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

GOE 517 AIRFOIL (goe517-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 517 AIRFOIL (goe517-il)
Reynolds number: 50,000
Max Cl/Cd: 40.93 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe517-il-50000-n5.txt
Download as CSV file: xf-goe517-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 517 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3183   0.09885   0.09252  -0.0285   1.0000   0.0810
  -7.250  -0.3221   0.09688   0.09067  -0.0276   1.0000   0.0823
  -7.000  -0.3256   0.09488   0.08878  -0.0273   1.0000   0.0839
  -6.750  -0.3301   0.09317   0.08718  -0.0277   1.0000   0.0856
  -6.500  -0.3355   0.09203   0.08614  -0.0293   1.0000   0.0869
  -6.250  -0.3382   0.09076   0.08493  -0.0314   1.0000   0.0875
  -6.000  -0.3371   0.08895   0.08312  -0.0331   1.0000   0.0878
  -5.750  -0.3364   0.08439   0.07869  -0.0290   1.0000   0.0888
  -5.500  -0.3351   0.08129   0.07564  -0.0265   1.0000   0.0904
  -5.250  -0.3327   0.07866   0.07305  -0.0254   1.0000   0.0926
  -5.000  -0.3282   0.07611   0.07051  -0.0256   1.0000   0.0954
  -4.750  -0.3190   0.07371   0.06806  -0.0278   1.0000   0.0995
  -4.250  -0.2484   0.06191   0.05577  -0.0414   0.9881   0.0558
  -4.000  -0.2160   0.05747   0.05117  -0.0459   0.9808   0.0518
  -3.750  -0.1761   0.05237   0.04575  -0.0520   0.9736   0.0481
  -3.500  -0.1373   0.04815   0.04119  -0.0569   0.9669   0.0482
  -3.250  -0.1023   0.04463   0.03735  -0.0604   0.9592   0.0499
  -3.000  -0.0657   0.04090   0.03316  -0.0636   0.9518   0.0502
  -2.750  -0.0257   0.03730   0.02902  -0.0668   0.9456   0.0513
  -2.500   0.0093   0.03423   0.02532  -0.0685   0.9375   0.0553
  -2.250   0.0489   0.03127   0.02165  -0.0707   0.9316   0.0584
  -2.000   0.0820   0.02953   0.01941  -0.0716   0.9230   0.0664
  -1.750   0.1199   0.02789   0.01727  -0.0732   0.9160   0.0766
  -1.500   0.1542   0.02674   0.01593  -0.0743   0.9075   0.0915
  -1.250   0.1927   0.02572   0.01454  -0.0759   0.9001   0.1176
  -1.000   0.2301   0.02502   0.01362  -0.0776   0.8918   0.1520
  -0.750   0.2667   0.02436   0.01286  -0.0792   0.8836   0.1813
  -0.500   0.3037   0.02376   0.01211  -0.0807   0.8753   0.2061
  -0.250   0.3349   0.02329   0.01166  -0.0813   0.8652   0.2364
   0.000   0.3724   0.02255   0.01112  -0.0831   0.8579   0.2952
   0.500   0.4510   0.02088   0.01048  -0.0870   0.8382   1.0000
   0.750   0.4858   0.02089   0.01021  -0.0878   0.8276   1.0000
   1.000   0.5155   0.02092   0.01004  -0.0877   0.8132   1.0000
   1.250   0.5449   0.02091   0.00985  -0.0874   0.7982   1.0000
   1.500   0.5735   0.02091   0.00971  -0.0869   0.7829   1.0000
   1.750   0.6009   0.02096   0.00964  -0.0863   0.7678   1.0000
   2.000   0.6273   0.02107   0.00967  -0.0857   0.7534   1.0000
   2.250   0.6534   0.02119   0.00973  -0.0850   0.7391   1.0000
   2.500   0.6794   0.02132   0.00980  -0.0843   0.7248   1.0000
   2.750   0.7052   0.02144   0.00991  -0.0835   0.7105   1.0000
   3.000   0.7310   0.02155   0.01000  -0.0827   0.6959   1.0000
