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

GOE 591 AIRFOIL (goe591-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 591 AIRFOIL (goe591-il)
Reynolds number: 500,000
Max Cl/Cd: 107.98 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe591-il-500000.txt
Download as CSV file: xf-goe591-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 591 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2410   0.08423   0.08209  -0.0619   0.9862   0.0279
  -8.500  -0.2259   0.07858   0.07644  -0.0693   0.9810   0.0280
  -8.250  -0.2156   0.07414   0.07201  -0.0720   0.9754   0.0288
  -8.000  -0.1926   0.07148   0.06933  -0.0755   0.9717   0.0295
  -7.750  -0.1750   0.06780   0.06565  -0.0808   0.9628   0.0305
  -7.500  -0.1478   0.06188   0.05969  -0.0918   0.9567   0.0320
  -7.250  -0.1392   0.04528   0.04275  -0.1128   0.9338   0.0350
  -7.000  -0.1190   0.04395   0.04138  -0.1137   0.9232   0.0357
  -6.750  -0.0977   0.04291   0.04028  -0.1142   0.9116   0.0367
  -6.500  -0.0803   0.04072   0.03797  -0.1150   0.8979   0.0385
  -6.250  -0.0793   0.03174   0.02833  -0.1169   0.8817   0.0437
  -6.000  -0.0589   0.03081   0.02737  -0.1164   0.8704   0.0446
  -5.750  -0.0385   0.02958   0.02602  -0.1160   0.8602   0.0461
  -5.500  -0.0404   0.01829   0.01299  -0.1121   0.8492   0.0345
  -5.250  -0.0209   0.01571   0.01002  -0.1110   0.8399   0.0351
  -5.000   0.0028   0.01441   0.00854  -0.1103   0.8317   0.0361
  -4.750   0.0282   0.01384   0.00787  -0.1098   0.8231   0.0373
  -4.500   0.0536   0.01321   0.00711  -0.1093   0.8149   0.0381
  -4.250   0.0794   0.01263   0.00641  -0.1087   0.8069   0.0389
  -4.000   0.1051   0.01211   0.00579  -0.1082   0.7987   0.0397
  -3.750   0.1312   0.01173   0.00530  -0.1076   0.7907   0.0409
  -3.500   0.1570   0.01137   0.00487  -0.1071   0.7824   0.0418
  -3.250   0.1832   0.01104   0.00443  -0.1066   0.7746   0.0423
  -3.000   0.2078   0.01047   0.00381  -0.1058   0.7658   0.0435
  -2.750   0.2334   0.01013   0.00341  -0.1053   0.7578   0.0451
  -2.500   0.2591   0.00990   0.00316  -0.1047   0.7485   0.0471
  -2.250   0.2854   0.00975   0.00296  -0.1042   0.7402   0.0497
  -2.000   0.3114   0.00958   0.00274  -0.1037   0.7307   0.0522
  -1.750   0.3373   0.00938   0.00255  -0.1031   0.7213   0.0582
  -1.250   0.3889   0.00905   0.00228  -0.1021   0.7017   0.1011
  -1.000   0.4146   0.00893   0.00220  -0.1016   0.6916   0.1278
  -0.750   0.4400   0.00881   0.00217  -0.1011   0.6812   0.1714
  -0.500   0.4654   0.00868   0.00214  -0.1006   0.6699   0.2185
  -0.250   0.4905   0.00854   0.00211  -0.1000   0.6589   0.2708
   0.000   0.5133   0.00817   0.00213  -0.0992   0.6478   0.4241
   0.250   0.5242   0.00712   0.00223  -0.0956   0.6369   0.8089
   0.500   0.6117   0.00708   0.00233  -0.1080   0.6220   0.9925
   0.750   0.6525   0.00718   0.00231  -0.1109   0.6076   1.0000
   1.000   0.6748   0.00730   0.00232  -0.1097   0.5939   1.0000
   1.250   0.6971   0.00742   0.00235  -0.1085   0.5805   1.0000
   1.500   0.7196   0.00755   0.00239  -0.1073   0.5686   1.0000
   1.750   0.7420   0.00769   0.00244  -0.1062   0.5576   1.0000
   2.000   0.7647   0.00782   0.00250  -0.1051   0.5464   1.0000
   2.250   0.7877   0.00795   0.00257  -0.1041   0.5360   1.0000
   2.500   0.8105   0.00811   0.00266  -0.1031   0.5266   1.0000
   2.750   0.8335   0.00825   0.00275  -0.1021   0.5168   1.0000
   3.000   0.8565   0.00840   0.00285  -0.1011   0.5065   1.0000
