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

GOE 371 AIRFOIL (goe371-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 371 AIRFOIL (goe371-il)
Reynolds number: 100,000
Max Cl/Cd: 61.09 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe371-il-100000.txt
Download as CSV file: xf-goe371-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 371 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3348   0.09989   0.09509  -0.0292   1.0000   0.0602
  -7.750  -0.3434   0.09900   0.09433  -0.0295   1.0000   0.0606
  -7.500  -0.3483   0.09807   0.09351  -0.0331   1.0000   0.0609
  -7.250  -0.3489   0.09662   0.09211  -0.0371   1.0000   0.0612
  -7.000  -0.3397   0.08909   0.08462  -0.0271   1.0000   0.0635
  -6.750  -0.3381   0.08653   0.08212  -0.0253   1.0000   0.0652
  -6.500  -0.3398   0.08440   0.08006  -0.0243   1.0000   0.0671
  -6.250  -0.3431   0.08240   0.07813  -0.0238   1.0000   0.0694
  -6.000  -0.3460   0.08057   0.07635  -0.0242   1.0000   0.0716
  -5.750  -0.3416   0.07955   0.07530  -0.0311   1.0000   0.0742
  -5.500  -0.3338   0.07679   0.07246  -0.0357   1.0000   0.0752
  -5.250  -0.3395   0.07292   0.06874  -0.0294   1.0000   0.0766
  -5.000  -0.3386   0.07042   0.06629  -0.0263   1.0000   0.0788
  -4.750  -0.3316   0.06799   0.06386  -0.0263   1.0000   0.0830
  -4.500  -0.3040   0.06448   0.06012  -0.0354   1.0000   0.0902
  -4.250  -0.3038   0.06177   0.05751  -0.0312   1.0000   0.0933
  -4.000  -0.2688   0.05905   0.05440  -0.0397   1.0000   0.1042
  -3.750  -0.2657   0.05560   0.05113  -0.0365   1.0000   0.1066
  -3.500  -0.2206   0.05175   0.04702  -0.0441   0.9950   0.1200
  -3.250  -0.1816   0.04829   0.04338  -0.0493   0.9891   0.1346
  -3.000  -0.1464   0.04511   0.04016  -0.0528   0.9838   0.1520
  -2.750  -0.1073   0.04246   0.03725  -0.0575   0.9769   0.1786
  -2.500  -0.0682   0.03981   0.03448  -0.0614   0.9712   0.2078
  -1.500   0.1185   0.02780   0.02025  -0.0760   0.9460   0.1210
  -1.250   0.1656   0.02583   0.01781  -0.0789   0.9407   0.1096
  -1.000   0.2027   0.02466   0.01614  -0.0798   0.9316   0.1020
  -0.750   0.2467   0.02324   0.01448  -0.0824   0.9251   0.0990
  -0.500   0.2848   0.02222   0.01331  -0.0838   0.9161   0.0983
  -0.250   0.3242   0.02142   0.01240  -0.0855   0.9073   0.1007
   0.000   0.3703   0.02043   0.01146  -0.0885   0.9001   0.1075
   0.250   0.4080   0.01975   0.01083  -0.0899   0.8898   0.1136
   0.500   0.4583   0.01887   0.00997  -0.0933   0.8840   0.1294
   0.750   0.5111   0.01625   0.00935  -0.0977   0.8750   1.0000
   1.000   0.5493   0.01607   0.00893  -0.0989   0.8635   1.0000
   1.250   0.5873   0.01584   0.00855  -0.1000   0.8519   1.0000
   1.500   0.6232   0.01561   0.00822  -0.1006   0.8396   1.0000
   1.750   0.6537   0.01550   0.00802  -0.1002   0.8242   1.0000
   2.000   0.6822   0.01545   0.00790  -0.0995   0.8075   1.0000
   2.250   0.7103   0.01540   0.00778  -0.0987   0.7899   1.0000
   2.500   0.7385   0.01528   0.00759  -0.0977   0.7715   1.0000
   2.750   0.7661   0.01516   0.00736  -0.0965   0.7525   1.0000
