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

GOE 374 AIRFOIL (goe374-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 374 AIRFOIL (goe374-il)
Reynolds number: 500,000
Max Cl/Cd: 97.07 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe374-il-500000-n5.txt
Download as CSV file: xf-goe374-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 374 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3157   0.09211   0.09000  -0.0203   1.0000   0.0065
  -7.500  -0.3165   0.08984   0.08777  -0.0199   1.0000   0.0065
  -7.250  -0.3197   0.08767   0.08565  -0.0193   1.0000   0.0065
  -7.000  -0.3036   0.08374   0.08174  -0.0239   0.9949   0.0065
  -6.750  -0.2894   0.07939   0.07740  -0.0264   0.9883   0.0069
  -6.500  -0.2675   0.07595   0.07394  -0.0315   0.9799   0.0076
  -6.250  -0.2411   0.07173   0.06970  -0.0382   0.9712   0.0080
  -5.750  -0.1651   0.06257   0.06041  -0.0569   0.9497   0.0096
  -4.250  -0.0080   0.03811   0.03511  -0.0775   0.8522   0.0064
  -4.000   0.0176   0.03465   0.03143  -0.0785   0.8360   0.0065
  -3.750   0.0430   0.03120   0.02774  -0.0791   0.8200   0.0066
  -3.500   0.0663   0.02989   0.02628  -0.0792   0.8020   0.0073
  -3.250   0.0915   0.02785   0.02401  -0.0791   0.7825   0.0081
  -3.000   0.1177   0.02509   0.02095  -0.0787   0.7608   0.0085
  -2.750   0.1435   0.02226   0.01773  -0.0779   0.7351   0.0087
  -2.500   0.1691   0.01916   0.01414  -0.0766   0.7058   0.0092
  -2.250   0.1939   0.01685   0.01136  -0.0754   0.6770   0.0101
  -2.000   0.2186   0.01678   0.01111  -0.0749   0.6513   0.0115
  -1.750   0.2447   0.01468   0.00854  -0.0737   0.6334   0.0133
  -1.500   0.2707   0.01333   0.00685  -0.0729   0.6172   0.0153
  -1.250   0.2965   0.01339   0.00682  -0.0726   0.6009   0.0168
  -1.000   0.3227   0.01270   0.00590  -0.0720   0.5851   0.0188
  -0.750   0.3495   0.01249   0.00542  -0.0714   0.5680   0.0215
  -0.500   0.3751   0.01178   0.00454  -0.0709   0.5497   0.0231
  -0.250   0.4009   0.01140   0.00408  -0.0705   0.5283   0.0242
   0.000   0.4265   0.01119   0.00374  -0.0700   0.5047   0.0260
   0.250   0.4522   0.01096   0.00336  -0.0695   0.4830   0.0274
   0.500   0.4776   0.01081   0.00306  -0.0690   0.4622   0.0284
   0.750   0.5033   0.01075   0.00288  -0.0685   0.4435   0.0293
   1.000   0.5291   0.01074   0.00276  -0.0681   0.4280   0.0298
   1.500   0.5801   0.01034   0.00229  -0.0672   0.4077   0.0319
   1.750   0.6062   0.01032   0.00222  -0.0668   0.3978   0.0313
   2.000   0.6322   0.01035   0.00219  -0.0665   0.3864   0.0309
   2.250   0.6581   0.01038   0.00218  -0.0661   0.3747   0.0305
   2.500   0.6844   0.01038   0.00217  -0.0658   0.3661   0.0303
   2.750   0.7105   0.01041   0.00218  -0.0655   0.3579   0.0301
   3.000   0.7366   0.01044   0.00218  -0.0652   0.3472   0.0300
   3.250   0.7627   0.01049   0.00221  -0.0648   0.3364   0.0301
   3.500   0.7887   0.01057   0.00224  -0.0645   0.3263   0.0303
   3.750   0.8146   0.01067   0.00229  -0.0642   0.3162   0.0310
   4.000   0.8406   0.01077   0.00237  -0.0639   0.3061   0.0326
   4.250   0.8665   0.01089   0.00248  -0.0636   0.2960   0.0369