   3.250   0.8790   0.00858   0.00295  -0.1000   0.4964   1.0000
   3.500   0.9018   0.00873   0.00307  -0.0989   0.4862   1.0000
   3.750   0.9251   0.00888   0.00319  -0.0980   0.4773   1.0000
   4.000   0.9479   0.00907   0.00332  -0.0970   0.4691   1.0000
   4.250   0.9716   0.00921   0.00345  -0.0962   0.4612   1.0000
   4.750   1.0180   0.00955   0.00375  -0.0944   0.4443   1.0000
   5.000   1.0407   0.00976   0.00392  -0.0935   0.4364   1.0000
   5.250   1.0642   0.00991   0.00409  -0.0927   0.4279   1.0000
   5.500   1.0869   0.01012   0.00427  -0.0917   0.4198   1.0000
   5.750   1.1100   0.01029   0.00445  -0.0909   0.4110   1.0000
   6.000   1.1327   0.01049   0.00464  -0.0900   0.4028   1.0000
   6.250   1.1542   0.01073   0.00485  -0.0888   0.3924   1.0000
   6.500   1.1764   0.01092   0.00505  -0.0878   0.3805   1.0000
   6.750   1.1977   0.01115   0.00526  -0.0867   0.3677   1.0000
   7.000   1.2184   0.01141   0.00550  -0.0855   0.3544   1.0000
   7.250   1.2376   0.01171   0.00575  -0.0840   0.3369   1.0000
   7.500   1.2565   0.01204   0.00603  -0.0825   0.3206   1.0000
   7.750   1.2747   0.01239   0.00634  -0.0808   0.3050   1.0000
   8.000   1.2910   0.01279   0.00668  -0.0789   0.2885   1.0000
   8.250   1.3055   0.01322   0.00705  -0.0766   0.2709   1.0000
   8.500   1.3194   0.01368   0.00745  -0.0742   0.2536   1.0000
   8.750   1.3321   0.01423   0.00792  -0.0718   0.2349   1.0000
   9.000   1.3433   0.01485   0.00845  -0.0691   0.2160   1.0000
   9.250   1.3554   0.01546   0.00900  -0.0668   0.1969   1.0000
   9.500   1.3641   0.01627   0.00967  -0.0639   0.1714   1.0000
   9.750   1.3687   0.01731   0.01052  -0.0607   0.1401   1.0000
  10.000   1.3707   0.01854   0.01157  -0.0572   0.1132   1.0000
  10.250   1.3689   0.02004   0.01286  -0.0535   0.0767   1.0000
  10.500   1.3581   0.02217   0.01473  -0.0490   0.0405   1.0000
  10.750   1.3560   0.02392   0.01638  -0.0459   0.0239   1.0000
  11.000   1.3615   0.02527   0.01779  -0.0438   0.0205   1.0000
  11.250   1.3673   0.02665   0.01924  -0.0418   0.0192   1.0000
  11.500   1.3714   0.02824   0.02091  -0.0399   0.0180   1.0000
  11.750   1.3753   0.02990   0.02267  -0.0382   0.0171   1.0000
  12.000   1.3804   0.03153   0.02440  -0.0367   0.0168   1.0000
  12.250   1.3834   0.03342   0.02638  -0.0353   0.0160   1.0000
  12.500   1.3848   0.03551   0.02859  -0.0340   0.0156   1.0000
  12.750   1.3844   0.03787   0.03104  -0.0328   0.0152   1.0000
  13.000   1.3824   0.04049   0.03376  -0.0317   0.0148   1.0000
  13.250   1.3784   0.04342   0.03679  -0.0309   0.0145   1.0000
  13.500   1.3701   0.04698   0.04046  -0.0303   0.0142   1.0000
  13.750   1.3582   0.05111   0.04472  -0.0299   0.0138   1.0000
  14.000   1.3490   0.05510   0.04882  -0.0298   0.0137   1.0000
  14.250   1.3476   0.05830   0.05212  -0.0299   0.0135   1.0000
  14.500   1.3446   0.06177   0.05569  -0.0301   0.0134   1.0000
  14.750   1.3422   0.06525   0.05927  -0.0304   0.0132   1.0000
  15.000   1.3384   0.06898   0.06309  -0.0309   0.0130   1.0000
  15.250   1.3360   0.07259   0.06680  -0.0314   0.0129   1.0000
  15.500   1.3325   0.07640   0.07070  -0.0321   0.0125   1.0000
  15.750   1.3297   0.08016   0.07455  -0.0328   0.0123   1.0000
  16.000   1.3268   0.08399   0.07846  -0.0336   0.0120   1.0000
<< Back to GOE 591 AIRFOIL (goe591-il)

Polar data table (+)

Polar graphs


<< Back to GOE 591 AIRFOIL (goe591-il)