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

GOE 615 AIRFOIL (goe615-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 615 AIRFOIL (goe615-il)
Reynolds number: 100,000
Max Cl/Cd: 51.01 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe615-il-100000.txt
Download as CSV file: xf-goe615-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 615 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2825   0.12379   0.11917  -0.0332   1.0000   0.1000
  -9.750  -0.3151   0.12486   0.12037  -0.0324   1.0000   0.1007
  -9.500  -0.3444   0.12531   0.12094  -0.0315   1.0000   0.1010
  -9.250  -0.3122   0.11861   0.11424  -0.0275   1.0000   0.1027
  -9.000  -0.3090   0.11657   0.11224  -0.0248   1.0000   0.1042
  -8.750  -0.3058   0.11453   0.11023  -0.0240   0.9990   0.1066
  -8.500  -0.2880   0.11087   0.10656  -0.0301   0.9934   0.1120
  -8.250  -0.3188   0.11012   0.10588  -0.0448   0.9787   0.1152
  -8.000  -0.2540   0.10263   0.09833  -0.0403   0.9810   0.1195
  -7.750  -0.2338   0.09925   0.09494  -0.0440   0.9735   0.1247
  -7.500  -0.2674   0.09725   0.09294  -0.0626   0.9531   0.1305
  -7.250  -0.2160   0.09181   0.08753  -0.0555   0.9565   0.1333
  -7.000  -0.1943   0.08889   0.08459  -0.0563   0.9480   0.1378
  -6.500  -0.1690   0.08140   0.07710  -0.0668   0.9279   0.1503
  -6.250  -0.1684   0.07747   0.07301  -0.0802   0.9116   0.1627
  -6.000  -0.1288   0.07396   0.06960  -0.0764   0.9101   0.1665
  -5.750  -0.1148   0.07155   0.06716  -0.0778   0.8982   0.1744
  -5.500  -0.0857   0.06708   0.06262  -0.0841   0.8929   0.1837
  -5.250  -0.0786   0.06417   0.05956  -0.0892   0.8788   0.1975
  -5.000  -0.0394   0.06103   0.05646  -0.0897   0.8762   0.2040
  -4.750  -0.0297   0.05852   0.05388  -0.0913   0.8634   0.2175
  -4.500   0.0066   0.05522   0.05048  -0.0957   0.8591   0.2354
  -4.250   0.0187   0.05341   0.04855  -0.0961   0.8466   0.2518
  -4.000   0.0615   0.04252   0.03668  -0.1078   0.8412   0.1669
  -3.750   0.0738   0.03484   0.02746  -0.1068   0.8290   0.1127
  -3.500   0.1124   0.03247   0.02467  -0.1086   0.8248   0.1123
  -3.250   0.1317   0.03128   0.02320  -0.1071   0.8149   0.1116
  -3.000   0.1670   0.02961   0.02118  -0.1080   0.8092   0.1112
  -2.750   0.2096   0.02800   0.01923  -0.1098   0.8058   0.1126
  -2.500   0.2240   0.02786   0.01887  -0.1073   0.7942   0.1147
  -2.250   0.2622   0.02643   0.01740  -0.1087   0.7901   0.1184
  -2.000   0.2808   0.02619   0.01715  -0.1070   0.7811   0.1212
  -1.750   0.3129   0.02548   0.01638  -0.1073   0.7750   0.1263
  -1.500   0.3521   0.02447   0.01540  -0.1086   0.7714   0.1353
  -1.250   0.3633   0.02469   0.01566  -0.1058   0.7606   0.1422
  -1.000   0.4018   0.02367   0.01471  -0.1066   0.7555   0.1592
  -0.750   0.4213   0.02335   0.01459  -0.1048   0.7454   0.1893
  -0.500   0.4894   0.01996   0.01373  -0.1105   0.7402   0.8844
  -0.250   0.5629   0.01945   0.01281  -0.1184   0.7347   1.0000
   0.000   0.5764   0.01992   0.01313  -0.1159   0.7249   1.0000
   0.250   0.6128   0.01968   0.01263  -0.1167   0.7196   1.0000
   0.500   0.6240   0.02026   0.01312  -0.1138   0.7097   1.0000
   0.750   0.6576   0.02009   0.01276  -0.1142   0.7037   1.0000
   1.000   0.6736   0.02053   0.01311  -0.1121   0.6946   1.0000
   1.250   0.7042   0.02045   0.01287  -0.1120   0.6877   1.0000
   1.500   0.7252   0.02073   0.01306  -0.1106   0.6795   1.0000
   1.750   0.7524   0.02075   0.01296  -0.1100   0.6715   1.0000
   2.000   0.7770   0.02090   0.01301  -0.1091   0.6635   1.0000
   2.250   0.8025   0.02097   0.01299  -0.1083   0.6548   1.0000
   2.500   0.8273   0.02109   0.01303  -0.1073   0.6462   1.0000
   2.750   0.8548   0.02108   0.01292  -0.1068   0.6372   1.0000
   3.000   0.8765   0.02129   0.01308  -0.1054   0.6274   1.0000
   3.250   0.9091   0.02112   0.01278  -0.1056   0.6187   1.0000
   3.500   0.9274   0.02142   0.01307  -0.1037   0.6073   1.0000
