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

GOE 382 AIRFOIL (goe382-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 382 AIRFOIL (goe382-il)
Reynolds number: 200,000
Max Cl/Cd: 60.87 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe382-il-200000.txt
Download as CSV file: xf-goe382-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 382 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.0333   0.10221   0.09835  -0.0867   0.8408   0.0774
 -10.500  -0.0289   0.09951   0.09557  -0.0869   0.8298   0.0789
 -10.250  -0.0693   0.09109   0.08716  -0.0973   0.8196   0.0831
 -10.000  -0.0486   0.08985   0.08584  -0.0942   0.8092   0.0837
  -9.750  -0.0328   0.08827   0.08420  -0.0928   0.7983   0.0846
  -9.500  -0.0225   0.08610   0.08197  -0.0928   0.7887   0.0861
  -9.250  -0.0194   0.08296   0.07879  -0.0944   0.7797   0.0883
  -9.000  -0.0562   0.07351   0.06933  -0.1050   0.7735   0.0927
  -8.750  -0.0390   0.07236   0.06815  -0.1033   0.7644   0.0935
  -8.500  -0.0287   0.07029   0.06601  -0.1032   0.7574   0.0949
  -8.250  -0.1112   0.05795   0.05335  -0.1173   0.7510   0.1021
  -8.000  -0.0859   0.05647   0.05193  -0.1165   0.7438   0.1030
  -7.750  -0.0684   0.05502   0.05042  -0.1158   0.7380   0.1044
  -7.500  -0.1444   0.03616   0.02965  -0.1151   0.7350   0.0663
  -7.250  -0.1271   0.03388   0.02723  -0.1144   0.7283   0.0651
  -7.000  -0.1105   0.03177   0.02477  -0.1133   0.7228   0.0644
  -6.750  -0.0917   0.02995   0.02258  -0.1123   0.7178   0.0639
  -6.500  -0.0719   0.02833   0.02068  -0.1113   0.7116   0.0632
  -6.250  -0.0497   0.02687   0.01890  -0.1105   0.7061   0.0626
  -6.000  -0.0252   0.02569   0.01743  -0.1099   0.7014   0.0625
  -5.750  -0.0009   0.02475   0.01633  -0.1094   0.6959   0.0627
  -5.500   0.0242   0.02390   0.01535  -0.1089   0.6901   0.0631
  -5.250   0.0507   0.02313   0.01441  -0.1086   0.6852   0.0636
  -5.000   0.0779   0.02249   0.01360  -0.1084   0.6806   0.0644
  -4.750   0.1028   0.02199   0.01303  -0.1078   0.6747   0.0657
  -4.500   0.1293   0.02163   0.01248  -0.1074   0.6692   0.0671
  -4.250   0.1569   0.02080   0.01168  -0.1074   0.6647   0.0685
  -4.000   0.1830   0.02033   0.01122  -0.1071   0.6598   0.0701
  -3.750   0.2080   0.01990   0.01081  -0.1066   0.6539   0.0719
  -3.500   0.2346   0.01954   0.01037  -0.1062   0.6487   0.0740
  -3.250   0.2607   0.01897   0.00981  -0.1059   0.6443   0.0768
  -3.000   0.2838   0.01862   0.00953  -0.1052   0.6389   0.0805
  -2.750   0.3067   0.01813   0.00910  -0.1044   0.6332   0.0863
  -2.500   0.3309   0.01759   0.00858  -0.1038   0.6283   0.0981
  -2.250   0.3559   0.01709   0.00836  -0.1034   0.6238   0.1657
  -1.750   0.4001   0.01637   0.00830  -0.1018   0.6124   0.3149
  -1.500   0.4243   0.01596   0.00818  -0.1013   0.6077   0.4060
  -1.250   0.4436   0.01557   0.00836  -0.0997   0.6027   0.5615
  -1.000   0.4649   0.01546   0.00849  -0.0983   0.5965   0.6417
  -0.750   0.4891   0.01534   0.00852  -0.0973   0.5914   0.7029
  -0.500   0.5155   0.01532   0.00857  -0.0965   0.5871   0.7600
  -0.250   0.5368   0.01539   0.00882  -0.0948   0.5811   0.8137
   0.000   0.5627   0.01546   0.00896  -0.0938   0.5752   0.8641
   0.250   0.5962   0.01556   0.00897  -0.0943   0.5702   0.9002
   0.500   0.6308   0.01579   0.00913  -0.0954   0.5649   0.9278
   0.750   0.6678   0.01598   0.00929  -0.0971   0.5585   0.9497
   1.000   0.7140   0.01611   0.00928  -0.1007   0.5530   0.9660
   1.250   0.7662   0.01630   0.00930  -0.1056   0.5475   0.9771
   1.500   0.8179   0.01645   0.00943  -0.1108   0.5403   0.9888
   1.750   0.8779   0.01650   0.00933  -0.1175   0.5346   0.9992
   2.000   0.9009   0.01661   0.00929  -0.1171   0.5303   1.0000
   2.250   0.9119   0.01677   0.00950  -0.1145   0.5245   1.0000
   2.500   0.9287   0.01687   0.00953  -0.1128   0.5194   1.0000
   2.750   0.9498   0.01695   0.00948  -0.1117   0.5151   1.0000
   3.000   0.9665   0.01716   0.00966  -0.1100   0.5102   1.0000
   3.250   0.9817   0.01736   0.00986  -0.1079   0.5049   1.0000
