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

GOE 432 AIRFOIL (goe432-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 432 AIRFOIL (goe432-il)
Reynolds number: 50,000
Max Cl/Cd: 35.44 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe432-il-50000.txt
Download as CSV file: xf-goe432-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 432 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3586   0.10306   0.09627  -0.0183   1.0000   0.1650
  -7.750  -0.3641   0.10157   0.09490  -0.0180   1.0000   0.1723
  -7.500  -0.3841   0.10181   0.09535  -0.0190   1.0000   0.1743
  -7.250  -0.3644   0.09642   0.08993  -0.0161   1.0000   0.1861
  -7.000  -0.3823   0.09651   0.09019  -0.0189   1.0000   0.1893
  -6.750  -0.3670   0.09160   0.08531  -0.0150   1.0000   0.2015
  -6.500  -0.3661   0.08865   0.08246  -0.0144   1.0000   0.2086
  -6.250  -0.3727   0.08758   0.08149  -0.0166   1.0000   0.2189
  -6.000  -0.3619   0.08338   0.07734  -0.0129   1.0000   0.2287
  -5.750  -0.3618   0.08078   0.07483  -0.0129   1.0000   0.2388
  -5.500  -0.3617   0.07850   0.07262  -0.0137   1.0000   0.2516
  -5.250  -0.3584   0.07592   0.07011  -0.0134   1.0000   0.2661
  -5.000  -0.3535   0.07298   0.06724  -0.0118   1.0000   0.2816
  -4.750  -0.3480   0.07000   0.06430  -0.0092   1.0000   0.2990
  -4.500  -0.3449   0.06793   0.06228  -0.0094   1.0000   0.3251
  -4.250  -0.3415   0.06540   0.05982  -0.0067   1.0000   0.3546
  -4.000  -0.3387   0.06256   0.05708  -0.0024   1.0000   0.3853
  -3.750  -0.3366   0.05970   0.05432   0.0030   1.0000   0.4182
  -3.000  -0.3457   0.05268   0.04766   0.0238   1.0000   0.5657
  -2.750  -0.3482   0.05001   0.04510   0.0332   1.0000   0.6195
  -2.500  -0.1273   0.03722   0.02941  -0.0420   1.0000   0.1868
  -2.250  -0.0948   0.03406   0.02560  -0.0436   1.0000   0.1729
  -2.000  -0.0681   0.03180   0.02297  -0.0438   1.0000   0.1709
  -1.750  -0.0407   0.02980   0.02054  -0.0439   1.0000   0.1681
  -1.500  -0.0148   0.02846   0.01872  -0.0437   1.0000   0.1731
  -1.250   0.0100   0.02719   0.01715  -0.0434   1.0000   0.1767
  -1.000   0.0314   0.02638   0.01623  -0.0427   1.0000   0.1835
  -0.750   0.0517   0.02584   0.01553  -0.0420   1.0000   0.1939
  -0.500   0.0736   0.02552   0.01499  -0.0416   1.0000   0.2031
  -0.250   0.0941   0.02542   0.01485  -0.0414   1.0000   0.2166
   0.000   0.1184   0.02554   0.01502  -0.0424   0.9975   0.2411
   0.250   0.1835   0.02363   0.01474  -0.0494   0.9829   0.5048
   0.500   0.2572   0.02355   0.01440  -0.0579   0.9572   1.0000
   0.750   0.3146   0.02424   0.01475  -0.0641   0.9311   1.0000
   1.000   0.3685   0.02470   0.01497  -0.0693   0.9064   1.0000
   1.250   0.4250   0.02493   0.01503  -0.0745   0.8850   1.0000
   1.500   0.4695   0.02511   0.01510  -0.0774   0.8618   1.0000
   1.750   0.5239   0.02499   0.01491  -0.0816   0.8430   1.0000
