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

GOE 321 (HANSA-BRANDENBURG III.1) AIRFOIL (goe321-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 321 (HANSA-BRANDENBURG III.1) AIRFOIL (goe321-il)
Reynolds number: 50,000
Max Cl/Cd: 34.14 at α=7.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe321-il-50000-n5.txt
Download as CSV file: xf-goe321-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 321 (HANSA-BRANDENBURG III.1) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.2884   0.11383   0.10728  -0.0230   1.0000   0.1018
  -9.000  -0.2913   0.11228   0.10583  -0.0263   1.0000   0.1052
  -8.750  -0.3022   0.11164   0.10535  -0.0307   1.0000   0.1063
  -8.500  -0.2855   0.10622   0.09998  -0.0297   1.0000   0.1083
  -8.250  -0.2709   0.10240   0.09621  -0.0289   1.0000   0.1117
  -8.000  -0.2677   0.09994   0.09388  -0.0295   1.0000   0.1145
  -7.750  -0.2761   0.09861   0.09272  -0.0290   0.9999   0.1167
  -7.500  -0.2663   0.09615   0.09022  -0.0435   0.9620   0.1219
  -7.250  -0.2449   0.09100   0.08508  -0.0463   0.9471   0.1239
  -7.000  -0.2217   0.08682   0.08088  -0.0464   0.9331   0.1278
  -6.750  -0.2041   0.08358   0.07756  -0.0518   0.9150   0.1353
  -6.250  -0.1720   0.07685   0.07074  -0.0572   0.8846   0.1478
  -6.000  -0.1584   0.07393   0.06768  -0.0632   0.8698   0.1575
  -5.750  -0.1414   0.07080   0.06454  -0.0623   0.8587   0.1632
  -5.250  -0.1085   0.06579   0.05923  -0.0680   0.8354   0.1878
  -5.000  -0.0928   0.06259   0.05602  -0.0676   0.8258   0.1951
  -4.750  -0.0759   0.06011   0.05345  -0.0687   0.8161   0.2112
  -4.250  -0.0013   0.05118   0.04333  -0.0804   0.7988   0.1035
  -4.000   0.0280   0.04812   0.03970  -0.0823   0.7906   0.0908
  -3.750   0.0506   0.04570   0.03717  -0.0826   0.7827   0.0883
  -3.500   0.0758   0.04347   0.03466  -0.0832   0.7751   0.0854
  -3.250   0.1024   0.04137   0.03219  -0.0838   0.7672   0.0827
  -3.000   0.1310   0.03940   0.02973  -0.0841   0.7614   0.0807
  -2.750   0.1562   0.03807   0.02818  -0.0844   0.7533   0.0816
  -2.500   0.1832   0.03675   0.02659  -0.0845   0.7472   0.0829
  -2.250   0.2105   0.03556   0.02510  -0.0847   0.7405   0.0834
  -2.000   0.2381   0.03443   0.02367  -0.0847   0.7336   0.0830
  -1.750   0.2670   0.03334   0.02226  -0.0846   0.7286   0.0830
  -1.500   0.2937   0.03261   0.02132  -0.0846   0.7215   0.0834
  -1.250   0.3214   0.03191   0.02039  -0.0845   0.7155   0.0848
  -1.000   0.3502   0.03127   0.01950  -0.0842   0.7108   0.0880
  -0.750   0.3759   0.03102   0.01907  -0.0841   0.7031   0.0906
  -0.500   0.4036   0.03052   0.01839  -0.0836   0.6976   0.0924
  -0.250   0.4312   0.03006   0.01786  -0.0833   0.6919   0.0948
   0.000   0.4567   0.02987   0.01762  -0.0829   0.6844   0.0983
   0.250   0.4845   0.02951   0.01710  -0.0822   0.6789   0.1050
   0.500   0.5074   0.02944   0.01707  -0.0817   0.6698   0.1112
   0.750   0.5342   0.02910   0.01662  -0.0809   0.6635   0.1194
   1.000   0.5573   0.02912   0.01669  -0.0804   0.6547   0.1327
   1.250   0.5840   0.02881   0.01643  -0.0799   0.6483   0.1568