   3.250   0.7568   0.02168   0.01012  -0.0819   0.6813   1.0000
   3.500   0.7824   0.02181   0.01028  -0.0811   0.6662   1.0000
   3.750   0.8077   0.02195   0.01044  -0.0802   0.6507   1.0000
   4.000   0.8328   0.02213   0.01064  -0.0793   0.6348   1.0000
   4.250   0.8577   0.02234   0.01091  -0.0785   0.6187   1.0000
   4.500   0.8824   0.02258   0.01120  -0.0775   0.6020   1.0000
   4.750   0.9071   0.02285   0.01153  -0.0766   0.5852   1.0000
   5.000   0.9316   0.02319   0.01196  -0.0757   0.5683   1.0000
   5.250   0.9559   0.02357   0.01243  -0.0748   0.5513   1.0000
   5.500   0.9800   0.02402   0.01297  -0.0739   0.5343   1.0000
   5.750   1.0037   0.02453   0.01363  -0.0731   0.5177   1.0000
   6.000   1.0269   0.02509   0.01434  -0.0721   0.5013   1.0000
   6.250   1.0497   0.02570   0.01513  -0.0712   0.4856   1.0000
   6.500   1.0722   0.02635   0.01601  -0.0702   0.4706   1.0000
   6.750   1.0944   0.02701   0.01689  -0.0692   0.4562   1.0000
   7.000   1.1172   0.02768   0.01780  -0.0683   0.4430   1.0000
   7.250   1.1392   0.02837   0.01878  -0.0672   0.4300   1.0000
   7.500   1.1595   0.02918   0.01998  -0.0660   0.4169   1.0000
   7.750   1.1719   0.02939   0.02036  -0.0630   0.3901   1.0000
   8.000   1.1742   0.02947   0.02026  -0.0585   0.3473   1.0000
   8.250   1.1712   0.03031   0.02086  -0.0539   0.2970   1.0000
   8.500   1.1670   0.03164   0.02203  -0.0498   0.2358   1.0000
   8.750   1.1575   0.03375   0.02359  -0.0455   0.1362   1.0000
   9.000   1.1429   0.03709   0.02619  -0.0417   0.0818   1.0000
   9.250   1.1351   0.04019   0.02913  -0.0388   0.0646   1.0000
   9.500   1.1294   0.04314   0.03211  -0.0365   0.0559   1.0000
   9.750   1.1249   0.04606   0.03519  -0.0348   0.0505   1.0000
  10.000   1.1195   0.04919   0.03843  -0.0334   0.0468   1.0000
  10.250   1.1163   0.05227   0.04172  -0.0323   0.0439   1.0000
  10.500   1.1124   0.05554   0.04513  -0.0315   0.0413   1.0000
  10.750   1.1073   0.05906   0.04873  -0.0310   0.0392   1.0000
  11.000   1.1077   0.06212   0.05198  -0.0302   0.0371   1.0000
  11.250   1.1102   0.06502   0.05510  -0.0294   0.0351   1.0000
  11.500   1.1156   0.06769   0.05795  -0.0282   0.0336   1.0000
  11.750   1.1223   0.07033   0.06075  -0.0271   0.0318   1.0000
  12.000   1.1353   0.07269   0.06314  -0.0253   0.0298   1.0000
  12.250   1.1428   0.07583   0.06669  -0.0245   0.0287   1.0000
  12.500   1.1465   0.07954   0.07082  -0.0241   0.0280   1.0000
  12.750   1.1456   0.08386   0.07547  -0.0242   0.0277   1.0000
  13.000   1.1400   0.08869   0.08060  -0.0252   0.0276   1.0000
  13.250   1.1309   0.09403   0.08621  -0.0268   0.0275   1.0000
  13.500   1.1192   0.09991   0.09235  -0.0291   0.0275   1.0000
  13.750   1.1059   0.10625   0.09891  -0.0321   0.0276   1.0000
  14.000   1.0923   0.11293   0.10579  -0.0356   0.0278   1.0000
  14.250   1.0780   0.12018   0.11321  -0.0397   0.0279   1.0000
<< Back to GOE 517 AIRFOIL (goe517-il)

Polar data table (+)

Polar graphs


<< Back to GOE 517 AIRFOIL (goe517-il)