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

GOE 425 AIRFOIL (goe425-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 425 AIRFOIL (goe425-il)
Reynolds number: 500,000
Max Cl/Cd: 111 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe425-il-500000.txt
Download as CSV file: xf-goe425-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 425 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.2990   0.10977   0.10674  -0.0406   0.7819   0.0353
 -11.000  -0.2955   0.10623   0.10314  -0.0423   0.7696   0.0362
 -10.750  -0.3319   0.09212   0.08905  -0.0535   0.7662   0.0387
 -10.500  -0.3234   0.09018   0.08705  -0.0532   0.7557   0.0389
 -10.250  -0.3145   0.08820   0.08501  -0.0535   0.7461   0.0392
 -10.000  -0.3067   0.08585   0.08263  -0.0542   0.7379   0.0396
  -9.750  -0.3008   0.08307   0.07981  -0.0555   0.7301   0.0400
  -9.500  -0.2978   0.07949   0.07620  -0.0577   0.7235   0.0407
  -9.250  -0.2997   0.07454   0.07125  -0.0613   0.7171   0.0416
  -9.000  -0.3688   0.03643   0.03267  -0.0886   0.7053   0.0452
  -8.750  -0.3557   0.03427   0.03048  -0.0886   0.6985   0.0455
  -8.500  -0.3446   0.03209   0.02820  -0.0885   0.6922   0.0459
  -8.250  -0.3321   0.02996   0.02601  -0.0885   0.6856   0.0464
  -8.000  -0.3204   0.02763   0.02354  -0.0885   0.6791   0.0473
  -7.750  -0.3302   0.02207   0.01713  -0.0876   0.6744   0.0519
  -7.500  -0.3086   0.02028   0.01543  -0.0877   0.6681   0.0526
  -7.250  -0.2877   0.01898   0.01410  -0.0876   0.6616   0.0533
  -7.000  -0.2672   0.01780   0.01280  -0.0873   0.6557   0.0546
  -6.750  -0.2532   0.01553   0.01011  -0.0865   0.6496   0.0602
  -6.500  -0.2488   0.02212   0.01556  -0.0863   0.6532   0.0425
  -6.250  -0.2252   0.02003   0.01312  -0.0856   0.6468   0.0403
  -6.000  -0.1999   0.01900   0.01187  -0.0852   0.6407   0.0401
  -5.750  -0.1736   0.01821   0.01100  -0.0850   0.6344   0.0405
  -5.500  -0.1472   0.01739   0.01005  -0.0847   0.6278   0.0405
  -5.250  -0.1207   0.01669   0.00922  -0.0844   0.6221   0.0405
  -5.000  -0.0938   0.01604   0.00851  -0.0842   0.6165   0.0406
  -4.750  -0.0669   0.01547   0.00788  -0.0839   0.6108   0.0409
  -4.500  -0.0403   0.01498   0.00731  -0.0837   0.6057   0.0411
  -4.250  -0.0134   0.01451   0.00681  -0.0835   0.6006   0.0415
  -4.000   0.0135   0.01408   0.00635  -0.0832   0.5952   0.0419
  -3.750   0.0402   0.01373   0.00594  -0.0830   0.5904   0.0425
  -3.500   0.0672   0.01345   0.00561  -0.0828   0.5861   0.0433
  -3.250   0.0945   0.01316   0.00530  -0.0826   0.5818   0.0439
  -3.000   0.1217   0.01290   0.00501  -0.0825   0.5774   0.0443
  -2.750   0.1478   0.01250   0.00456  -0.0821   0.5733   0.0450
  -2.500   0.1742   0.01218   0.00421  -0.0819   0.5694   0.0460
  -2.250   0.2015   0.01192   0.00397  -0.0818   0.5658   0.0471
  -2.000   0.2291   0.01173   0.00376  -0.0817   0.5620   0.0484
  -1.750   0.2567   0.01159   0.00358  -0.0816   0.5584   0.0499
  -1.500   0.2844   0.01151   0.00343  -0.0816   0.5547   0.0520
  -1.250   0.3122   0.01135   0.00327  -0.0815   0.5516   0.0557
  -1.000   0.3401   0.01122   0.00315  -0.0815   0.5482   0.0620
  -0.750   0.3674   0.01096   0.00310  -0.0814   0.5448   0.1172
  -0.500   0.3917   0.01024   0.00297  -0.0812   0.5416   0.2991
  -0.250   0.4166   0.00983   0.00303  -0.0809   0.5384   0.4599
   0.000   0.4439   0.00978   0.00311  -0.0808   0.5353   0.5135
   0.250   0.4714   0.00972   0.00316  -0.0808   0.5322   0.5506
   0.500   0.4992   0.00970   0.00320  -0.0807   0.5290   0.5791
   0.750   0.5267   0.00968   0.00325  -0.0806   0.5259   0.6048
   1.000   0.5541   0.00968   0.00330  -0.0805   0.5230   0.6364
   1.250   0.5811   0.00972   0.00340  -0.0803   0.5197   0.6721
   1.500   0.6069   0.00956   0.00347  -0.0799   0.5171   0.7246
   1.750   0.6294   0.00924   0.00357  -0.0785   0.5143   0.8344
   2.250   0.7421   0.00928   0.00374  -0.0905   0.5076   1.0000
   2.500   0.7683   0.00944   0.00382  -0.0903   0.5045   1.0000
   2.750   0.7943   0.00958   0.00393  -0.0900   0.5017   1.0000
   3.000   0.8202   0.00966   0.00402  -0.0896   0.4990   1.0000