   4.500   0.9084   0.00947   0.00286  -0.0675   0.2831   1.0000
   4.750   0.9335   0.00968   0.00302  -0.0670   0.2720   1.0000
   5.000   0.9585   0.00990   0.00323  -0.0665   0.2603   1.0000
   5.250   0.9833   0.01013   0.00343  -0.0661   0.2477   1.0000
   5.500   1.0069   0.01051   0.00369  -0.0655   0.2227   1.0000
   5.750   1.0301   0.01094   0.00399  -0.0648   0.1957   1.0000
   6.000   1.0520   0.01150   0.00437  -0.0640   0.1601   1.0000
   6.250   1.0736   0.01211   0.00481  -0.0631   0.1292   1.0000
   6.500   1.0961   0.01260   0.00521  -0.0624   0.1112   1.0000
   6.750   1.1177   0.01320   0.00568  -0.0616   0.0873   1.0000
   7.000   1.1307   0.01480   0.00686  -0.0595   0.0146   1.0000
   7.250   1.1532   0.01529   0.00742  -0.0587   0.0099   1.0000
   7.500   1.1755   0.01577   0.00801  -0.0578   0.0085   1.0000
   7.750   1.1965   0.01637   0.00872  -0.0568   0.0072   1.0000
   8.000   1.2171   0.01701   0.00947  -0.0558   0.0060   1.0000
   8.250   1.2375   0.01763   0.01020  -0.0547   0.0054   1.0000
   8.500   1.2568   0.01835   0.01104  -0.0535   0.0049   1.0000
   8.750   1.2745   0.01920   0.01198  -0.0521   0.0045   1.0000
   9.000   1.2878   0.02042   0.01334  -0.0501   0.0042   1.0000
   9.250   1.3028   0.02140   0.01445  -0.0483   0.0040   1.0000
   9.500   1.3156   0.02248   0.01566  -0.0463   0.0038   1.0000
   9.750   1.3282   0.02349   0.01678  -0.0443   0.0035   1.0000
  10.000   1.3371   0.02464   0.01805  -0.0418   0.0033   1.0000
  10.250   1.3447   0.02563   0.01912  -0.0391   0.0031   1.0000
  10.500   1.3507   0.02670   0.02030  -0.0363   0.0029   1.0000
  10.750   1.3514   0.02818   0.02188  -0.0333   0.0027   1.0000
  11.000   1.3502   0.02992   0.02374  -0.0304   0.0027   1.0000
  11.250   1.3451   0.03215   0.02610  -0.0278   0.0026   1.0000
  11.500   1.3413   0.03452   0.02862  -0.0258   0.0025   1.0000
  11.750   1.3385   0.03704   0.03128  -0.0244   0.0025   1.0000
  12.000   1.3366   0.03970   0.03409  -0.0235   0.0024   1.0000
  12.250   1.3333   0.04271   0.03724  -0.0230   0.0024   1.0000
  12.500   1.3282   0.04614   0.04082  -0.0228   0.0024   1.0000
  12.750   1.3232   0.04982   0.04466  -0.0230   0.0023   1.0000
  13.000   1.3175   0.05372   0.04871  -0.0234   0.0022   1.0000
  13.250   1.3119   0.05776   0.05291  -0.0241   0.0021   1.0000
  13.500   1.3045   0.06215   0.05746  -0.0249   0.0021   1.0000
  13.750   1.2969   0.06672   0.06219  -0.0261   0.0021   1.0000
  14.000   1.2889   0.07147   0.06709  -0.0275   0.0021   1.0000
  14.250   1.2798   0.07660   0.07238  -0.0293   0.0021   1.0000
  14.500   1.2700   0.08204   0.07798  -0.0314   0.0021   1.0000
  14.750   1.2597   0.08779   0.08388  -0.0338   0.0021   1.0000
  15.000   1.2472   0.09412   0.09037  -0.0366   0.0020   1.0000
  15.250   1.2362   0.10037   0.09678  -0.0396   0.0021   1.0000
  15.500   1.2245   0.10703   0.10358  -0.0430   0.0020   1.0000
  15.750   1.2123   0.11395   0.11064  -0.0466   0.0021   1.0000
<< Back to GOE 374 AIRFOIL (goe374-il)

Polar data table (+)

Polar graphs


<< Back to GOE 374 AIRFOIL (goe374-il)