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

GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 457 AIRFOIL (goe457-il)
Reynolds number: 200,000
Max Cl/Cd: 72.9 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe457-il-200000-n5.txt
Download as CSV file: xf-goe457-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 457 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3666   0.08794   0.08468  -0.0203   1.0000   0.0176
  -7.250  -0.3795   0.08611   0.08292  -0.0179   1.0000   0.0175
  -7.000  -0.3871   0.08341   0.08028  -0.0174   1.0000   0.0172
  -6.750  -0.3656   0.07842   0.07526  -0.0246   0.9958   0.0178
  -6.500  -0.3406   0.07307   0.06988  -0.0329   0.9905   0.0188
  -6.250  -0.3118   0.06649   0.06324  -0.0431   0.9856   0.0196
  -6.000  -0.2848   0.06003   0.05669  -0.0517   0.9796   0.0197
  -5.750  -0.2511   0.05163   0.04810  -0.0624   0.9745   0.0202
  -5.500  -0.2179   0.04268   0.03884  -0.0715   0.9687   0.0218
  -5.250  -0.1855   0.04141   0.03750  -0.0744   0.9656   0.0232
  -5.000  -0.1573   0.03606   0.03184  -0.0780   0.9590   0.0245
  -4.750  -0.1220   0.02917   0.02430  -0.0825   0.9547   0.0270
  -4.500  -0.0899   0.02469   0.01901  -0.0843   0.9498   0.0292
  -4.250  -0.0629   0.02186   0.01582  -0.0852   0.9443   0.0310
  -4.000  -0.0291   0.02090   0.01470  -0.0868   0.9407   0.0334
  -3.750   0.0049   0.01921   0.01264  -0.0881   0.9374   0.0351
  -3.500   0.0322   0.01798   0.01112  -0.0878   0.9306   0.0367
  -3.250   0.0673   0.01692   0.00975  -0.0891   0.9264   0.0387
  -3.000   0.1018   0.01592   0.00852  -0.0902   0.9215   0.0394
  -2.750   0.1313   0.01500   0.00746  -0.0902   0.9139   0.0401
  -2.500   0.1677   0.01399   0.00636  -0.0918   0.9088   0.0415
  -2.250   0.1954   0.01349   0.00584  -0.0916   0.8984   0.0436
  -2.000   0.2272   0.01296   0.00528  -0.0921   0.8886   0.0450
  -1.750   0.2595   0.01245   0.00472  -0.0927   0.8773   0.0462
  -1.500   0.2891   0.01204   0.00426  -0.0927   0.8626   0.0475
  -1.250   0.3175   0.01171   0.00387  -0.0925   0.8428   0.0494
  -1.000   0.3481   0.01140   0.00348  -0.0926   0.8159   0.0519
  -0.750   0.3821   0.01111   0.00309  -0.0934   0.7743   0.0580
  -0.500   0.4135   0.01101   0.00275  -0.0936   0.7229   0.0704
  -0.250   0.4393   0.01103   0.00257  -0.0928   0.6757   0.0970
   0.000   0.4625   0.01093   0.00253  -0.0918   0.6336   0.1719
   0.500   0.5066   0.01062   0.00264  -0.0895   0.5726   0.4397
   1.000   0.5765   0.00989   0.00274  -0.0922   0.5322   1.0000
   1.250   0.6011   0.01011   0.00280  -0.0914   0.5178   1.0000
   1.500   0.6259   0.01033   0.00288  -0.0906   0.5049   1.0000
   2.000   0.6761   0.01073   0.00308  -0.0893   0.4816   1.0000
   2.250   0.7014   0.01092   0.00320  -0.0887   0.4712   1.0000
   2.500   0.7265   0.01113   0.00332  -0.0880   0.4608   1.0000
   2.750   0.7517   0.01132   0.00347  -0.0874   0.4499   1.0000
   3.000   0.7770   0.01151   0.00362  -0.0868   0.4394   1.0000
   3.250   0.8019   0.01172   0.00378  -0.0862   0.4286   1.0000