   3.500   1.0020   0.01750   0.00992  -0.1067   0.5004   1.0000
   3.750   1.0269   0.01766   0.00993  -0.1063   0.4965   1.0000
   4.000   1.0415   0.01795   0.01025  -0.1042   0.4915   1.0000
   4.250   1.0588   0.01817   0.01046  -0.1025   0.4864   1.0000
   4.500   1.0814   0.01833   0.01053  -0.1017   0.4820   1.0000
   4.750   1.1095   0.01856   0.01062  -0.1020   0.4782   1.0000
   5.000   1.1223   0.01892   0.01106  -0.0996   0.4734   1.0000
   5.250   1.1418   0.01920   0.01134  -0.0984   0.4688   1.0000
   5.500   1.1660   0.01943   0.01150  -0.0980   0.4648   1.0000
   5.750   1.1965   0.01970   0.01162  -0.0988   0.4611   1.0000
   6.000   1.2083   0.02012   0.01215  -0.0964   0.4565   1.0000
   6.250   1.2269   0.02046   0.01251  -0.0951   0.4520   1.0000
   6.500   1.2509   0.02074   0.01273  -0.0948   0.4482   1.0000
   6.750   1.2808   0.02104   0.01293  -0.0955   0.4448   1.0000
   7.000   1.2971   0.02155   0.01350  -0.0940   0.4410   1.0000
   7.250   1.3118   0.02202   0.01404  -0.0922   0.4369   1.0000
   7.500   1.3325   0.02239   0.01440  -0.0914   0.4330   1.0000
   7.750   1.3595   0.02267   0.01463  -0.0917   0.4294   1.0000
   8.000   1.3889   0.02312   0.01500  -0.0925   0.4260   1.0000
   8.250   1.3950   0.02375   0.01577  -0.0894   0.4223   1.0000
   8.500   1.4074   0.02429   0.01637  -0.0874   0.4187   1.0000
   8.750   1.4274   0.02472   0.01682  -0.0866   0.4153   1.0000
   9.000   1.4563   0.02505   0.01708  -0.0873   0.4122   1.0000
   9.250   1.4880   0.02555   0.01752  -0.0885   0.4089   1.0000
   9.500   1.4836   0.02639   0.01854  -0.0840   0.4054   1.0000
   9.750   1.4913   0.02716   0.01940  -0.0816   0.4019   1.0000
  10.000   1.5085   0.02773   0.02001  -0.0806   0.3987   1.0000
  10.250   1.5356   0.02809   0.02033  -0.0810   0.3958   1.0000
  10.500   1.5768   0.02835   0.02050  -0.0836   0.3928   1.0000
  10.750   1.5673   0.02958   0.02192  -0.0789   0.3896   1.0000
  11.000   1.5635   0.03082   0.02331  -0.0754   0.3861   1.0000
  11.250   1.5730   0.03171   0.02427  -0.0737   0.3828   1.0000
  11.500   1.5978   0.03197   0.02451  -0.0738   0.3793   1.0000
  11.750   1.6460   0.03168   0.02405  -0.0768   0.3754   1.0000
  12.000   1.6155   0.03382   0.02645  -0.0705   0.3717   1.0000
  12.250   1.6063   0.03543   0.02820  -0.0671   0.3672   1.0000
  12.500   1.6272   0.03559   0.02833  -0.0667   0.3627   1.0000
  12.750   1.6747   0.03489   0.02746  -0.0691   0.3587   1.0000
  13.000   1.6380   0.03799   0.03085  -0.0635   0.3549   1.0000
  13.250   1.6257   0.04030   0.03331  -0.0607   0.3508   1.0000
  13.500   1.6371   0.04128   0.03434  -0.0598   0.3474   1.0000
  13.750   1.6679   0.04110   0.03411  -0.0604   0.3445   1.0000
  14.000   1.6870   0.04174   0.03476  -0.0602   0.3412   1.0000
  14.250   1.6374   0.04689   0.04022  -0.0556   0.3368   1.0000
  14.500   1.6280   0.04953   0.04297  -0.0539   0.3324   1.0000
  14.750   1.6624   0.04861   0.04198  -0.0544   0.3287   1.0000
  15.000   1.6776   0.04933   0.04269  -0.0540   0.3246   1.0000
  15.250   1.6102   0.05748   0.05118  -0.0506   0.3193   1.0000
  15.500   1.6201   0.05865   0.05237  -0.0503   0.3147   1.0000
  15.750   1.6760   0.05538   0.04893  -0.0512   0.3107   1.0000
  16.000   1.5999   0.06542   0.05932  -0.0488   0.3051   1.0000
  16.250   1.5803   0.07001   0.06403  -0.0483   0.2995   1.0000
  16.500   1.6364   0.06620   0.06009  -0.0488   0.2959   1.0000
  16.750   1.4268   0.09446   0.08888  -0.0480   0.2826   1.0000
  17.000   1.5474   0.08166   0.07593  -0.0478   0.2832   1.0000
  17.250   1.6108   0.07660   0.07072  -0.0480   0.2801   1.0000
  17.500   1.4609   0.09842   0.09295  -0.0487   0.2680   1.0000
  17.750   1.5408   0.09067   0.08509  -0.0484   0.2667   1.0000
  18.000   1.6092   0.08471   0.07897  -0.0482   0.2637   1.0000
  18.250   1.4944   0.10260   0.09721  -0.0497   0.2524   1.0000
  18.500   1.5576   0.09694   0.09144  -0.0493   0.2497   1.0000
<< Back to GOE 382 AIRFOIL (goe382-il)

Polar data table (+)

Polar graphs


<< Back to GOE 382 AIRFOIL (goe382-il)