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

GOE 777 AIRFOIL (goe777-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 777 AIRFOIL (goe777-il)
Reynolds number: 100,000
Max Cl/Cd: 15.47 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe777-il-100000.txt
Download as CSV file: xf-goe777-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 777 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.1795   0.10894   0.10339  -0.0604   0.7063   0.1886
 -10.750  -0.4347   0.06315   0.05675  -0.0895   0.7158   0.1087
 -10.500  -0.4358   0.05959   0.05300  -0.0890   0.7125   0.1082
 -10.250  -0.4423   0.05607   0.04925  -0.0881   0.7091   0.1080
 -10.000  -0.4477   0.05284   0.04568  -0.0867   0.7056   0.1081
  -9.750  -0.4497   0.04993   0.04236  -0.0850   0.7023   0.1083
  -9.500  -0.4447   0.04732   0.03936  -0.0834   0.6991   0.1088
  -9.250  -0.4238   0.04512   0.03715  -0.0831   0.6956   0.1102
  -9.000  -0.4047   0.04368   0.03561  -0.0824   0.6924   0.1123
  -8.750  -0.3894   0.04221   0.03395  -0.0815   0.6890   0.1147
  -8.500  -0.3759   0.04067   0.03209  -0.0802   0.6857   0.1170
  -8.250  -0.3622   0.03937   0.03032  -0.0787   0.6826   0.1192
  -8.000  -0.3393   0.03782   0.02880  -0.0785   0.6796   0.1219
  -7.750  -0.3162   0.03697   0.02792  -0.0781   0.6768   0.1258
  -7.500  -0.2963   0.03612   0.02672  -0.0769   0.6740   0.1300
  -7.250  -0.2725   0.03497   0.02550  -0.0764   0.6713   0.1340
  -7.000  -0.2492   0.03445   0.02495  -0.0759   0.6685   0.1390
  -6.750  -0.2288   0.03416   0.02447  -0.0753   0.6653   0.1441
  -6.500  -0.2049   0.03358   0.02411  -0.0754   0.6624   0.1495
  -6.250  -0.1830   0.03339   0.02383  -0.0748   0.6591   0.1558
  -6.000  -0.1590   0.03285   0.02337  -0.0744   0.6558   0.1621
  -5.750  -0.1347   0.03255   0.02303  -0.0738   0.6529   0.1697
  -5.500  -0.1096   0.03209   0.02261  -0.0732   0.6504   0.1773
  -5.250  -0.0849   0.03194   0.02240  -0.0725   0.6479   0.1863
  -5.000  -0.0711   0.03236   0.02306  -0.0716   0.6428   0.1935
  -4.750  -0.0542   0.03263   0.02339  -0.0705   0.6373   0.2030
  -4.500  -0.0323   0.03228   0.02317  -0.0694   0.6328   0.2151
  -4.250  -0.0068   0.03167   0.02264  -0.0683   0.6294   0.2332
  -4.000   0.0176   0.03103   0.02220  -0.0670   0.6266   0.2664
  -3.750   0.0175   0.03204   0.02369  -0.0647   0.6202   0.3133
  -3.500   0.0200   0.03223   0.02458  -0.0618   0.6147   0.4220
  -3.250   0.0343   0.03220   0.02512  -0.0589   0.6107   0.5386
  -3.000   0.0597   0.03213   0.02510  -0.0569   0.6077   0.6068
  -2.750   0.0898   0.03196   0.02488  -0.0552   0.6054   0.6527
  -2.500   0.0693   0.03524   0.02840  -0.0521   0.5946   0.6699
  -2.250   0.0895   0.03558   0.02877  -0.0500   0.5897   0.7017
  -2.000   0.1196   0.03525   0.02836  -0.0483   0.5868   0.7346
  -1.750   0.1534   0.03478   0.02784  -0.0466   0.5847   0.7646
  -1.500   0.1159   0.03908   0.03241  -0.0426   0.5704   0.7777
  -1.250   0.1431   0.03884   0.03214  -0.0400   0.5670   0.8092
  -1.000   0.1777   0.03824   0.03147  -0.0378   0.5649   0.8392
  -0.750   0.1364   0.04254   0.03598  -0.0335   0.5490   0.8561
  -0.500   0.1764   0.04195   0.03530  -0.0324   0.5463   0.8854
  -0.250   0.2269   0.04120   0.03440  -0.0328   0.5444   0.9121
   0.000   0.2617   0.04289   0.03605  -0.0350   0.5373   0.9328
   0.250   0.3754   0.04061   0.03341  -0.0436   0.5394   0.9440
   0.500   0.3808   0.04436   0.03728  -0.0461   0.5248   0.9611
   0.750   0.4066   0.04727   0.04020  -0.0504   0.5141   0.9754
   1.000   0.5035   0.04391   0.03654  -0.0565   0.5172   0.9787
   1.250   0.6037   0.04054   0.03279  -0.0638   0.5202   0.9808
   1.500   0.6071   0.04531   0.03780  -0.0679   0.5041   0.9986
   1.750   0.6429   0.04417   0.03651  -0.0684   0.5026   1.0000
   2.000   0.6678   0.04316   0.03535  -0.0673   0.5013   1.0000
   2.250   0.3108   0.06254   0.05515  -0.0387   0.4677   1.0000
   2.500   0.3415   0.06260   0.05507  -0.0386   0.4654   1.0000
   2.750   0.3783   0.06222   0.05455  -0.0386   0.4638   1.0000
   3.000   0.3025   0.07180   0.06426  -0.0373   0.4561   1.0000
   3.250   0.2979   0.07517   0.06759  -0.0370   0.4533   1.0000
   3.500   0.3068   0.07740   0.06976  -0.0369   0.4507   1.0000
   3.750   0.2755   0.08508   0.07754  -0.0381   0.4668   1.0000
   4.000   0.2313   0.09327   0.08586  -0.0393   0.4953   1.0000
   4.250   0.2391   0.09481   0.08735  -0.0389   0.4893   1.0000
   4.500   0.2580   0.09662   0.08909  -0.0391   0.4863   1.0000
   4.750   0.2812   0.09864   0.09103  -0.0394   0.4843   1.0000
   5.000   0.3097   0.10098   0.09329  -0.0401   0.4828   1.0000
   5.250   0.3431   0.10382   0.09605  -0.0411   0.4817   1.0000
   5.500   0.3719   0.10718   0.09934  -0.0420   0.4810   1.0000
   5.750   0.3035   0.10720   0.09945  -0.0388   0.4713   1.0000
   6.000   0.3209   0.10907   0.10127  -0.0389   0.4680   1.0000
   6.250   0.3428   0.11117   0.10332  -0.0393   0.4658   1.0000
   6.500   0.3694   0.11369   0.10577  -0.0399   0.4643   1.0000
   6.750   0.4024   0.11689   0.10892  -0.0409   0.4632   1.0000
   7.000   0.3604   0.11792   0.10999  -0.0392   0.4568   1.0000
   7.250   0.3714   0.11969   0.11174  -0.0393   0.4520   1.0000
<< Back to GOE 777 AIRFOIL (goe777-il)

Polar data table (+)

Polar graphs


<< Back to GOE 777 AIRFOIL (goe777-il)