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GOE 646 AIRFOIL (goe646-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 646 AIRFOIL (goe646-il)
Reynolds number: 50,000
Max Cl/Cd: 18.51 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe646-il-50000-n5.txt
Download as CSV file: xf-goe646-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 646 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.3464   0.12477   0.11677  -0.0490   1.0000   0.1104
 -11.750  -0.3630   0.12025   0.11231  -0.0494   1.0000   0.1116
 -11.500  -0.3812   0.11562   0.10774  -0.0498   1.0000   0.1131
 -11.250  -0.3770   0.11500   0.10718  -0.0477   1.0000   0.1148
 -11.000  -0.3810   0.11310   0.10534  -0.0464   1.0000   0.1169
 -10.750  -0.3936   0.10986   0.10215  -0.0459   1.0000   0.1188
 -10.500  -0.4166   0.10500   0.09735  -0.0462   1.0000   0.1209
 -10.250  -0.4812   0.09347   0.08592  -0.0501   1.0000   0.1239
 -10.000  -0.4492   0.09526   0.08772  -0.0490   0.9970   0.1263
  -9.750  -0.4405   0.09108   0.08350  -0.0531   0.9909   0.1303
  -9.500  -0.6224   0.06086   0.05291  -0.0742   0.9721   0.1337
  -9.250  -0.6382   0.05505   0.04678  -0.0781   0.9607   0.1373
  -9.000  -0.6071   0.05465   0.04642  -0.0796   0.9543   0.1421
  -8.750  -0.6137   0.05002   0.04136  -0.0816   0.9430   0.1472
  -8.500  -0.5881   0.04888   0.04020  -0.0830   0.9360   0.1524
  -8.250  -0.5719   0.04728   0.03847  -0.0835   0.9266   0.1582
  -8.000  -0.5512   0.04525   0.03621  -0.0852   0.9191   0.1650
  -7.750  -0.5303   0.04456   0.03551  -0.0850   0.9100   0.1708
  -7.500  -0.5023   0.04275   0.03344  -0.0873   0.9041   0.1794
  -7.250  -0.4858   0.04234   0.03307  -0.0860   0.8935   0.1851
  -7.000  -0.4577   0.04085   0.03132  -0.0877   0.8870   0.1945
  -6.750  -0.4370   0.04059   0.03111  -0.0869   0.8778   0.2013
  -6.500  -0.4120   0.03949   0.02980  -0.0876   0.8700   0.2109
  -6.250  -0.3763   0.03894   0.02919  -0.0894   0.8651   0.2216
  -6.000  -0.3643   0.03855   0.02874  -0.0873   0.8534   0.2290
  -5.750  -0.3328   0.03786   0.02788  -0.0885   0.8475   0.2409
  -5.500  -0.3124   0.03781   0.02786  -0.0874   0.8384   0.2489
  -5.250  -0.2876   0.03712   0.02693  -0.0877   0.8304   0.2611
  -5.000  -0.2517   0.03702   0.02688  -0.0887   0.8258   0.2715
  -4.750  -0.2401   0.03677   0.02647  -0.0867   0.8146   0.2812
  -4.500  -0.2096   0.03662   0.02630  -0.0871   0.8087   0.2924
  -4.250  -0.1730   0.03627   0.02587  -0.0885   0.8049   0.3049
  -4.000  -0.1658   0.03633   0.02581  -0.0857   0.7926   0.3147
  -3.750  -0.1329   0.03612   0.02560  -0.0863   0.7879   0.3259
  -3.500  -0.1170   0.03599   0.02523  -0.0851   0.7789   0.3375
  -3.250  -0.0930   0.03601   0.02535  -0.0843   0.7721   0.3457
  -3.000  -0.0591   0.03549   0.02462  -0.0854   0.7681   0.3587
  -2.750  -0.0478   0.03579   0.02498  -0.0830   0.7583   0.3652
