4.750  -0.2265   0.02815   0.02389  -0.0779   0.9478   0.0239
  -4.000  -0.1688   0.01989   0.01469  -0.0729   0.9378   0.0199
  -3.750  -0.1420   0.01692   0.01124  -0.0721   0.9366   0.0193
  -3.500  -0.1114   0.01542   0.00949  -0.0724   0.9357   0.0207
  -3.250  -0.0786   0.01483   0.00875  -0.0734   0.9349   0.0225
  -3.000  -0.0495   0.01310   0.00684  -0.0737   0.9343   0.0249
  -2.750  -0.0358   0.01276   0.00647  -0.0706   0.9282   0.0269
  -2.500  -0.0078   0.01228   0.00595  -0.0706   0.9261   0.0298
  -2.250   0.0238   0.01201   0.00565  -0.0715   0.9247   0.0335
  -2.000   0.0539   0.01132   0.00493  -0.0721   0.9235   0.0415
  -1.750   0.0861   0.01086   0.00445  -0.0731   0.9226   0.0508
  -1.500   0.1192   0.01050   0.00408  -0.0743   0.9219   0.0628
  -1.250   0.1529   0.01015   0.00377  -0.0757   0.9213   0.0786
  -1.000   0.1654   0.00998   0.00376  -0.0725   0.9148   0.1262
  -0.750   0.1783   0.00838   0.00373  -0.0700   0.9118   0.6113
  -0.500   0.3715   0.00800   0.00433  -0.1079   0.9346   1.0000
  -0.250   0.3883   0.00800   0.00430  -0.1056   0.9286   1.0000
   0.000   0.4199   0.00791   0.00418  -0.1066   0.9263   1.0000
   0.250   0.4570   0.00778   0.00404  -0.1088   0.9248   1.0000
   0.500   0.5010   0.00753   0.00379  -0.1124   0.9233   1.0000
   0.750   0.5553   0.00709   0.00334  -0.1182   0.9209   1.0000
   1.000   0.5786   0.00695   0.00321  -0.1170   0.9122   1.0000
   1.250   0.6217   0.00674   0.00302  -0.1204   0.9082   1.0000
   1.500   0.6511   0.00657   0.00285  -0.1206   0.8965   1.0000
   1.750   0.6833   0.00639   0.00266  -0.1213   0.8813   1.0000
   2.000   0.7078   0.00633   0.00258  -0.1203   0.8649   1.0000
   2.250   0.7256   0.00628   0.00247  -0.1177   0.8354   1.0000
   2.500   0.7438   0.00631   0.00241  -0.1152   0.7947   1.0000
   2.750   0.7594   0.00651   0.00240  -0.1122   0.7402   1.0000
   3.000   0.7610   0.00705   0.00249  -0.1062   0.6424   1.0000
   3.250   0.7440   0.00822   0.00285  -0.0963   0.4772   1.0000
   3.500   0.7391   0.00940   0.00330  -0.0896   0.3174   1.0000
   3.750   0.7396   0.01061   0.00378  -0.0843   0.1602   1.0000
   4.000   0.7494   0.01144   0.00420  -0.0808   0.0767   1.0000
   4.250   0.7655   0.01193   0.00462  -0.0784   0.0522   1.0000
   4.500   0.7817   0.01243   0.00504  -0.0760   0.0352   1.0000
   4.750   0.7983   0.01293   0.00559  -0.0736   0.0291   1.0000
   5.000   0.8145   0.01347   0.00613  -0.0712   0.0245   1.0000
   5.250   0.8272   0.01428   0.00703  -0.0681   0.0217   1.0000
   5.500   0.8441   0.01478   0.00758  -0.0658   0.0199   1.0000
   5.750   0.8597   0.01548   0.00838  -0.0633   0.0185   1.0000
   6.000   0.8766   0.01625   0.00920  -0.0611   0.0171   1.0000
   6.250   0.8956   0.01730   0.01032  -0.0595   0.0159   1.0000
   6.500   0.9291   0.02065   0.01388  -0.0610   0.0142   1.0000
   6.750   0.9502   0.02120   0.01454  -0.0596   0.0137   1.0000
   7.000   0.9761   0.02291   0.01646  -0.0592   0.0132   1.0000
   7.250   1.0002   0.02509   0.01895  -0.0583   0.0126   1.0000
   7.500   1.0194   0.02864   0.02292  -0.0563   0.0127   1.0000
   7.750   1.0333   0.03325   0.02785  -0.0535   0.0150   1.0000
   8.000   1.0226   0.04247   0.03804  -0.0443   0.0205   1.0000
   8.250   1.0287   0.04487   0.04071  -0.0403   0.0191   1.0000
   8.500   1.0304   0.04754   0.04361  -0.0361   0.0179   1.0000
   8.750   1.0313   0.04989   0.04614  -0.0323   0.0171   1.0000
   9.000   1.0321   0.05170   0.04807  -0.0289   0.0163   1.0000
   9.250   1.0317   0.05338   0.04982  -0.0256   0.0158   1.0000
   9.500   1.0274   0.05561   0.05215  -0.0219   0.0154   1.0000
   9.750   1.0148   0.05857   0.05529  -0.0167   0.0154   1.0000
  10.000   0.9982   0.06103   0.05786  -0.0116   0.0151   1.0000
  10.250   0.9800   0.06399   0.06095  -0.0071   0.0150   1.0000
  10.500   0.9565   0.06767   0.06477  -0.0032   0.0148   1.0000
  10.750   0.9321   0.07180   0.06903  -0.0005   0.0147   1.0000
  11.000   0.9157   0.07542   0.07278   0.0009   0.0148   1.0000
  11.250   0.8983   0.07961   0.07708   0.0011   0.0148   1.0000
  11.500   0.8746   0.08547   0.08305  -0.0005   0.0147   1.0000
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