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GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 391 AIRFOIL (goe391-il)
Reynolds number: 1,000,000
Max Cl/Cd: 104.7 at α=1.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe391-il-1000000.txt
Download as CSV file: xf-goe391-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 391 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4436   0.08840   0.08684  -0.0180   1.0000   0.0074
  -7.500  -0.4549   0.08646   0.08494  -0.0161   1.0000   0.0074
  -7.250  -0.4589   0.08378   0.08228  -0.0161   1.0000   0.0074
  -7.000  -0.4472   0.07966   0.07815  -0.0204   0.9991   0.0074
  -6.750  -0.4223   0.07398   0.07244  -0.0285   0.9969   0.0075
  -6.500  -0.4054   0.06778   0.06621  -0.0344   0.9951   0.0080
  -6.250  -0.3811   0.06399   0.06238  -0.0394   0.9933   0.0083
  -6.000  -0.3557   0.06017   0.05852  -0.0444   0.9907   0.0092
  -5.750  -0.3240   0.05551   0.05378  -0.0504   0.9882   0.0106
  -5.500  -0.2842   0.05099   0.04913  -0.0568   0.9862   0.0113
  -5.250  -0.2495   0.04652   0.04452  -0.0616   0.9846   0.0114
  -5.000  -0.2148   0.04187   0.03973  -0.0661   0.9833   0.0114
  -4.750  -0.1881   0.03788   0.03559  -0.0679   0.9793   0.0114
  -4.500  -0.1587   0.03376   0.03130  -0.0699   0.9762   0.0114
  -4.000  -0.0988   0.02203   0.01885  -0.0739   0.9720   0.0104
  -3.750  -0.0656   0.01687   0.01319  -0.0745   0.9705   0.0113
  -3.500  -0.0454   0.01118   0.00667  -0.0726   0.9652   0.0124
  -3.250  -0.0138   0.01154   0.00707  -0.0735   0.9618   0.0135
  -3.000   0.0173   0.01080   0.00621  -0.0741   0.9591   0.0154
  -2.750   0.0484   0.01014   0.00541  -0.0747   0.9567   0.0166
  -2.500   0.0728   0.00890   0.00402  -0.0739   0.9517   0.0184
  -2.250   0.1013   0.00898   0.00414  -0.0741   0.9471   0.0204
  -2.000   0.1310   0.00865   0.00376  -0.0744   0.9430   0.0226
  -1.750   0.1567   0.00839   0.00346  -0.0738   0.9359   0.0243
  -1.500   0.1857   0.00855   0.00359  -0.0739   0.9300   0.0253
  -1.250   0.2087   0.00734   0.00233  -0.0729   0.9223   0.0292
  -1.000   0.2359   0.00705   0.00201  -0.0727   0.9150   0.0307
  -0.750   0.2611   0.00678   0.00171  -0.0720   0.9051   0.0316
  -0.500   0.2873   0.00656   0.00146  -0.0715   0.8962   0.0331
  -0.250   0.3140   0.00639   0.00124  -0.0712   0.8877   0.0349
   0.000   0.3400   0.00628   0.00109  -0.0706   0.8746   0.0363
   0.250   0.3652   0.00622   0.00097  -0.0699   0.8528   0.0369
   0.500   0.3905   0.00614   0.00079  -0.0692   0.8342   0.0383
   0.750   0.4157   0.00612   0.00069  -0.0685   0.8116   0.0395
   1.000   0.4394   0.00620   0.00062  -0.0675   0.7731   0.0416
   1.250   0.4603   0.00642   0.00059  -0.0658   0.7047   0.0471
   1.500   0.5256   0.00502   0.00095  -0.0755   0.5716   1.0000
   1.750   0.5454   0.00540   0.00103  -0.0738   0.5027   1.0000
   2.000   0.5622   0.00605   0.00117  -0.0717   0.3790   1.0000