   3.250   0.8461   0.00975   0.00410  -0.0893   0.4961   1.0000
   3.500   0.8721   0.00986   0.00419  -0.0890   0.4932   1.0000
   3.750   0.8981   0.00998   0.00427  -0.0886   0.4902   1.0000
   4.000   0.9245   0.01020   0.00441  -0.0884   0.4864   1.0000
   4.250   0.9498   0.01025   0.00451  -0.0880   0.4832   1.0000
   4.500   0.9755   0.01033   0.00461  -0.0876   0.4796   1.0000
   4.750   1.0015   0.01043   0.00471  -0.0873   0.4762   1.0000
   5.000   1.0275   0.01056   0.00481  -0.0870   0.4729   1.0000
   5.250   1.0537   0.01077   0.00497  -0.0868   0.4693   1.0000
   5.500   1.0792   0.01084   0.00511  -0.0865   0.4658   1.0000
   5.750   1.1049   0.01094   0.00523  -0.0862   0.4620   1.0000
   6.000   1.1306   0.01104   0.00533  -0.0859   0.4580   1.0000
   6.250   1.1565   0.01124   0.00547  -0.0856   0.4537   1.0000
   6.500   1.1817   0.01132   0.00564  -0.0852   0.4496   1.0000
   6.750   1.2070   0.01141   0.00577  -0.0849   0.4450   1.0000
   7.000   1.2321   0.01154   0.00588  -0.0845   0.4401   1.0000
   7.250   1.2571   0.01170   0.00606  -0.0841   0.4354   1.0000
   7.500   1.2822   0.01181   0.00624  -0.0838   0.4305   1.0000
   7.750   1.3069   0.01196   0.00639  -0.0834   0.4256   1.0000
   8.000   1.3312   0.01216   0.00659  -0.0829   0.4206   1.0000
   8.250   1.3557   0.01228   0.00679  -0.0825   0.4146   1.0000
   8.500   1.3791   0.01246   0.00697  -0.0819   0.4082   1.0000
   8.750   1.4029   0.01266   0.00722  -0.0814   0.4026   1.0000
   9.000   1.4263   0.01285   0.00745  -0.0809   0.3961   1.0000
   9.250   1.4481   0.01311   0.00770  -0.0801   0.3892   1.0000
   9.500   1.4702   0.01333   0.00798  -0.0793   0.3801   1.0000
   9.750   1.4904   0.01364   0.00829  -0.0783   0.3707   1.0000
  10.000   1.5085   0.01402   0.00865  -0.0770   0.3600   1.0000
  10.250   1.5273   0.01438   0.00904  -0.0758   0.3488   1.0000
  10.500   1.5422   0.01483   0.00950  -0.0739   0.3374   1.0000
  10.750   1.5540   0.01541   0.01005  -0.0717   0.3244   1.0000
  11.000   1.5640   0.01612   0.01074  -0.0693   0.3096   1.0000
  11.250   1.5720   0.01700   0.01159  -0.0670   0.2932   1.0000
  11.500   1.5763   0.01817   0.01271  -0.0644   0.2732   1.0000
  11.750   1.5739   0.01991   0.01434  -0.0616   0.2474   1.0000
  12.000   1.5615   0.02263   0.01690  -0.0587   0.2136   1.0000
  12.250   1.5459   0.02604   0.02017  -0.0564   0.1862   1.0000
  12.500   1.5348   0.02933   0.02339  -0.0548   0.1703   1.0000
  12.750   1.5285   0.03228   0.02634  -0.0536   0.1598   1.0000
  13.000   1.5217   0.03532   0.02939  -0.0524   0.1516   1.0000
  13.250   1.5189   0.03802   0.03212  -0.0514   0.1439   1.0000
  13.500   1.5128   0.04106   0.03519  -0.0505   0.1370   1.0000
  13.750   1.5100   0.04385   0.03801  -0.0497   0.1293   1.0000
  14.000   1.5053   0.04691   0.04110  -0.0490   0.1226   1.0000
  14.250   1.5021   0.04990   0.04410  -0.0485   0.1153   1.0000
  14.500   1.4985   0.05300   0.04723  -0.0480   0.1090   1.0000
  14.750   1.4946   0.05620   0.05044  -0.0477   0.1027   1.0000
  15.000   1.4907   0.05944   0.05371  -0.0474   0.0975   1.0000
  15.250   1.4874   0.06266   0.05696  -0.0472   0.0927   1.0000
  15.500   1.4825   0.06611   0.06043  -0.0470   0.0885   1.0000
  15.750   1.4818   0.06908   0.06346  -0.0469   0.0851   1.0000
  16.000   1.4785   0.07239   0.06680  -0.0469   0.0816   1.0000
  16.250   1.4731   0.07598   0.07042  -0.0469   0.0782   1.0000
  16.500   1.4748   0.07876   0.07328  -0.0470   0.0752   1.0000
  16.750   1.4729   0.08198   0.07653  -0.0471   0.0716   1.0000
  17.000   1.4677   0.08566   0.08024  -0.0473   0.0681   1.0000
  17.250   1.4699   0.08841   0.08305  -0.0475   0.0638   1.0000
  17.500   1.4649   0.09214   0.08680  -0.0479   0.0590   1.0000
  17.750   1.4619   0.09562   0.09027  -0.0483   0.0521   1.0000
  18.000   1.4553   0.09960   0.09423  -0.0488   0.0439   1.0000
  18.250   1.4469   0.10385   0.09845  -0.0494   0.0360   1.0000
  18.500   1.4383   0.10822   0.10281  -0.0502   0.0310   1.0000
<< Back to GOE 425 AIRFOIL (goe425-il)

Polar data table (+)

Polar graphs


<< Back to GOE 425 AIRFOIL (goe425-il)