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

GOE 386 AIRFOIL (goe386-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 386 AIRFOIL (goe386-il)
Reynolds number: 100,000
Max Cl/Cd: 33.24 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe386-il-100000.txt
Download as CSV file: xf-goe386-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 386 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.0225   0.10521   0.10016  -0.0941   0.9218   0.1880
  -9.250   0.0082   0.09982   0.09474  -0.0956   0.9118   0.1895
  -9.000   0.0382   0.09637   0.09125  -0.0951   0.8967   0.1914
  -8.750   0.0570   0.09366   0.08850  -0.0950   0.8815   0.1945
  -8.500   0.0635   0.09119   0.08597  -0.0955   0.8678   0.2000
  -8.250   0.0214   0.08820   0.08298  -0.1002   0.8505   0.2058
  -8.000   0.0537   0.08549   0.08022  -0.0975   0.8377   0.2078
  -7.750  -0.1645   0.05156   0.04531  -0.1207   0.8230   0.1344
  -7.500  -0.1215   0.05296   0.04699  -0.1196   0.8130   0.1375
  -7.250  -0.1593   0.04433   0.03746  -0.1191   0.8047   0.1333
  -7.000  -0.1594   0.04148   0.03421  -0.1172   0.7939   0.1337
  -6.750  -0.1473   0.03849   0.03067  -0.1163   0.7880   0.1345
  -6.500  -0.1344   0.03689   0.02884  -0.1149   0.7795   0.1355
  -6.250  -0.1132   0.03570   0.02756  -0.1142   0.7724   0.1368
  -6.000  -0.0858   0.03458   0.02633  -0.1142   0.7675   0.1389
  -5.750  -0.0671   0.03380   0.02540  -0.1131   0.7604   0.1412
  -5.500  -0.0485   0.03281   0.02415  -0.1120   0.7532   0.1437
  -5.250  -0.0235   0.03152   0.02243  -0.1116   0.7484   0.1462
  -5.000   0.0075   0.03053   0.02140  -0.1120   0.7448   0.1489
  -4.750   0.0210   0.03065   0.02158  -0.1102   0.7372   0.1516
  -4.500   0.0443   0.03025   0.02106  -0.1095   0.7319   0.1556
  -4.250   0.0729   0.02948   0.02005  -0.1095   0.7277   0.1600
  -4.000   0.1052   0.02883   0.01943  -0.1099   0.7243   0.1646
  -3.750   0.1169   0.02919   0.01981  -0.1078   0.7171   0.1689
  -3.500   0.1387   0.02904   0.01960  -0.1070   0.7119   0.1750
  -3.250   0.1689   0.02856   0.01916  -0.1070   0.7074   0.1833
  -3.000   0.2044   0.02774   0.01828  -0.1077   0.7035   0.1945
  -2.750   0.2142   0.02814   0.01872  -0.1051   0.6944   0.2058
  -2.500   0.2429   0.02749   0.01819  -0.1048   0.6881   0.2272
  -2.250   0.2783   0.02675   0.01740  -0.1052   0.6834   0.2622
  -2.000   0.2906   0.02725   0.01805  -0.1031   0.6759   0.2823
  -1.750   0.3120   0.02740   0.01828  -0.1021   0.6699   0.3026
  -1.500   0.3426   0.02715   0.01805  -0.1021   0.6654   0.3226
  -1.250   0.3776   0.02682   0.01768  -0.1028   0.6616   0.3430
  -1.000   0.3841   0.02773   0.01873  -0.1000   0.6538   0.3553
  -0.750   0.4083   0.02783   0.01884  -0.0993   0.6478   0.3736
  -0.500   0.4417   0.02741   0.01846  -0.0997   0.6433   0.3950
  -0.250   0.4747   0.02713   0.01821  -0.1001   0.6391   0.4199
   0.000   0.4775   0.02809   0.01945  -0.0968   0.6302   0.4396
   0.250   0.5055   0.02775   0.01932  -0.0964   0.6246   0.4826
   0.500   0.5398   0.02689   0.01885  -0.0965   0.6204   0.5748
   0.750   0.5454   0.02714   0.01986  -0.0921   0.6130   0.7335
   1.000   0.6348   0.02694   0.01991  -0.1010   0.6046   0.9774
   1.250   0.7158   0.02647   0.01908  -0.1102   0.5989   1.0000
   1.500   0.7165   0.02728   0.01984  -0.1061   0.5912   1.0000
   1.750   0.7272   0.02768   0.02015  -0.1034   0.5833   1.0000
   2.000   0.7623   0.02738   0.01961  -0.1040   0.5778   1.0000
   2.250   0.7672   0.02811   0.02031  -0.1004   0.5694   1.0000
   2.500   0.7826   0.02840   0.02051  -0.0983   0.5611   1.0000
   2.750   0.8223   0.02801   0.01989  -0.0996   0.5558   1.0000
   3.000   0.8207   0.02902   0.02093  -0.0951   0.5460   1.0000
   3.250   0.8459   0.02906   0.02085  -0.0943   0.5386   1.0000
   3.500   0.8898   0.02861   0.02015  -0.0963   0.5337   1.0000
   3.750   0.8805   0.02985   0.02151  -0.0908   0.5224   1.0000
   4.000   0.9161   0.02953   0.02101  -0.0915   0.5162   1.0000
   4.250   0.9381   0.02985   0.02124  -0.0905   0.5091   1.0000
   4.500   0.9448   0.03060   0.02202  -0.0874   0.4999   1.0000
   4.750   0.9844   0.03024   0.02146  -0.0887   0.4950   1.0000
   5.000   0.9878   0.03134   0.02261  -0.0853   0.4868   1.0000
   5.250   1.0031   0.03187   0.02311  -0.0835   0.4796   1.0000
   5.500   1.0448   0.03143   0.02248  -0.0852   0.4752   1.0000
   5.750   1.0421   0.03280   0.02394  -0.0811   0.4676   1.0000
   6.000   1.0532   0.03353   0.02468  -0.0788   0.4610   1.0000
   6.250   1.0934   0.03317   0.02416  -0.0803   0.4569   1.0000
   6.500   1.1120   0.03387   0.02483  -0.0792   0.4521   1.0000
   6.750   1.0755   0.03663   0.02780  -0.0712   0.4444   1.0000
   7.000   1.1079   0.03660   0.02769  -0.0717   0.4404   1.0000
   7.250   1.1632   0.03579   0.02670  -0.0752   0.4373   1.0000
   7.500   1.0152   0.04566   0.03707  -0.0584   0.4252   1.0000
   7.750   1.0597   0.04445   0.03577  -0.0593   0.4228   1.0000
   8.000   1.1188   0.04255   0.03374  -0.0616   0.4210   1.0000
   8.250   1.1829   0.04084   0.03186  -0.0649   0.4192   1.0000
  10.250   0.5788   0.14086   0.13333  -0.0651   0.4090   1.0000
  10.500   0.6005   0.14277   0.13523  -0.0652   0.4041   1.0000
<< Back to GOE 386 AIRFOIL (goe386-il)

Polar data table (+)

Polar graphs


<< Back to GOE 386 AIRFOIL (goe386-il)