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

GOE 623 AIRFOIL (goe623-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 623 AIRFOIL (goe623-il)
Reynolds number: 100,000
Max Cl/Cd: 53.55 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe623-il-100000.txt
Download as CSV file: xf-goe623-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 623 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2555   0.10284   0.09820  -0.0373   1.0000   0.1237
  -9.500  -0.2846   0.10126   0.09674  -0.0401   1.0000   0.1264
  -9.250  -0.3173   0.09978   0.09541  -0.0419   1.0000   0.1270
  -9.000  -0.2681   0.09354   0.08909  -0.0370   1.0000   0.1304
  -8.750  -0.2656   0.09116   0.08676  -0.0352   1.0000   0.1344
  -8.500  -0.3510   0.09428   0.08953  -0.0360   1.0000   0.1309
  -8.250  -0.3482   0.09174   0.08705  -0.0351   1.0000   0.1345
  -8.000  -0.3626   0.08983   0.08526  -0.0345   1.0000   0.1384
  -7.750  -0.3964   0.08908   0.08469  -0.0320   1.0000   0.1400
  -7.500  -0.4433   0.08820   0.08395  -0.0330   1.0000   0.1414
  -7.250  -0.4674   0.08503   0.08081  -0.0336   1.0000   0.1425
  -7.000  -0.4508   0.08218   0.07805  -0.0274   1.0000   0.1444
  -6.750  -0.4477   0.08047   0.07639  -0.0235   1.0000   0.1478
  -6.500  -0.4569   0.07829   0.07425  -0.0226   1.0000   0.1521
  -6.250  -0.4828   0.07398   0.06978  -0.0294   1.0000   0.1584
  -6.000  -0.4663   0.07134   0.06727  -0.0259   0.9983   0.1612
  -5.750  -0.4340   0.06523   0.06085  -0.0374   0.9894   0.1748
  -5.500  -0.3997   0.04418   0.03840  -0.0520   0.9838   0.0961
  -5.250  -0.3658   0.03892   0.03255  -0.0558   0.9764   0.0937
  -5.000  -0.3249   0.03364   0.02641  -0.0599   0.9714   0.0907
  -4.750  -0.2882   0.03071   0.02291  -0.0622   0.9634   0.0911
  -4.500  -0.2435   0.02881   0.02058  -0.0658   0.9577   0.0953
  -4.250  -0.2065   0.02706   0.01836  -0.0675   0.9488   0.0979
  -4.000  -0.1609   0.02539   0.01651  -0.0709   0.9433   0.1020
  -3.750  -0.1248   0.02463   0.01567  -0.0726   0.9335   0.1098
  -3.500  -0.0784   0.02357   0.01459  -0.0760   0.9281   0.1205
  -3.250  -0.0437   0.02277   0.01380  -0.0773   0.9182   0.1368
  -3.000   0.0021   0.02148   0.01271  -0.0806   0.9130   0.1914
  -2.750   0.0356   0.02083   0.01228  -0.0819   0.9032   0.2411
  -2.500   0.0775   0.02029   0.01189  -0.0846   0.8970   0.2894
  -2.250   0.1071   0.01996   0.01174  -0.0850   0.8865   0.3366
  -2.000   0.1454   0.01912   0.01130  -0.0867   0.8805   0.4083
  -1.750   0.1659   0.01803   0.01131  -0.0848   0.8698   0.6223
  -1.500   0.2619   0.01710   0.01090  -0.0951   0.8698   1.0000
  -1.250   0.2958   0.01700   0.01053  -0.0960   0.8619   1.0000
  -1.000   0.3191   0.01712   0.01047  -0.0951   0.8502   1.0000
  -0.750   0.3419   0.01730   0.01049  -0.0942   0.8391   1.0000
  -0.500   0.3732   0.01722   0.01023  -0.0944   0.8318   1.0000
  -0.250   0.3933   0.01750   0.01038  -0.0930   0.8199   1.0000
   0.000   0.4184   0.01763   0.01038  -0.0923   0.8106   1.0000
   0.250   0.4452   0.01765   0.01028  -0.0917   0.8016   1.0000
   0.500   0.4675   0.01786   0.01039  -0.0905   0.7906   1.0000
   0.750   0.4982   0.01768   0.01007  -0.0902   0.7835   1.0000
   1.000   0.5195   0.01792   0.01025  -0.0890   0.7714   1.0000
   1.250   0.5438   0.01805   0.01030  -0.0880   0.7612   1.0000
   1.500   0.5729   0.01790   0.01004  -0.0875   0.7529   1.0000
   1.750   0.5957   0.01807   0.01017  -0.0864   0.7410   1.0000
   2.000   0.6212   0.01808   0.01012  -0.0855   0.7303   1.0000
   2.250   0.6506   0.01786   0.00979  -0.0849   0.7212   1.0000
   2.500   0.6741   0.01799   0.00989  -0.0839   0.7088   1.0000
   2.750   0.6991   0.01805   0.00992  -0.0830   0.6973   1.0000
   3.000   0.7266   0.01800   0.00980  -0.0823   0.6870   1.0000
   3.250   0.7535   0.01800   0.00975  -0.0816   0.6759   1.0000
   3.500   0.7780   0.01814   0.00988  -0.0808   0.6635   1.0000