   3.500   0.8265   0.01194   0.00394  -0.0855   0.4164   1.0000
   3.750   0.8511   0.01215   0.00411  -0.0848   0.4032   1.0000
   4.000   0.8756   0.01237   0.00429  -0.0840   0.3892   1.0000
   4.250   0.8997   0.01260   0.00449  -0.0833   0.3744   1.0000
   4.500   0.9235   0.01285   0.00469  -0.0825   0.3585   1.0000
   4.750   0.9472   0.01312   0.00490  -0.0817   0.3411   1.0000
   5.000   0.9709   0.01339   0.00515  -0.0809   0.3239   1.0000
   5.250   0.9945   0.01368   0.00541  -0.0801   0.3085   1.0000
   5.500   1.0182   0.01398   0.00569  -0.0793   0.2951   1.0000
   5.750   1.0417   0.01429   0.00599  -0.0785   0.2816   1.0000
   6.000   1.0650   0.01462   0.00633  -0.0777   0.2666   1.0000
   6.250   1.0879   0.01498   0.00668  -0.0768   0.2501   1.0000
   6.500   1.1102   0.01538   0.00706  -0.0759   0.2303   1.0000
   6.750   1.1315   0.01588   0.00749  -0.0749   0.2048   1.0000
   7.000   1.1505   0.01660   0.00802  -0.0736   0.1639   1.0000
   7.250   1.1673   0.01755   0.00871  -0.0720   0.1269   1.0000
   7.500   1.1858   0.01834   0.00940  -0.0707   0.1082   1.0000
   7.750   1.2050   0.01902   0.01005  -0.0694   0.0960   1.0000
   8.000   1.2243   0.01966   0.01068  -0.0681   0.0865   1.0000
   8.250   1.2444   0.02020   0.01129  -0.0670   0.0779   1.0000
   8.500   1.2636   0.02080   0.01191  -0.0657   0.0706   1.0000
   8.750   1.2816   0.02148   0.01256  -0.0643   0.0604   1.0000
   9.000   1.2992   0.02215   0.01324  -0.0628   0.0494   1.0000
   9.250   1.3105   0.02332   0.01423  -0.0606   0.0221   1.0000
   9.500   1.3204   0.02453   0.01546  -0.0579   0.0177   1.0000
   9.750   1.3295   0.02564   0.01671  -0.0551   0.0157   1.0000
  10.000   1.3373   0.02680   0.01803  -0.0522   0.0140   1.0000
  10.250   1.3470   0.02782   0.01919  -0.0497   0.0134   1.0000
  10.500   1.3552   0.02894   0.02049  -0.0471   0.0128   1.0000
  10.750   1.3617   0.03020   0.02192  -0.0445   0.0124   1.0000
  11.000   1.3663   0.03163   0.02352  -0.0418   0.0119   1.0000
  11.250   1.3692   0.03320   0.02527  -0.0393   0.0115   1.0000
  11.500   1.3697   0.03503   0.02726  -0.0368   0.0110   1.0000
  11.750   1.3687   0.03706   0.02946  -0.0347   0.0107   1.0000
  12.000   1.3634   0.03962   0.03218  -0.0327   0.0103   1.0000
  12.250   1.3546   0.04271   0.03547  -0.0312   0.0099   1.0000
  12.500   1.3458   0.04613   0.03905  -0.0304   0.0098   1.0000
  12.750   1.3330   0.05038   0.04345  -0.0304   0.0095   1.0000
  13.000   1.3282   0.05403   0.04726  -0.0309   0.0094   1.0000
  13.250   1.3235   0.05791   0.05130  -0.0317   0.0093   1.0000
  13.500   1.3171   0.06223   0.05578  -0.0329   0.0092   1.0000
  13.750   1.3098   0.06682   0.06051  -0.0344   0.0091   1.0000
  14.000   1.3026   0.07154   0.06537  -0.0360   0.0090   1.0000
  14.250   1.2948   0.07639   0.07037  -0.0377   0.0090   1.0000
  14.500   1.2866   0.08143   0.07554  -0.0396   0.0088   1.0000
  14.750   1.2790   0.08644   0.08069  -0.0415   0.0087   1.0000
<< Back to GOE 457 AIRFOIL (goe457-il)

Polar data table (+)

Polar graphs


<< Back to GOE 457 AIRFOIL (goe457-il)