   3.750   0.9586   0.02133   0.01285  -0.1037   0.5979   1.0000
   4.000   0.9821   0.02144   0.01291  -0.1026   0.5863   1.0000
   4.250   1.0032   0.02167   0.01310  -0.1011   0.5739   1.0000
   4.500   1.0307   0.02174   0.01306  -0.1006   0.5624   1.0000
   4.750   1.0589   0.02179   0.01299  -0.1002   0.5503   1.0000
   5.000   1.0770   0.02212   0.01332  -0.0983   0.5364   1.0000
   5.250   1.0980   0.02244   0.01359  -0.0968   0.5229   1.0000
   5.500   1.1222   0.02269   0.01376  -0.0959   0.5101   1.0000
   5.750   1.1499   0.02287   0.01379  -0.0956   0.4978   1.0000
   6.000   1.1645   0.02330   0.01426  -0.0932   0.4839   1.0000
   6.250   1.1819   0.02369   0.01465  -0.0913   0.4710   1.0000
   6.500   1.2034   0.02396   0.01486  -0.0900   0.4590   1.0000
   6.750   1.2265   0.02413   0.01492  -0.0890   0.4475   1.0000
   7.000   1.2391   0.02455   0.01540  -0.0864   0.4350   1.0000
   7.250   1.2560   0.02489   0.01573  -0.0844   0.4237   1.0000
   7.500   1.2788   0.02507   0.01578  -0.0834   0.4133   1.0000
   7.750   1.2914   0.02554   0.01630  -0.0809   0.4023   1.0000
   8.000   1.3074   0.02599   0.01673  -0.0790   0.3922   1.0000
   8.250   1.3302   0.02631   0.01694  -0.0782   0.3829   1.0000
   8.500   1.3410   0.02698   0.01769  -0.0756   0.3734   1.0000
   8.750   1.3661   0.02745   0.01803  -0.0752   0.3653   1.0000
   9.000   1.3768   0.02822   0.01893  -0.0727   0.3571   1.0000
   9.250   1.4039   0.02878   0.01934  -0.0728   0.3495   1.0000
   9.500   1.4111   0.02962   0.02034  -0.0698   0.3419   1.0000
   9.750   1.4423   0.03018   0.02071  -0.0706   0.3340   1.0000
  10.000   1.4427   0.03109   0.02186  -0.0666   0.3275   1.0000
  10.250   1.4641   0.03180   0.02254  -0.0659   0.3209   1.0000
  10.500   1.4827   0.03269   0.02348  -0.0649   0.3151   1.0000
  10.750   1.4870   0.03366   0.02460  -0.0617   0.3093   1.0000
  11.000   1.5172   0.03425   0.02509  -0.0624   0.3025   1.0000
  11.250   1.5144   0.03534   0.02637  -0.0583   0.2969   1.0000
  11.500   1.5198   0.03625   0.02740  -0.0555   0.2908   1.0000
  11.750   1.5469   0.03684   0.02789  -0.0557   0.2841   1.0000
  12.000   1.5347   0.03820   0.02952  -0.0510   0.2791   1.0000
  12.250   1.5433   0.03898   0.03033  -0.0489   0.2725   1.0000
  12.500   1.5567   0.03988   0.03126  -0.0475   0.2669   1.0000
  12.750   1.5445   0.04153   0.03315  -0.0436   0.2616   1.0000
  13.000   1.5577   0.04224   0.03386  -0.0423   0.2554   1.0000
  13.250   1.5565   0.04371   0.03547  -0.0398   0.2499   1.0000
  13.500   1.5455   0.04567   0.03763  -0.0369   0.2442   1.0000
  13.750   1.5666   0.04592   0.03775  -0.0362   0.2370   1.0000
  14.000   1.5414   0.04899   0.04115  -0.0330   0.2319   1.0000
  14.250   1.5417   0.05045   0.04265  -0.0314   0.2248   1.0000
  14.500   1.5318   0.05290   0.04523  -0.0296   0.2182   1.0000
  14.750   1.5216   0.05542   0.04787  -0.0281   0.2107   1.0000
  15.000   1.5111   0.05820   0.05073  -0.0270   0.2030   1.0000
  15.250   1.5026   0.06091   0.05349  -0.0261   0.1947   1.0000
  15.500   1.4834   0.06512   0.05783  -0.0256   0.1864   1.0000
  15.750   1.4859   0.06698   0.05956  -0.0251   0.1774   1.0000
  16.000   1.4604   0.07260   0.06542  -0.0255   0.1700   1.0000
  16.250   1.4626   0.07484   0.06755  -0.0253   0.1626   1.0000
  16.500   1.4425   0.08036   0.07327  -0.0261   0.1566   1.0000
  16.750   1.4566   0.08111   0.07381  -0.0257   0.1505   1.0000
  17.000   1.4315   0.08772   0.08072  -0.0272   0.1466   1.0000
  17.250   1.4206   0.09241   0.08553  -0.0283   0.1425   1.0000
  17.500   1.4369   0.09288   0.08588  -0.0279   0.1381   1.0000
  17.750   1.4219   0.09840   0.09160  -0.0294   0.1357   1.0000
  18.000   1.3925   0.10640   0.09987  -0.0322   0.1335   1.0000
  18.250   1.3429   0.11841   0.11217  -0.0373   0.1322   1.0000
<< Back to GOE 615 AIRFOIL (goe615-il)

Polar data table (+)

Polar graphs


<< Back to GOE 615 AIRFOIL (goe615-il)