  -2.500  -0.0209   0.03554   0.02457  -0.0831   0.7523   0.3767
  -2.250   0.0131   0.03518   0.02421  -0.0838   0.7487   0.3862
  -2.000   0.0213   0.03564   0.02460  -0.0814   0.7391   0.3946
  -1.750   0.0481   0.03551   0.02442  -0.0813   0.7335   0.4043
  -1.500   0.0828   0.03515   0.02398  -0.0822   0.7300   0.4159
  -1.250   0.0911   0.03572   0.02453  -0.0798   0.7206   0.4236
  -1.000   0.1156   0.03576   0.02456  -0.0793   0.7150   0.4332
  -0.750   0.1487   0.03546   0.02419  -0.0800   0.7115   0.4449
  -0.500   0.1580   0.03615   0.02491  -0.0777   0.7028   0.4530
  -0.250   0.1816   0.03630   0.02499  -0.0773   0.6967   0.4641
   0.000   0.2149   0.03605   0.02476  -0.0777   0.6930   0.4759
   0.500   0.2494   0.03702   0.02573  -0.0753   0.6781   0.4968
   0.750   0.2844   0.03667   0.02537  -0.0759   0.6742   0.5109
   1.250   0.3231   0.03734   0.02610  -0.0736   0.6582   0.5362
   1.500   0.3622   0.03671   0.02548  -0.0744   0.6544   0.5543
   2.000   0.3993   0.03744   0.02630  -0.0719   0.6378   0.5879
   2.250   0.4393   0.03676   0.02566  -0.0727   0.6343   0.6129
   2.750   0.4703   0.03777   0.02685  -0.0695   0.6174   0.6592
   3.000   0.5110   0.03692   0.02609  -0.0702   0.6143   0.6962
   3.500   0.5392   0.03780   0.02738  -0.0664   0.5969   0.7787
   3.750   0.6038   0.03634   0.02618  -0.0709   0.5939   1.0000
   4.250   0.6355   0.03795   0.02756  -0.0689   0.5756   1.0000
   4.750   0.6671   0.03955   0.02900  -0.0666   0.5571   1.0000
   5.000   0.7148   0.03861   0.02794  -0.0681   0.5540   1.0000
   5.250   0.6985   0.04117   0.03052  -0.0642   0.5386   1.0000
   5.500   0.7436   0.04023   0.02947  -0.0652   0.5350   1.0000
   6.000   0.7713   0.04199   0.03117  -0.0624   0.5159   1.0000
   6.500   0.7991   0.04377   0.03290  -0.0597   0.4969   1.0000
   7.000   0.8248   0.04578   0.03489  -0.0570   0.4779   1.0000
   7.500   0.8497   0.04787   0.03696  -0.0544   0.4592   1.0000
   8.000   0.8733   0.05011   0.03920  -0.0519   0.4407   1.0000
   8.500   0.8939   0.05273   0.04183  -0.0496   0.4227   1.0000
   9.000   0.9109   0.05588   0.04502  -0.0474   0.4051   1.0000
   9.500   0.9244   0.05960   0.04877  -0.0455   0.3880   1.0000
  10.000   0.9324   0.06421   0.05346  -0.0440   0.3713   1.0000
  10.250   0.9707   0.06269   0.05193  -0.0433   0.3687   1.0000
  10.750   0.9675   0.06884   0.05817  -0.0421   0.3519   1.0000
  11.250   0.9582   0.07623   0.06568  -0.0417   0.3349   1.0000
  11.500   0.9777   0.07688   0.06635  -0.0410   0.3306   1.0000
  12.000   0.9682   0.08490   0.07450  -0.0412   0.3154   1.0000
  12.500   0.9393   0.09630   0.08603  -0.0430   0.2979   1.0000
  12.750   0.9694   0.09526   0.08504  -0.0418   0.2961   1.0000
  13.250   0.9309   0.10900   0.09893  -0.0451   0.2791   1.0000
  13.500   0.8858   0.12011   0.11010  -0.0489   0.2674   1.0000
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