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

GOE 426 AIRFOIL (goe426-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 426 AIRFOIL (goe426-il)
Reynolds number: 200,000
Max Cl/Cd: 77.87 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe426-il-200000-n5.txt
Download as CSV file: xf-goe426-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 426 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.1341   0.08841   0.08490  -0.0648   0.9398   0.0463
  -9.750  -0.1264   0.08417   0.08063  -0.0679   0.9277   0.0483
  -9.500  -0.1442   0.07723   0.07368  -0.0753   0.9135   0.0506
  -9.000  -0.2042   0.08258   0.07894  -0.0719   0.9440   0.0497
  -8.500  -0.2239   0.06043   0.05664  -0.0931   0.8993   0.0357
  -8.250  -0.2304   0.05460   0.05067  -0.0991   0.8750   0.0354
  -8.000  -0.2374   0.04936   0.04518  -0.1024   0.8531   0.0352
  -7.750  -0.2414   0.04442   0.03991  -0.1042   0.8340   0.0352
  -7.500  -0.2434   0.03925   0.03421  -0.1052   0.8166   0.0355
  -7.250  -0.2349   0.03571   0.03020  -0.1052   0.8013   0.0358
  -7.000  -0.2208   0.03298   0.02708  -0.1050   0.7876   0.0358
  -6.750  -0.2041   0.03059   0.02428  -0.1047   0.7752   0.0358
  -6.500  -0.1852   0.02856   0.02182  -0.1042   0.7637   0.0360
  -6.250  -0.1646   0.02676   0.01968  -0.1038   0.7522   0.0361
  -6.000  -0.1428   0.02512   0.01778  -0.1035   0.7412   0.0364
  -5.750  -0.1197   0.02382   0.01624  -0.1032   0.7308   0.0367
  -5.500  -0.0957   0.02279   0.01503  -0.1029   0.7201   0.0372
  -5.250  -0.0709   0.02200   0.01410  -0.1026   0.7103   0.0379
  -5.000  -0.0459   0.02121   0.01313  -0.1023   0.7009   0.0386
  -4.750  -0.0205   0.02036   0.01208  -0.1020   0.6915   0.0392
  -4.500   0.0051   0.01953   0.01103  -0.1016   0.6825   0.0395
  -4.250   0.0310   0.01878   0.01012  -0.1013   0.6726   0.0399
  -4.000   0.0570   0.01813   0.00929  -0.1009   0.6638   0.0404
  -3.750   0.0833   0.01753   0.00858  -0.1006   0.6547   0.0409
  -3.500   0.1096   0.01701   0.00794  -0.1003   0.6465   0.0415
  -3.250   0.1359   0.01654   0.00737  -0.1000   0.6376   0.0420
  -3.000   0.1623   0.01619   0.00691  -0.0996   0.6296   0.0427
  -2.750   0.1879   0.01563   0.00637  -0.0993   0.6215   0.0440
  -2.500   0.2140   0.01528   0.00598  -0.0991   0.6145   0.0453
  -2.250   0.2404   0.01497   0.00564  -0.0988   0.6068   0.0465
  -2.000   0.2667   0.01469   0.00529  -0.0985   0.6004   0.0476
  -1.750   0.2933   0.01444   0.00501  -0.0983   0.5935   0.0489
  -1.500   0.3200   0.01426   0.00474  -0.0981   0.5873   0.0502
  -1.250   0.3467   0.01404   0.00447  -0.0979   0.5815   0.0520
  -1.000   0.3736   0.01387   0.00427  -0.0978   0.5753   0.0550
  -0.750   0.4006   0.01378   0.00409  -0.0976   0.5700   0.0594
  -0.500   0.4276   0.01364   0.00396  -0.0974   0.5646   0.0668
  -0.250   0.4545   0.01344   0.00388  -0.0973   0.5591   0.1013
   0.000   0.4778   0.01249   0.00386  -0.0972   0.5543   0.3931
   0.250   0.5017   0.01212   0.00399  -0.0966   0.5496   0.5462
   0.500   0.5261   0.01196   0.00411  -0.0957   0.5445   0.6396
   0.750   0.5495   0.01184   0.00421  -0.0945   0.5399   0.7200
   1.000   0.5730   0.01177   0.00426  -0.0933   0.5359   0.7929
   1.250   0.6042   0.01154   0.00431  -0.0933   0.5310   0.9082
   1.500   0.6472   0.01160   0.00431  -0.0966   0.5258   1.0000
   1.750   0.6740   0.01176   0.00435  -0.0964   0.5217   1.0000
   2.000   0.7009   0.01192   0.00443  -0.0963   0.5176   1.0000
   2.250   0.7276   0.01206   0.00453  -0.0962   0.5128   1.0000
