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

GOE 647 AIRFOIL (goe647-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 647 AIRFOIL (goe647-il)
Reynolds number: 50,000
Max Cl/Cd: 28.3 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe647-il-50000-n5.txt
Download as CSV file: xf-goe647-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 647 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.2622   0.11632   0.10957  -0.0434   1.0000   0.0866
 -10.250  -0.2689   0.11416   0.10751  -0.0420   1.0000   0.0865
 -10.000  -0.2766   0.11220   0.10564  -0.0403   1.0000   0.0862
  -9.750  -0.2861   0.11038   0.10392  -0.0385   1.0000   0.0859
  -9.500  -0.2960   0.10853   0.10217  -0.0368   0.9998   0.0854
  -9.250  -0.2803   0.10369   0.09732  -0.0419   0.9918   0.0845
  -9.000  -0.2686   0.09867   0.09230  -0.0472   0.9831   0.0837
  -8.750  -0.2607   0.09358   0.08721  -0.0524   0.9731   0.0832
  -8.500  -0.2577   0.08821   0.08184  -0.0580   0.9619   0.0834
  -8.250  -0.2595   0.08251   0.07614  -0.0639   0.9496   0.0835
  -8.000  -0.2681   0.07552   0.06911  -0.0712   0.9356   0.0834
  -7.750  -0.2825   0.06598   0.05936  -0.0805   0.9210   0.0830
  -7.500  -0.2916   0.05885   0.05187  -0.0858   0.9076   0.0829
  -7.250  -0.2935   0.05387   0.04648  -0.0877   0.8941   0.0832
  -7.000  -0.2824   0.04943   0.04154  -0.0899   0.8842   0.0839
  -6.750  -0.2681   0.04588   0.03745  -0.0909   0.8738   0.0852
  -6.500  -0.2506   0.04313   0.03417  -0.0913   0.8639   0.0871
  -6.250  -0.2220   0.04158   0.03257  -0.0922   0.8555   0.0898
  -6.000  -0.1975   0.03991   0.03066  -0.0924   0.8464   0.0923
  -5.750  -0.1717   0.03813   0.02854  -0.0926   0.8372   0.0948
  -5.500  -0.1454   0.03651   0.02650  -0.0927   0.8284   0.0977
  -5.250  -0.1187   0.03533   0.02529  -0.0926   0.8189   0.1015
  -5.000  -0.0907   0.03425   0.02407  -0.0926   0.8100   0.1064
  -4.750  -0.0631   0.03311   0.02264  -0.0924   0.8002   0.1119
  -4.500  -0.0367   0.03217   0.02177  -0.0920   0.7905   0.1178
  -4.250  -0.0074   0.03119   0.02065  -0.0919   0.7809   0.1272
  -4.000   0.0169   0.03048   0.01995  -0.0911   0.7697   0.1380
  -3.750   0.0503   0.02956   0.01906  -0.0916   0.7609   0.1557
  -3.500   0.0721   0.02916   0.01875  -0.0904   0.7483   0.1762
  -3.250   0.1083   0.02839   0.01799  -0.0912   0.7401   0.2144
  -3.000   0.1288   0.02816   0.01782  -0.0897   0.7271   0.2496
  -2.750   0.1532   0.02790   0.01758  -0.0887   0.7158   0.2881
  -2.500   0.1831   0.02754   0.01719  -0.0884   0.7062   0.3266
  -2.250   0.2041   0.02750   0.01709  -0.0869   0.6937   0.3562
  -2.000   0.2345   0.02716   0.01671  -0.0866   0.6842   0.3846
  -1.750   0.2586   0.02705   0.01651  -0.0855   0.6723   0.4113
  -1.500   0.2820   0.02693   0.01641  -0.0843   0.6606   0.4408
  -1.250   0.3136   0.02654   0.01593  -0.0841   0.6511   0.4735
  -1.000   0.3351   0.02645   0.01586  -0.0827   0.6382   0.4993