   2.250   0.5822   0.00651   0.00133  -0.0702   0.3112   1.0000
   2.500   0.6038   0.00686   0.00146  -0.0690   0.2593   1.0000
   2.750   0.6219   0.00754   0.00171  -0.0673   0.1576   1.0000
   3.000   0.6415   0.00813   0.00196  -0.0657   0.0792   1.0000
   3.250   0.6632   0.00854   0.00220  -0.0645   0.0420   1.0000
   3.500   0.6869   0.00875   0.00241  -0.0637   0.0382   1.0000
   3.750   0.7102   0.00901   0.00267  -0.0627   0.0351   1.0000
   4.000   0.7335   0.00929   0.00300  -0.0617   0.0326   1.0000
   4.250   0.7565   0.00961   0.00337  -0.0607   0.0311   1.0000
   4.500   0.7784   0.01006   0.00389  -0.0595   0.0295   1.0000
   4.750   0.8014   0.01036   0.00423  -0.0586   0.0288   1.0000
   5.000   0.8257   0.01051   0.00438  -0.0579   0.0280   1.0000
   5.250   0.8494   0.01074   0.00465  -0.0572   0.0267   1.0000
   5.500   0.8729   0.01101   0.00494  -0.0564   0.0247   1.0000
   5.750   0.8966   0.01126   0.00519  -0.0556   0.0221   1.0000
   6.000   0.9161   0.01200   0.00600  -0.0541   0.0187   1.0000
   6.250   0.9419   0.01200   0.00600  -0.0538   0.0173   1.0000
   6.500   0.9671   0.01209   0.00608  -0.0534   0.0150   1.0000
   6.750   0.9858   0.01298   0.00703  -0.0517   0.0122   1.0000
   7.000   1.0094   0.01325   0.00734  -0.0509   0.0115   1.0000
   7.250   1.0332   0.01349   0.00758  -0.0503   0.0103   1.0000
   7.500   1.0562   0.01384   0.00793  -0.0495   0.0091   1.0000
   7.750   1.0725   0.01509   0.00932  -0.0474   0.0080   1.0000
   8.000   1.0882   0.01654   0.01096  -0.0453   0.0075   1.0000
   8.250   1.1088   0.01728   0.01183  -0.0441   0.0073   1.0000
   8.500   1.1295   0.01798   0.01263  -0.0429   0.0068   1.0000
   8.750   1.1481   0.01905   0.01385  -0.0415   0.0065   1.0000
   9.000   1.1667   0.02013   0.01507  -0.0400   0.0061   1.0000
   9.250   1.1863   0.02084   0.01587  -0.0388   0.0056   1.0000
   9.500   1.2036   0.02194   0.01711  -0.0373   0.0053   1.0000
   9.750   1.2159   0.02384   0.01922  -0.0351   0.0049   1.0000
  10.000   1.2113   0.02860   0.02456  -0.0306   0.0046   1.0000
  10.250   1.1975   0.03400   0.03054  -0.0253   0.0045   1.0000
  10.500   1.1799   0.03884   0.03581  -0.0200   0.0045   1.0000
  10.750   1.1142   0.02975   0.02703  -0.0108   0.0046   1.0000
  11.000   1.0978   0.03262   0.03008  -0.0072   0.0046   1.0000
  11.250   1.0730   0.03669   0.03435  -0.0043   0.0046   1.0000
  11.500   1.0505   0.04119   0.03902  -0.0031   0.0046   1.0000
  11.750   1.0323   0.04591   0.04387  -0.0033   0.0046   1.0000
  12.000   1.0108   0.05214   0.05024  -0.0051   0.0046   1.0000
  12.250   0.9953   0.05843   0.05664  -0.0080   0.0046   1.0000
  12.500   0.9769   0.06647   0.06478  -0.0124   0.0047   1.0000
  12.750   0.9592   0.07566   0.07408  -0.0177   0.0047   1.0000
  13.000   0.9467   0.08440   0.08290  -0.0227   0.0048   1.0000
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