   2.500   0.7544   0.01221   0.00463  -0.0960   0.5084   1.0000
   2.750   0.7812   0.01239   0.00473  -0.0959   0.5045   1.0000
   3.000   0.8079   0.01257   0.00486  -0.0957   0.5003   1.0000
   3.250   0.8344   0.01272   0.00501  -0.0956   0.4955   1.0000
   3.500   0.8609   0.01289   0.00516  -0.0954   0.4910   1.0000
   3.750   0.8876   0.01309   0.00529  -0.0952   0.4873   1.0000
   4.000   0.9138   0.01327   0.00549  -0.0950   0.4827   1.0000
   4.250   0.9399   0.01344   0.00568  -0.0948   0.4778   1.0000
   4.500   0.9660   0.01363   0.00585  -0.0945   0.4733   1.0000
   4.750   0.9923   0.01384   0.00603  -0.0943   0.4693   1.0000
   5.000   1.0179   0.01403   0.00628  -0.0940   0.4644   1.0000
   5.250   1.0435   0.01422   0.00650  -0.0937   0.4595   1.0000
   5.500   1.0692   0.01444   0.00669  -0.0934   0.4551   1.0000
   5.750   1.0944   0.01465   0.00696  -0.0930   0.4500   1.0000
   6.000   1.1195   0.01486   0.00723  -0.0927   0.4449   1.0000
   6.250   1.1446   0.01508   0.00747  -0.0923   0.4404   1.0000
   6.500   1.1694   0.01532   0.00775  -0.0919   0.4356   1.0000
   6.750   1.1936   0.01555   0.00806  -0.0914   0.4298   1.0000
   7.000   1.2179   0.01579   0.00832  -0.0908   0.4248   1.0000
   7.250   1.2417   0.01604   0.00864  -0.0903   0.4193   1.0000
   7.500   1.2649   0.01629   0.00897  -0.0896   0.4129   1.0000
   7.750   1.2877   0.01655   0.00922  -0.0889   0.4068   1.0000
   8.000   1.3098   0.01682   0.00962  -0.0880   0.3995   1.0000
   8.250   1.3312   0.01710   0.00990  -0.0871   0.3924   1.0000
   8.500   1.3512   0.01739   0.01027  -0.0859   0.3827   1.0000
   8.750   1.3694   0.01771   0.01061  -0.0844   0.3717   1.0000
   9.000   1.3855   0.01808   0.01099  -0.0827   0.3593   1.0000
   9.250   1.3996   0.01852   0.01141  -0.0806   0.3457   1.0000
   9.500   1.4097   0.01902   0.01190  -0.0779   0.3296   1.0000
   9.750   1.4176   0.01963   0.01248  -0.0749   0.3134   1.0000
  10.000   1.4242   0.02035   0.01318  -0.0720   0.2959   1.0000
  10.250   1.4290   0.02122   0.01403  -0.0690   0.2766   1.0000
  10.500   1.4321   0.02228   0.01504  -0.0661   0.2570   1.0000
  10.750   1.4335   0.02353   0.01625  -0.0632   0.2350   1.0000
  11.000   1.4291   0.02525   0.01785  -0.0601   0.2083   1.0000
  11.250   1.4198   0.02750   0.01996  -0.0571   0.1785   1.0000
  11.500   1.4072   0.03027   0.02259  -0.0544   0.1505   1.0000
  12.000   1.3858   0.03636   0.02853  -0.0508   0.1128   1.0000
  12.250   1.3778   0.03947   0.03163  -0.0497   0.0977   1.0000
  12.500   1.3708   0.04264   0.03481  -0.0488   0.0826   1.0000
  12.750   1.3599   0.04635   0.03847  -0.0481   0.0669   1.0000
  13.000   1.3503   0.05011   0.04219  -0.0477   0.0558   1.0000
  13.250   1.3420   0.05388   0.04597  -0.0476   0.0470   1.0000
  13.500   1.3356   0.05757   0.04969  -0.0476   0.0402   1.0000
  13.750   1.3304   0.06122   0.05339  -0.0477   0.0350   1.0000
  14.000   1.3259   0.06491   0.05713  -0.0480   0.0317   1.0000
  14.250   1.3219   0.06860   0.06089  -0.0483   0.0290   1.0000
  14.500   1.3179   0.07239   0.06474  -0.0488   0.0272   1.0000
  14.750   1.3161   0.07597   0.06842  -0.0493   0.0258   1.0000
  15.000   1.3131   0.07976   0.07231  -0.0500   0.0245   1.0000
  15.250   1.3096   0.08371   0.07633  -0.0508   0.0234   1.0000
  15.500   1.3053   0.08785   0.08055  -0.0517   0.0226   1.0000
<< Back to GOE 426 AIRFOIL (goe426-il)

Polar data table (+)

Polar graphs


<< Back to GOE 426 AIRFOIL (goe426-il)