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

GOE 523 AIRFOIL (goe523-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 523 AIRFOIL (goe523-il)
Reynolds number: 200,000
Max Cl/Cd: 74.29 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe523-il-200000.txt
Download as CSV file: xf-goe523-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 523 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000   0.1678   0.11355   0.10977  -0.1222   0.9016   0.0566
  -9.750   0.1832   0.11033   0.10654  -0.1258   0.8987   0.0588
  -9.500   0.1810   0.10734   0.10355  -0.1335   0.8957   0.0605
  -9.250   0.2181   0.10274   0.09892  -0.1349   0.8952   0.0613
  -9.000   0.2190   0.10150   0.09771  -0.1316   0.8859   0.0623
  -8.750   0.2389   0.09850   0.09469  -0.1344   0.8830   0.0643
  -8.500   0.2405   0.09578   0.09196  -0.1421   0.8796   0.0681
  -8.000   0.2676   0.08950   0.08569  -0.1421   0.8684   0.0695
  -7.750   0.2996   0.08599   0.08215  -0.1451   0.8668   0.0712
  -7.500   0.3266   0.08250   0.07862  -0.1498   0.8648   0.0740
  -7.250   0.3052   0.08145   0.07760  -0.1498   0.8531   0.0771
  -7.000   0.3716   0.07488   0.07094  -0.1613   0.8597   0.0795
  -6.750   0.3649   0.07405   0.07015  -0.1564   0.8484   0.0805
  -6.500   0.3962   0.07079   0.06684  -0.1615   0.8462   0.0839
  -6.250   0.3739   0.07021   0.06630  -0.1562   0.8332   0.0860
  -6.000   0.3844   0.06567   0.06172  -0.1642   0.8284   0.0891
  -5.500   0.4115   0.06237   0.05841  -0.1609   0.8138   0.0922
  -5.250   0.3963   0.05942   0.05545  -0.1692   0.7990   0.0995
  -5.000   0.4283   0.05566   0.05164  -0.1712   0.7963   0.1012
  -4.750   0.5817   0.02617   0.02023  -0.2412   0.7940   0.0681
  -4.500   0.6066   0.02439   0.01824  -0.2422   0.7839   0.0676
  -4.250   0.5833   0.03283   0.02801  -0.2240   0.7774   0.1017
  -4.000   0.7068   0.02071   0.01380  -0.2525   0.7739   0.0691
  -3.750   0.7349   0.01989   0.01292  -0.2528   0.7648   0.0706
  -3.500   0.7806   0.01886   0.01170  -0.2563   0.7596   0.0723
  -3.250   0.8036   0.01844   0.01122  -0.2554   0.7496   0.0738
  -3.000   0.8439   0.01776   0.01036  -0.2577   0.7433   0.0764
  -2.750   0.8702   0.01732   0.00991  -0.2575   0.7339   0.0792
  -2.500   0.9070   0.01688   0.00942  -0.2592   0.7265   0.0850
  -2.250   0.9353   0.01659   0.00913  -0.2594   0.7172   0.0914
  -2.000   0.9721   0.01622   0.00874  -0.2612   0.7096   0.1043
  -1.750   1.0033   0.01596   0.00857  -0.2621   0.7007   0.1445
  -1.500   1.0401   0.01561   0.00834  -0.2642   0.6924   0.2181
  -1.250   1.0676   0.01563   0.00844  -0.2643   0.6833   0.2684
  -1.000   1.0990   0.01565   0.00843  -0.2650   0.6750   0.3173
  -0.750   1.1253   0.01578   0.00862  -0.2648   0.6663   0.3636
  -0.500   1.1537   0.01594   0.00877  -0.2648   0.6578   0.4093
  -0.250   1.1784   0.01618   0.00902  -0.2641   0.6491   0.4422
   0.000   1.2045   0.01643   0.00921  -0.2636   0.6404   0.4708
   0.250   1.2298   0.01672   0.00944  -0.2630   0.6321   0.4935
   0.500   1.2534   0.01698   0.00966  -0.2621   0.6234   0.5123
   0.750   1.2794   0.01729   0.00990  -0.2616   0.6154   0.5303
   1.000   1.2999   0.01757   0.01018  -0.2601   0.6065   0.5454
   1.250   1.3280   0.01788   0.01037  -0.2601   0.5989   0.5612
   1.500   1.3457   0.01818   0.01069  -0.2581   0.5901   0.5762
   1.750   1.3743   0.01850   0.01087  -0.2582   0.5827   0.5931
   2.000   1.3887   0.01882   0.01126  -0.2556   0.5741   0.6058
   2.250   1.4121   0.01908   0.01145  -0.2547   0.5666   0.6185
   2.500   1.4317   0.01940   0.01175  -0.2532   0.5588   0.6321
   2.750   1.4501   0.01968   0.01201  -0.2515   0.5510   0.6466
   3.000   1.4741   0.01995   0.01221  -0.2508   0.5442   0.6583
   3.250   1.4864   0.02022   0.01252  -0.2480   0.5363   0.6680
   3.500   1.5108   0.02047   0.01266  -0.2475   0.5296   0.6792
   3.750   1.5260   0.02079   0.01302  -0.2453   0.5225   0.6881
   4.000   1.5438   0.02110   0.01330  -0.2436   0.5156   0.6987