   1.500   0.6115   0.02668   0.01653  -0.0798   0.6408   1.0000
   1.750   0.6370   0.02701   0.01649  -0.0790   0.6336   1.0000
   2.000   0.6615   0.02741   0.01665  -0.0784   0.6264   1.0000
   2.250   0.6854   0.02782   0.01690  -0.0777   0.6184   1.0000
   2.500   0.7105   0.02812   0.01704  -0.0770   0.6113   1.0000
   3.000   0.7582   0.02879   0.01751  -0.0755   0.5938   1.0000
   3.250   0.7841   0.02878   0.01740  -0.0746   0.5846   1.0000
   3.500   0.8066   0.02901   0.01756  -0.0736   0.5727   1.0000
   3.750   0.8325   0.02884   0.01727  -0.0725   0.5617   1.0000
   4.000   0.8595   0.02854   0.01685  -0.0714   0.5506   1.0000
   4.250   0.8814   0.02875   0.01703  -0.0703   0.5372   1.0000
   4.500   0.9047   0.02891   0.01716  -0.0693   0.5254   1.0000
   4.750   0.9322   0.02876   0.01692  -0.0685   0.5162   1.0000
   5.000   0.9521   0.02928   0.01749  -0.0675   0.5031   1.0000
   5.250   0.9735   0.02966   0.01788  -0.0665   0.4907   1.0000
   5.500   0.9975   0.02981   0.01801  -0.0656   0.4794   1.0000
   5.750   1.0190   0.03013   0.01836  -0.0645   0.4667   1.0000
   6.000   1.0375   0.03069   0.01897  -0.0634   0.4522   1.0000
   6.250   1.0561   0.03120   0.01955  -0.0622   0.4374   1.0000
   6.500   1.0745   0.03170   0.02008  -0.0610   0.4225   1.0000
   6.750   1.0928   0.03218   0.02058  -0.0597   0.4075   1.0000
   7.000   1.1113   0.03265   0.02106  -0.0585   0.3928   1.0000
   7.250   1.1299   0.03315   0.02151  -0.0572   0.3789   1.0000
   7.500   1.1484   0.03369   0.02200  -0.0559   0.3661   1.0000
   7.750   1.1680   0.03421   0.02243  -0.0547   0.3548   1.0000
   8.000   1.1855   0.03500   0.02319  -0.0535   0.3438   1.0000
   8.250   1.2015   0.03599   0.02419  -0.0523   0.3336   1.0000
   8.750   1.2353   0.03798   0.02616  -0.0501   0.3163   1.0000
   9.000   1.2582   0.03863   0.02674  -0.0494   0.3094   1.0000
   9.250   1.2675   0.04025   0.02852  -0.0480   0.3019   1.0000
   9.500   1.2863   0.04122   0.02951  -0.0471   0.2957   1.0000
   9.750   1.3029   0.04243   0.03078  -0.0461   0.2904   1.0000
  10.000   1.3107   0.04420   0.03276  -0.0447   0.2854   1.0000
  10.250   1.3268   0.04553   0.03418  -0.0439   0.2811   1.0000
  10.500   1.3540   0.04629   0.03496  -0.0437   0.2775   1.0000
  10.750   1.3542   0.04871   0.03762  -0.0422   0.2739   1.0000
  11.000   1.3521   0.05142   0.04057  -0.0408   0.2703   1.0000
  11.250   1.3565   0.05374   0.04308  -0.0398   0.2669   1.0000
  11.500   1.3732   0.05512   0.04457  -0.0392   0.2637   1.0000
  11.750   1.3894   0.05662   0.04617  -0.0386   0.2607   1.0000
  12.000   1.3420   0.06385   0.05374  -0.0380   0.2574   1.0000
  12.250   1.2534   0.07811   0.06825  -0.0410   0.2525   1.0000
  12.500   1.2176   0.08721   0.07748  -0.0435   0.2487   1.0000
  12.750   1.2275   0.08951   0.07991  -0.0432   0.2467   1.0000
<< Back to GOE 321 (HANSA-BRANDENBURG III.1) AIRFOIL (goe321-il)

Polar data table (+)

Polar graphs


<< Back to GOE 321 (HANSA-BRANDENBURG III.1) AIRFOIL (goe321-il)