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

GOE 405 AIRFOIL (goe405-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 405 AIRFOIL (goe405-il)
Reynolds number: 200,000
Max Cl/Cd: 82.26 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe405-il-200000.txt
Download as CSV file: xf-goe405-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 405 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.1275   0.09853   0.09521  -0.0739   0.9657   0.0478
  -8.500  -0.1179   0.09531   0.09200  -0.0845   0.9588   0.0506
  -8.000  -0.0839   0.08674   0.08343  -0.0883   0.9495   0.0533
  -7.750  -0.0614   0.08336   0.08003  -0.0917   0.9459   0.0559
  -7.500  -0.0390   0.07931   0.07596  -0.0987   0.9427   0.0596
  -7.250   0.0617   0.05723   0.05398  -0.1091   0.9208   0.0667
  -7.000   0.0723   0.05392   0.05065  -0.1123   0.9147   0.0714
  -6.500   0.0299   0.06464   0.06120  -0.1181   0.9149   0.0708
  -6.250   0.0428   0.05892   0.05532  -0.1301   0.9021   0.0756
  -6.000   0.0662   0.05648   0.05290  -0.1292   0.8996   0.0775
  -5.750   0.0805   0.05447   0.05086  -0.1290   0.8915   0.0804
  -5.500   0.1085   0.04867   0.04471  -0.1392   0.8842   0.0889
  -5.250   0.1321   0.03115   0.02753  -0.1270   0.8579   0.0928
  -5.000   0.1361   0.03245   0.02766  -0.1440   0.8691   0.0643
  -4.750   0.1468   0.02718   0.02179  -0.1427   0.8596   0.0598
  -4.500   0.1743   0.02356   0.01745  -0.1434   0.8548   0.0611
  -4.250   0.1958   0.02190   0.01531  -0.1422   0.8468   0.0620
  -4.000   0.2234   0.01976   0.01275  -0.1423   0.8414   0.0633
  -3.750   0.2538   0.01860   0.01147  -0.1428   0.8371   0.0665
  -3.500   0.2761   0.01800   0.01074  -0.1416   0.8293   0.0694
  -3.250   0.3079   0.01707   0.00958  -0.1420   0.8249   0.0723
  -3.000   0.3324   0.01654   0.00890  -0.1411   0.8180   0.0749
  -2.750   0.3601   0.01583   0.00821  -0.1410   0.8122   0.0811
  -2.500   0.3908   0.01530   0.00755  -0.1413   0.8074   0.0886
  -2.250   0.4136   0.01496   0.00725  -0.1402   0.7998   0.1002
  -2.000   0.4448   0.01456   0.00690  -0.1405   0.7949   0.1323
  -1.750   0.4689   0.01476   0.00705  -0.1397   0.7873   0.1601
  -1.500   0.4983   0.01471   0.00695  -0.1398   0.7814   0.1779
  -1.250   0.5246   0.01474   0.00691  -0.1394   0.7745   0.1917
  -1.000   0.5519   0.01473   0.00684  -0.1392   0.7677   0.2062
  -0.750   0.5797   0.01471   0.00680  -0.1391   0.7614   0.2229
  -0.500   0.6051   0.01464   0.00672  -0.1386   0.7537   0.2359
  -0.250   0.6342   0.01441   0.00648  -0.1387   0.7476   0.2464
   0.000   0.6580   0.01429   0.00637  -0.1378   0.7393   0.2572
   0.250   0.6874   0.01412   0.00615  -0.1380   0.7330   0.2693
   0.500   0.7106   0.01405   0.00609  -0.1370   0.7241   0.2819
   0.750   0.7384   0.01387   0.00593  -0.1369   0.7172   0.2975
   1.000   0.7615   0.01372   0.00587  -0.1360   0.7083   0.3146
   1.250   0.7866   0.01355   0.00578  -0.1354   0.7004   0.3394
   1.500   0.8447   0.01208   0.00574  -0.1418   0.6915   1.0000
   1.750   0.8679   0.01221   0.00575  -0.1407   0.6822   1.0000
   2.000   0.8951   0.01228   0.00566  -0.1404   0.6739   1.0000
   2.250   0.9159   0.01241   0.00574  -0.1389   0.6632   1.0000
   2.500   0.9394   0.01252   0.00577  -0.1380   0.6533   1.0000
   2.750   0.9650   0.01260   0.00574  -0.1374   0.6437   1.0000
   3.000   0.9857   0.01272   0.00583  -0.1359   0.6320   1.0000
   3.250   1.0078   0.01285   0.00590  -0.1347   0.6207   1.0000
   3.500   1.0309   0.01297   0.00595  -0.1337   0.6096   1.0000
   3.750   1.0545   0.01311   0.00600  -0.1328   0.5987   1.0000
   4.000   1.0750   0.01328   0.00615  -0.1314   0.5865   1.0000