   4.250   1.5702   0.02137   0.01347  -0.2436   0.5096   0.7086
   4.500   1.5824   0.02177   0.01395  -0.2410   0.5027   0.7172
   4.750   1.6027   0.02209   0.01423  -0.2399   0.4965   0.7270
   5.000   1.6270   0.02243   0.01451  -0.2395   0.4908   0.7366
   5.250   1.6417   0.02290   0.01504  -0.2375   0.4845   0.7464
   5.500   1.6626   0.02324   0.01535  -0.2366   0.4789   0.7560
   5.750   1.6876   0.02362   0.01567  -0.2365   0.4734   0.7671
   6.250   1.7187   0.02454   0.01666  -0.2329   0.4618   0.7932
   6.500   1.7451   0.02488   0.01693  -0.2330   0.4566   0.8076
   6.750   1.7550   0.02547   0.01765  -0.2303   0.4510   0.8216
   7.000   1.7713   0.02590   0.01816  -0.2288   0.4458   0.8426
   7.250   1.7959   0.02610   0.01832  -0.2285   0.4410   1.0000
   7.500   1.8094   0.02687   0.01916  -0.2267   0.4359   1.0000
   7.750   1.8241   0.02760   0.01993  -0.2251   0.4306   1.0000
   8.000   1.8460   0.02817   0.02045  -0.2246   0.4257   1.0000
   8.250   1.8710   0.02876   0.02098  -0.2247   0.4209   1.0000
   8.500   1.8782   0.02971   0.02204  -0.2220   0.4158   1.0000
   8.750   1.8930   0.03046   0.02279  -0.2205   0.4106   1.0000
   9.000   1.9229   0.03087   0.02306  -0.2212   0.4054   1.0000
   9.250   1.9238   0.03207   0.02442  -0.2177   0.4003   1.0000
   9.500   1.9329   0.03304   0.02545  -0.2155   0.3950   1.0000
   9.750   1.9517   0.03371   0.02607  -0.2146   0.3903   1.0000
  10.000   1.9700   0.03452   0.02689  -0.2138   0.3857   1.0000
  10.250   1.9728   0.03585   0.02836  -0.2109   0.3809   1.0000
  10.500   1.9841   0.03685   0.02940  -0.2091   0.3763   1.0000
  10.750   2.0060   0.03743   0.02990  -0.2087   0.3719   1.0000
  11.000   2.0123   0.03877   0.03135  -0.2065   0.3673   1.0000
  11.250   2.0139   0.04030   0.03301  -0.2037   0.3625   1.0000
  11.500   2.0243   0.04140   0.03415  -0.2020   0.3580   1.0000
  11.750   2.0504   0.04180   0.03444  -0.2022   0.3538   1.0000
  12.000   2.0457   0.04383   0.03668  -0.1990   0.3496   1.0000
  12.250   2.0456   0.04564   0.03861  -0.1964   0.3448   1.0000
  12.500   2.0537   0.04693   0.03992  -0.1948   0.3403   1.0000
  12.750   2.0707   0.04777   0.04071  -0.1940   0.3357   1.0000
  13.000   2.0616   0.05043   0.04359  -0.1910   0.3312   1.0000
  13.250   2.0631   0.05240   0.04566  -0.1890   0.3267   1.0000
  13.500   2.0738   0.05367   0.04692  -0.1879   0.3227   1.0000
  13.750   2.0842   0.05508   0.04837  -0.1868   0.3186   1.0000
  14.000   2.0761   0.05811   0.05160  -0.1845   0.3145   1.0000
  14.250   2.0763   0.06046   0.05406  -0.1829   0.3102   1.0000
  14.500   2.0861   0.06192   0.05552  -0.1819   0.3064   1.0000
  14.750   2.0940   0.06364   0.05728  -0.1809   0.3024   1.0000
  15.000   2.0819   0.06746   0.06132  -0.1790   0.2982   1.0000
  15.250   2.0800   0.07029   0.06427  -0.1778   0.2940   1.0000
  15.500   2.0887   0.07191   0.06588  -0.1770   0.2899   1.0000
  15.750   2.0897   0.07454   0.06861  -0.1761   0.2859   1.0000
  16.000   2.0758   0.07904   0.07333  -0.1750   0.2816   1.0000
  16.250   2.0725   0.08230   0.07670  -0.1742   0.2774   1.0000
  16.500   2.0850   0.08350   0.07785  -0.1738   0.2732   1.0000
  16.750   2.0757   0.08769   0.08222  -0.1732   0.2694   1.0000
  17.000   2.0619   0.09258   0.08733  -0.1728   0.2654   1.0000
  17.250   2.0579   0.09616   0.09101  -0.1726   0.2613   1.0000
  17.500   2.0755   0.09656   0.09131  -0.1723   0.2566   1.0000
  17.750   2.0501   0.10336   0.09840  -0.1725   0.2527   1.0000
  18.000   2.0345   0.10877   0.10400  -0.1730   0.2479   1.0000
  18.250   2.0429   0.11049   0.10566  -0.1729   0.2423   1.0000
  18.500   2.0247   0.11645   0.11183  -0.1737   0.2379   1.0000
  18.750   2.0062   0.12256   0.11813  -0.1749   0.2331   1.0000
  19.000   2.0136   0.12443   0.11997  -0.1751   0.2273   1.0000
<< Back to GOE 523 AIRFOIL (goe523-il)

Polar data table (+)

Polar graphs


<< Back to GOE 523 AIRFOIL (goe523-il)