   3.750   0.8036   0.01824   0.00995  -0.0800   0.6513   1.0000
   4.000   0.8302   0.01831   0.00998  -0.0793   0.6392   1.0000
   4.250   0.8577   0.01834   0.00993  -0.0787   0.6269   1.0000
   4.500   0.8828   0.01848   0.01006  -0.0779   0.6127   1.0000
   4.750   0.9075   0.01864   0.01022  -0.0770   0.5978   1.0000
   5.000   0.9323   0.01879   0.01036  -0.0761   0.5820   1.0000
   5.250   0.9572   0.01893   0.01045  -0.0751   0.5652   1.0000
   5.500   0.9808   0.01909   0.01060  -0.0740   0.5467   1.0000
   5.750   1.0029   0.01929   0.01084  -0.0728   0.5261   1.0000
   6.000   1.0260   0.01947   0.01098  -0.0716   0.5051   1.0000
   6.250   1.0484   0.01971   0.01115  -0.0704   0.4830   1.0000
   6.500   1.0695   0.02001   0.01143  -0.0690   0.4590   1.0000
   6.750   1.0908   0.02037   0.01171  -0.0677   0.4363   1.0000
   7.000   1.1115   0.02084   0.01212  -0.0664   0.4143   1.0000
   7.250   1.1335   0.02140   0.01256  -0.0653   0.3959   1.0000
   7.500   1.1560   0.02206   0.01314  -0.0645   0.3800   1.0000
   7.750   1.1786   0.02281   0.01386  -0.0637   0.3659   1.0000
   8.000   1.2012   0.02361   0.01467  -0.0630   0.3535   1.0000
   8.250   1.2246   0.02445   0.01553  -0.0624   0.3429   1.0000
   8.500   1.2507   0.02530   0.01630  -0.0623   0.3338   1.0000
   8.750   1.2718   0.02623   0.01742  -0.0614   0.3252   1.0000
   9.000   1.2977   0.02712   0.01825  -0.0613   0.3169   1.0000
   9.250   1.3166   0.02792   0.01924  -0.0601   0.3073   1.0000
   9.500   1.3360   0.02870   0.02010  -0.0591   0.2971   1.0000
   9.750   1.3549   0.02930   0.02067  -0.0579   0.2851   1.0000
  10.000   1.3720   0.02980   0.02112  -0.0565   0.2722   1.0000
  10.250   1.3818   0.03034   0.02190  -0.0540   0.2605   1.0000
  10.500   1.3929   0.03098   0.02268  -0.0518   0.2495   1.0000
  10.750   1.4043   0.03157   0.02334  -0.0496   0.2388   1.0000
  11.000   1.4135   0.03214   0.02395  -0.0472   0.2278   1.0000
  11.250   1.4136   0.03285   0.02493  -0.0435   0.2171   1.0000
  11.500   1.4104   0.03364   0.02588  -0.0396   0.2059   1.0000
  11.750   1.4036   0.03471   0.02710  -0.0358   0.1926   1.0000
  12.000   1.3930   0.03628   0.02884  -0.0324   0.1756   1.0000
  12.250   1.3784   0.03868   0.03131  -0.0298   0.1519   1.0000
  12.500   1.3618   0.04209   0.03465  -0.0279   0.1194   1.0000
  12.750   1.3462   0.04598   0.03834  -0.0265   0.0984   1.0000
  13.000   1.3328   0.04995   0.04221  -0.0256   0.0872   1.0000
  13.250   1.3191   0.05412   0.04630  -0.0250   0.0808   1.0000
  13.500   1.3107   0.05791   0.05016  -0.0245   0.0750   1.0000
  13.750   1.3003   0.06192   0.05416  -0.0242   0.0712   1.0000
  14.000   1.2953   0.06537   0.05766  -0.0236   0.0677   1.0000
  14.250   1.2933   0.06856   0.06093  -0.0231   0.0645   1.0000
  14.500   1.2934   0.07136   0.06369  -0.0223   0.0615   1.0000
  14.750   1.3004   0.07338   0.06569  -0.0207   0.0584   1.0000
  15.000   1.3019   0.07641   0.06891  -0.0203   0.0561   1.0000
  15.250   1.3065   0.07901   0.07160  -0.0197   0.0540   1.0000
  15.500   1.3199   0.08062   0.07318  -0.0181   0.0519   1.0000
  15.750   1.3431   0.08201   0.07455  -0.0155   0.0499   1.0000
  16.000   1.3354   0.08626   0.07908  -0.0161   0.0493   1.0000
  16.250   1.3259   0.09089   0.08398  -0.0171   0.0488   1.0000
  16.500   1.3141   0.09595   0.08931  -0.0185   0.0484   1.0000
  16.750   1.3002   0.10152   0.09514  -0.0205   0.0480   1.0000
  17.000   1.2842   0.10764   0.10151  -0.0230   0.0478   1.0000
  17.250   1.2655   0.11449   0.10862  -0.0264   0.0477   1.0000
  17.500   1.2439   0.12222   0.11659  -0.0306   0.0478   1.0000
  17.750   1.2198   0.13098   0.12558  -0.0359   0.0483   1.0000
  18.000   1.1929   0.14099   0.13580  -0.0423   0.0489   1.0000
  18.250   1.1652   0.15199   0.14698  -0.0494   0.0496   1.0000
<< Back to GOE 623 AIRFOIL (goe623-il)

Polar data table (+)

Polar graphs


<< Back to GOE 623 AIRFOIL (goe623-il)