   4.250   1.0965   0.01347   0.00631  -0.1301   0.5750   1.0000
   4.500   1.1189   0.01367   0.00645  -0.1291   0.5641   1.0000
   4.750   1.1422   0.01389   0.00660  -0.1282   0.5537   1.0000
   5.000   1.1623   0.01413   0.00684  -0.1268   0.5423   1.0000
   5.250   1.1836   0.01439   0.00708  -0.1256   0.5317   1.0000
   5.500   1.2061   0.01467   0.00731  -0.1247   0.5219   1.0000
   5.750   1.2268   0.01495   0.00758  -0.1234   0.5116   1.0000
   6.000   1.2475   0.01526   0.00790  -0.1222   0.5017   1.0000
   6.250   1.2704   0.01560   0.00818  -0.1214   0.4929   1.0000
   6.500   1.2895   0.01592   0.00855  -0.1199   0.4831   1.0000
   6.750   1.3102   0.01628   0.00892  -0.1187   0.4739   1.0000
   7.000   1.3318   0.01665   0.00922  -0.1177   0.4640   1.0000
   7.250   1.3462   0.01697   0.00960  -0.1153   0.4525   1.0000
   7.500   1.3606   0.01732   0.00998  -0.1130   0.4402   1.0000
   7.750   1.3741   0.01768   0.01034  -0.1105   0.4282   1.0000
   8.000   1.3880   0.01808   0.01070  -0.1080   0.4174   1.0000
   8.250   1.3983   0.01846   0.01113  -0.1050   0.4063   1.0000
   8.500   1.4066   0.01887   0.01159  -0.1016   0.3941   1.0000
   8.750   1.4158   0.01932   0.01209  -0.0985   0.3821   1.0000
   9.000   1.4238   0.01983   0.01263  -0.0954   0.3694   1.0000
   9.250   1.4280   0.02042   0.01324  -0.0917   0.3534   1.0000
   9.500   1.4310   0.02114   0.01397  -0.0881   0.3352   1.0000
   9.750   1.4315   0.02208   0.01486  -0.0843   0.3147   1.0000
  10.000   1.4300   0.02327   0.01599  -0.0806   0.2873   1.0000
  10.250   1.4212   0.02506   0.01758  -0.0762   0.2482   1.0000
  10.500   1.4050   0.02761   0.01981  -0.0717   0.2002   1.0000
  10.750   1.3810   0.03110   0.02281  -0.0671   0.1290   1.0000
  11.000   1.3535   0.03526   0.02652  -0.0629   0.0643   1.0000
  11.250   1.3409   0.03855   0.02974  -0.0602   0.0472   1.0000
  11.500   1.3361   0.04130   0.03259  -0.0582   0.0410   1.0000
  11.750   1.3362   0.04372   0.03514  -0.0568   0.0376   1.0000
  12.000   1.3324   0.04660   0.03813  -0.0555   0.0351   1.0000
  12.250   1.3236   0.05014   0.04182  -0.0542   0.0335   1.0000
  12.500   1.3199   0.05330   0.04513  -0.0533   0.0325   1.0000
  12.750   1.3154   0.05666   0.04863  -0.0526   0.0316   1.0000
  13.000   1.3097   0.06026   0.05237  -0.0521   0.0308   1.0000
  13.250   1.3037   0.06402   0.05626  -0.0518   0.0302   1.0000
  13.500   1.2979   0.06786   0.06021  -0.0516   0.0295   1.0000
  13.750   1.2922   0.07175   0.06419  -0.0515   0.0287   1.0000
  14.000   1.2876   0.07554   0.06804  -0.0516   0.0279   1.0000
  14.250   1.2829   0.07932   0.07188  -0.0515   0.0272   1.0000
  14.500   1.2792   0.08266   0.07523  -0.0510   0.0262   1.0000
  14.750   1.2837   0.08507   0.07770  -0.0504   0.0256   1.0000
  15.000   1.2912   0.08708   0.07978  -0.0498   0.0251   1.0000
  15.250   1.3023   0.08849   0.08124  -0.0488   0.0247   1.0000
  15.500   1.3153   0.08967   0.08248  -0.0476   0.0242   1.0000
  15.750   1.3317   0.09043   0.08331  -0.0462   0.0239   1.0000
  16.000   1.3459   0.09167   0.08467  -0.0451   0.0234   1.0000
  16.250   1.3580   0.09331   0.08641  -0.0444   0.0228   1.0000
  16.500   1.3689   0.09516   0.08835  -0.0439   0.0221   1.0000
  16.750   1.3805   0.09702   0.09032  -0.0434   0.0216   1.0000
  17.000   1.3943   0.09893   0.09240  -0.0425   0.0217   1.0000
  17.250   1.3978   0.10236   0.09611  -0.0424   0.0220   1.0000
  17.500   1.3951   0.10673   0.10079  -0.0428   0.0225   1.0000
<< Back to GOE 405 AIRFOIL (goe405-il)

Polar data table (+)

Polar graphs


<< Back to GOE 405 AIRFOIL (goe405-il)