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

GOE 439 AIRFOIL (goe439-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 439 AIRFOIL (goe439-il)
Reynolds number: 100,000
Max Cl/Cd: 59.13 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe439-il-100000.txt
Download as CSV file: xf-goe439-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 439 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3060   0.10477   0.10011  -0.0323   1.0000   0.0611
  -8.000  -0.3157   0.10406   0.09952  -0.0325   1.0000   0.0614
  -7.750  -0.3241   0.10307   0.09865  -0.0335   1.0000   0.0615
  -7.500  -0.3284   0.10165   0.09733  -0.0358   1.0000   0.0617
  -7.250  -0.3040   0.09288   0.08853  -0.0280   1.0000   0.0644
  -7.000  -0.3074   0.09088   0.08661  -0.0260   1.0000   0.0657
  -6.750  -0.3122   0.08903   0.08485  -0.0245   1.0000   0.0672
  -6.500  -0.3192   0.08740   0.08332  -0.0231   1.0000   0.0691
  -6.250  -0.3281   0.08603   0.08204  -0.0222   1.0000   0.0712
  -6.000  -0.3360   0.08466   0.08075  -0.0218   1.0000   0.0722
  -5.750  -0.3387   0.08421   0.08030  -0.0284   1.0000   0.0742
  -5.500  -0.3371   0.08138   0.07748  -0.0309   1.0000   0.0751
  -5.250  -0.3436   0.07809   0.07431  -0.0238   1.0000   0.0766
  -5.000  -0.3403   0.07567   0.07191  -0.0220   0.9993   0.0793
  -4.750  -0.2761   0.07008   0.06602  -0.0421   0.9899   0.0885
  -4.500  -0.2550   0.06559   0.06160  -0.0417   0.9842   0.0916
  -4.250  -0.1973   0.06077   0.05644  -0.0561   0.9748   0.1024
  -4.000  -0.1714   0.05659   0.05233  -0.0571   0.9688   0.1066
  -3.750  -0.1247   0.05216   0.04763  -0.0659   0.9598   0.1179
  -3.500  -0.0852   0.04847   0.04376  -0.0712   0.9518   0.1322
  -3.250  -0.0439   0.04478   0.03991  -0.0761   0.9443   0.1473
  -3.000  -0.0060   0.04154   0.03649  -0.0801   0.9354   0.1621
  -2.750   0.0419   0.03832   0.03297  -0.0857   0.9291   0.1888
  -2.500   0.1018   0.03134   0.02472  -0.0908   0.9217   0.0993
  -2.250   0.1521   0.02714   0.01949  -0.0939   0.9163   0.0874
  -2.000   0.1897   0.02524   0.01744  -0.0960   0.9072   0.0960
  -1.750   0.2393   0.02303   0.01460  -0.0991   0.9016   0.1033
  -1.500   0.2772   0.02184   0.01320  -0.1006   0.8917   0.1170
  -1.250   0.3241   0.02037   0.01167  -0.1035   0.8858   0.1368
  -1.000   0.3579   0.01935   0.01053  -0.1039   0.8743   0.1578
  -0.750   0.3915   0.01819   0.00959  -0.1045   0.8633   0.2018
  -0.500   0.4267   0.01713   0.00879  -0.1053   0.8532   0.2670
  -0.250   0.4774   0.01464   0.00803  -0.1087   0.8446   1.0000
   0.000   0.5079   0.01459   0.00764  -0.1084   0.8310   1.0000
   0.250   0.5372   0.01456   0.00737  -0.1079   0.8170   1.0000
   0.500   0.5654   0.01457   0.00716  -0.1073   0.8027   1.0000
   0.750   0.5930   0.01461   0.00701  -0.1066   0.7881   1.0000
   1.000   0.6200   0.01467   0.00691  -0.1059   0.7732   1.0000
   1.250   0.6465   0.01477   0.00686  -0.1051   0.7582   1.0000
   1.500   0.6728   0.01489   0.00684  -0.1042   0.7432   1.0000
   1.750   0.6986   0.01505   0.00686  -0.1034   0.7280   1.0000
   2.000   0.7235   0.01524   0.00696  -0.1024   0.7121   1.0000
   2.250   0.7485   0.01544   0.00709  -0.1016   0.6965   1.0000
   2.500   0.7736   0.01565   0.00721  -0.1007   0.6811   1.0000