  -0.750   0.3633   0.02615   0.01551  -0.0823   0.6273   0.5272
  -0.500   0.3909   0.02582   0.01520  -0.0817   0.6161   0.5580
  -0.250   0.4128   0.02567   0.01518  -0.0803   0.6037   0.5955
   0.000   0.4418   0.02522   0.01487  -0.0796   0.5936   0.6568
   0.250   0.4817   0.02478   0.01488  -0.0808   0.5803   0.7965
   0.500   0.5471   0.02482   0.01470  -0.0879   0.5651   1.0000
   0.750   0.5705   0.02504   0.01463  -0.0871   0.5541   1.0000
   1.000   0.5936   0.02528   0.01461  -0.0862   0.5430   1.0000
   1.250   0.6126   0.02568   0.01482  -0.0847   0.5313   1.0000
   1.500   0.6395   0.02584   0.01469  -0.0842   0.5218   1.0000
   1.750   0.6565   0.02633   0.01506  -0.0825   0.5103   1.0000
   2.000   0.6798   0.02666   0.01518  -0.0817   0.5009   1.0000
   2.250   0.7005   0.02708   0.01544  -0.0804   0.4910   1.0000
   2.500   0.7219   0.02751   0.01571  -0.0794   0.4820   1.0000
   2.750   0.7434   0.02793   0.01599  -0.0783   0.4732   1.0000
   3.000   0.7650   0.02841   0.01632  -0.0773   0.4651   1.0000
   3.250   0.7848   0.02895   0.01676  -0.0761   0.4569   1.0000
   3.500   0.8101   0.02932   0.01696  -0.0756   0.4502   1.0000
   3.750   0.8251   0.03007   0.01769  -0.0738   0.4420   1.0000
   4.000   0.8512   0.03044   0.01789  -0.0734   0.4359   1.0000
   4.250   0.8669   0.03124   0.01867  -0.0719   0.4290   1.0000
   4.500   0.8857   0.03193   0.01930  -0.0707   0.4226   1.0000
   4.750   0.9137   0.03229   0.01950  -0.0706   0.4177   1.0000
   5.000   0.9241   0.03335   0.02061  -0.0684   0.4113   1.0000
   5.250   0.9410   0.03414   0.02138  -0.0671   0.4055   1.0000
   5.500   0.9679   0.03457   0.02169  -0.0669   0.4010   1.0000
   5.750   0.9793   0.03567   0.02282  -0.0650   0.3958   1.0000
   6.000   0.9880   0.03689   0.02409  -0.0628   0.3906   1.0000
   6.250   1.0075   0.03768   0.02486  -0.0619   0.3864   1.0000
   6.500   1.0372   0.03810   0.02517  -0.0622   0.3829   1.0000
   6.750   1.0318   0.03989   0.02710  -0.0585   0.3781   1.0000
   7.000   1.0282   0.04165   0.02896  -0.0552   0.3733   1.0000
   7.250   1.0430   0.04268   0.02998  -0.0539   0.3693   1.0000
   7.500   1.0727   0.04308   0.03031  -0.0541   0.3662   1.0000
   7.750   1.0625   0.04544   0.03278  -0.0507   0.3622   1.0000
   8.000   1.0091   0.05090   0.03851  -0.0453   0.3564   1.0000
   8.250   1.0022   0.05387   0.04154  -0.0437   0.3522   1.0000
   8.500   1.0250   0.05455   0.04220  -0.0432   0.3496   1.0000
   8.750   1.0659   0.05391   0.04150  -0.0435   0.3477   1.0000
   9.000   0.8918   0.07487   0.06290  -0.0424   0.3321   1.0000
   9.250   0.9148   0.07532   0.06333  -0.0417   0.3305   1.0000
  10.750   0.8023   0.11390   0.10232  -0.0499   0.2994   1.0000
  11.000   0.8208   0.11517   0.10360  -0.0496   0.2973   1.0000
<< Back to GOE 647 AIRFOIL (goe647-il)

Polar data table (+)

Polar graphs


<< Back to GOE 647 AIRFOIL (goe647-il)