   2.000   0.5710   0.02483   0.01469  -0.0841   0.8231   1.0000
   2.250   0.6129   0.02466   0.01450  -0.0855   0.8018   1.0000
   2.500   0.6605   0.02413   0.01392  -0.0872   0.7822   1.0000
   2.750   0.6922   0.02412   0.01388  -0.0866   0.7585   1.0000
   3.000   0.7320   0.02370   0.01341  -0.0867   0.7370   1.0000
   3.250   0.7581   0.02388   0.01356  -0.0853   0.7125   1.0000
   3.500   0.7894   0.02385   0.01346  -0.0843   0.6904   1.0000
   3.750   0.8141   0.02423   0.01383  -0.0830   0.6681   1.0000
   4.000   0.8430   0.02450   0.01408  -0.0820   0.6484   1.0000
   4.250   0.8660   0.02507   0.01467  -0.0806   0.6266   1.0000
   4.500   0.8927   0.02543   0.01499  -0.0793   0.6048   1.0000
   4.750   0.9154   0.02603   0.01563  -0.0777   0.5808   1.0000
   5.000   0.9410   0.02655   0.01607  -0.0762   0.5564   1.0000
   5.250   0.9643   0.02731   0.01680  -0.0746   0.5302   1.0000
   5.500   0.9853   0.02832   0.01782  -0.0728   0.5028   1.0000
   5.750   1.0064   0.02941   0.01893  -0.0711   0.4763   1.0000
   6.000   1.0308   0.03035   0.01977  -0.0698   0.4533   1.0000
   6.250   1.0485   0.03176   0.02135  -0.0681   0.4322   1.0000
   6.500   1.0692   0.03291   0.02256  -0.0667   0.4137   1.0000
   6.750   1.0909   0.03388   0.02360  -0.0653   0.3966   1.0000
   7.000   1.1120   0.03493   0.02472  -0.0640   0.3810   1.0000
   7.250   1.1314   0.03622   0.02615  -0.0626   0.3668   1.0000
   7.500   1.1489   0.03774   0.02788  -0.0611   0.3538   1.0000
   7.750   1.1650   0.03961   0.03000  -0.0597   0.3427   1.0000
   8.000   1.1843   0.04119   0.03177  -0.0584   0.3317   1.0000
   8.250   1.2075   0.04239   0.03301  -0.0573   0.3199   1.0000
   8.500   1.2140   0.04497   0.03602  -0.0552   0.3093   1.0000
   8.750   1.2223   0.04781   0.03917  -0.0533   0.3007   1.0000
   9.000   1.2466   0.04921   0.04064  -0.0523   0.2904   1.0000
   9.250   1.2324   0.05372   0.04566  -0.0494   0.2832   1.0000
   9.500   1.2455   0.05587   0.04802  -0.0478   0.2733   1.0000
   9.750   1.2243   0.06114   0.05362  -0.0452   0.2684   1.0000
  10.000   1.2293   0.06354   0.05619  -0.0432   0.2581   1.0000
  10.250   1.1849   0.07119   0.06399  -0.0415   0.2592   1.0000
  10.500   1.1392   0.07954   0.07228  -0.0413   0.2612   1.0000
  10.750   1.3430   0.05573   0.04765  -0.0368   0.1624   1.0000
  11.000   1.2181   0.07235   0.06552  -0.0333   0.2107   1.0000
  11.250   1.1916   0.07798   0.07121  -0.0325   0.2081   1.0000
  11.500   1.1639   0.08494   0.07819  -0.0334   0.2064   1.0000
  11.750   1.2536   0.07267   0.06579  -0.0242   0.1428   1.0000
  12.000   1.2201   0.07897   0.07225  -0.0241   0.1453   1.0000
  12.250   1.1885   0.08624   0.07959  -0.0260   0.1473   1.0000
  12.500   1.1585   0.09470   0.08809  -0.0293   0.1486   1.0000
<< Back to GOE 432 AIRFOIL (goe432-il)

Polar data table (+)

Polar graphs


<< Back to GOE 432 AIRFOIL (goe432-il)