   2.750   0.7987   0.01586   0.00734  -0.0999   0.6659   1.0000
   3.000   0.8239   0.01608   0.00750  -0.0991   0.6510   1.0000
   3.250   0.8490   0.01631   0.00770  -0.0983   0.6363   1.0000
   3.500   0.8741   0.01656   0.00791  -0.0975   0.6220   1.0000
   3.750   0.8992   0.01683   0.00813  -0.0968   0.6078   1.0000
   4.000   0.9244   0.01711   0.00839  -0.0961   0.5940   1.0000
   4.250   0.9495   0.01743   0.00873  -0.0954   0.5806   1.0000
   4.500   0.9746   0.01778   0.00906  -0.0948   0.5675   1.0000
   4.750   1.0000   0.01815   0.00942  -0.0941   0.5549   1.0000
   5.000   1.0257   0.01855   0.00979  -0.0936   0.5428   1.0000
   5.250   1.0498   0.01901   0.01035  -0.0929   0.5300   1.0000
   5.500   1.0738   0.01951   0.01095  -0.0922   0.5177   1.0000
   5.750   1.0979   0.02004   0.01157  -0.0915   0.5057   1.0000
   6.000   1.1216   0.02042   0.01201  -0.0906   0.4916   1.0000
   6.250   1.1446   0.02061   0.01219  -0.0894   0.4745   1.0000
   6.500   1.1674   0.02074   0.01232  -0.0882   0.4566   1.0000
   6.750   1.1864   0.02083   0.01255  -0.0864   0.4356   1.0000
   7.000   1.2052   0.02079   0.01250  -0.0844   0.4122   1.0000
   7.250   1.2230   0.02089   0.01274  -0.0825   0.3897   1.0000
   7.500   1.2376   0.02093   0.01287  -0.0799   0.3591   1.0000
   7.750   1.2499   0.02120   0.01319  -0.0770   0.3175   1.0000
   8.000   1.2553   0.02218   0.01382  -0.0734   0.2338   1.0000
   8.250   1.2421   0.02546   0.01584  -0.0684   0.0889   1.0000
   8.500   1.2437   0.02764   0.01788  -0.0649   0.0660   1.0000
   8.750   1.2483   0.02929   0.01963  -0.0617   0.0589   1.0000
   9.000   1.2499   0.03090   0.02136  -0.0582   0.0549   1.0000
   9.250   1.2453   0.03291   0.02345  -0.0545   0.0524   1.0000
   9.500   1.2457   0.03470   0.02540  -0.0515   0.0500   1.0000
   9.750   1.2453   0.03666   0.02750  -0.0489   0.0479   1.0000
  10.000   1.2443   0.03886   0.02976  -0.0466   0.0456   1.0000
  10.250   1.2461   0.04110   0.03207  -0.0445   0.0444   1.0000
  10.500   1.2585   0.04327   0.03417  -0.0425   0.0431   1.0000
  10.750   1.3182   0.04613   0.03691  -0.0433   0.0421   1.0000
  11.000   1.3450   0.04879   0.03986  -0.0428   0.0412   1.0000
  11.250   1.3631   0.05166   0.04306  -0.0418   0.0403   1.0000
  11.500   1.3827   0.05556   0.04731  -0.0410   0.0406   1.0000
  11.750   1.3982   0.06023   0.05234  -0.0403   0.0414   1.0000
  12.000   1.4169   0.06528   0.05771  -0.0401   0.0427   1.0000
  12.250   1.4043   0.06666   0.05945  -0.0360   0.0436   1.0000
  12.500   1.3810   0.06959   0.06282  -0.0327   0.0447   1.0000
  12.750   1.3561   0.07360   0.06724  -0.0308   0.0457   1.0000
  13.000   1.3339   0.07836   0.07233  -0.0302   0.0468   1.0000
  13.250   1.3122   0.08352   0.07778  -0.0306   0.0476   1.0000
  13.500   1.2905   0.08908   0.08358  -0.0318   0.0484   1.0000
  13.750   1.2682   0.09510   0.08982  -0.0339   0.0490   1.0000
  14.000   1.2452   0.10159   0.09651  -0.0368   0.0494   1.0000
  14.250   1.2215   0.10865   0.10370  -0.0406   0.0498   1.0000
  14.500   1.1963   0.11646   0.11168  -0.0455   0.0500   1.0000
  14.750   1.1706   0.12515   0.12050  -0.0514   0.0502   1.0000
<< Back to GOE 439 AIRFOIL (goe439-il)

Polar data table (+)

Polar graphs


<< Back to GOE 439 AIRFOIL (goe439-il)