   3.000   0.7900   0.01517   0.00730  -0.0949   0.7297   1.0000
   3.250   0.8161   0.01518   0.00721  -0.0938   0.7106   1.0000
   3.500   0.8408   0.01533   0.00736  -0.0927   0.6919   1.0000
   3.750   0.8653   0.01551   0.00753  -0.0917   0.6738   1.0000
   4.000   0.8902   0.01567   0.00767  -0.0907   0.6565   1.0000
   4.250   0.9153   0.01582   0.00780  -0.0897   0.6395   1.0000
   4.500   0.9405   0.01596   0.00794  -0.0888   0.6227   1.0000
   4.750   0.9644   0.01618   0.00819  -0.0877   0.6042   1.0000
   5.000   0.9887   0.01639   0.00841  -0.0867   0.5858   1.0000
   5.250   1.0134   0.01663   0.00863  -0.0857   0.5669   1.0000
   5.500   1.0367   0.01697   0.00898  -0.0845   0.5457   1.0000
   5.750   1.0605   0.01738   0.00933  -0.0834   0.5237   1.0000
   6.000   1.0823   0.01785   0.00979  -0.0820   0.4974   1.0000
   6.250   1.1027   0.01832   0.01020  -0.0803   0.4668   1.0000
   6.500   1.1220   0.01880   0.01062  -0.0784   0.4342   1.0000
   6.750   1.1399   0.01928   0.01113  -0.0765   0.3994   1.0000
   7.000   1.1584   0.01983   0.01167  -0.0747   0.3666   1.0000
   7.250   1.1763   0.02040   0.01219  -0.0728   0.3354   1.0000
   7.500   1.1930   0.02101   0.01274  -0.0709   0.3047   1.0000
   7.750   1.2083   0.02173   0.01339  -0.0690   0.2727   1.0000
   8.000   1.2228   0.02264   0.01420  -0.0670   0.2434   1.0000
   8.250   1.2367   0.02367   0.01516  -0.0651   0.2163   1.0000
   8.500   1.2488   0.02471   0.01620  -0.0629   0.1869   1.0000
   8.750   1.2578   0.02588   0.01733  -0.0604   0.1514   1.0000
   9.000   1.2593   0.02780   0.01895  -0.0570   0.0984   1.0000
   9.250   1.2584   0.03025   0.02114  -0.0531   0.0747   1.0000
   9.500   1.2621   0.03233   0.02320  -0.0497   0.0644   1.0000
   9.750   1.2688   0.03448   0.02534  -0.0471   0.0579   1.0000
  10.000   1.2801   0.03641   0.02734  -0.0450   0.0529   1.0000
  10.250   1.3035   0.03976   0.03059  -0.0448   0.0489   1.0000
  10.500   1.3235   0.04226   0.03342  -0.0436   0.0470   1.0000
  10.750   1.3387   0.04489   0.03635  -0.0421   0.0448   1.0000
  11.000   1.3483   0.04733   0.03906  -0.0402   0.0426   1.0000
  11.250   1.3565   0.04993   0.04183  -0.0384   0.0409   1.0000
  11.500   1.3651   0.05349   0.04560  -0.0370   0.0399   1.0000
  11.750   1.3659   0.05719   0.04962  -0.0348   0.0396   1.0000
  12.000   1.3599   0.06072   0.05347  -0.0322   0.0395   1.0000
  12.250   1.3489   0.06434   0.05741  -0.0297   0.0396   1.0000
  12.500   1.3339   0.06796   0.06135  -0.0277   0.0397   1.0000
  12.750   1.3155   0.07198   0.06569  -0.0264   0.0400   1.0000
  13.000   1.2940   0.07668   0.07070  -0.0261   0.0403   1.0000
  13.250   1.2681   0.08223   0.07657  -0.0271   0.0409   1.0000
  13.500   1.2343   0.08928   0.08394  -0.0301   0.0415   1.0000
  13.750   1.1986   0.09789   0.09283  -0.0351   0.0422   1.0000
  14.000   1.1607   0.10855   0.10372  -0.0423   0.0433   1.0000
<< Back to GOE 371 AIRFOIL (goe371-il)

Polar data table (+)

Polar graphs


<< Back to GOE 371 AIRFOIL (goe371-il)