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

GOE 210 (DAIMLER) AIRFOIL (goe210-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 210 (DAIMLER) AIRFOIL (goe210-il)
Reynolds number: 50,000
Max Cl/Cd: 43.81 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe210-il-50000-n5.txt
Download as CSV file: xf-goe210-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 210 (DAIMLER) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3527   0.10955   0.10283  -0.0255   1.0000   0.0712
  -8.000  -0.3567   0.10844   0.10185  -0.0276   1.0000   0.0717
  -7.750  -0.3561   0.10688   0.10040  -0.0314   1.0000   0.0720
  -7.500  -0.3445   0.10033   0.09390  -0.0270   1.0000   0.0738
  -7.250  -0.3374   0.09668   0.09029  -0.0258   1.0000   0.0764
  -7.000  -0.3340   0.09399   0.08769  -0.0260   1.0000   0.0795
  -6.750  -0.3321   0.09174   0.08554  -0.0276   1.0000   0.0831
  -6.500  -0.3296   0.09048   0.08436  -0.0332   1.0000   0.0857
  -6.000  -0.3219   0.08340   0.07746  -0.0302   1.0000   0.0909
  -5.750  -0.3174   0.08085   0.07493  -0.0306   1.0000   0.0956
  -5.500  -0.3044   0.07967   0.07366  -0.0393   1.0000   0.1006
  -5.250  -0.3033   0.07556   0.06969  -0.0357   1.0000   0.1024
  -5.000  -0.2994   0.07261   0.06680  -0.0338   1.0000   0.1054
  -4.750  -0.2905   0.06997   0.06416  -0.0347   1.0000   0.1102
  -4.500  -0.2722   0.06715   0.06120  -0.0399   1.0000   0.1163
  -4.250  -0.2649   0.06400   0.05811  -0.0384   1.0000   0.1191
  -4.000  -0.2325   0.06073   0.05460  -0.0452   0.9973   0.1305
  -3.750  -0.2018   0.05667   0.05046  -0.0490   0.9914   0.1463
  -3.500  -0.1416   0.05050   0.04369  -0.0578   0.9867   0.0833
  -3.250  -0.0978   0.04564   0.03840  -0.0631   0.9810   0.0660
  -3.000  -0.0477   0.04113   0.03321  -0.0690   0.9765   0.0599
  -2.750  -0.0127   0.03854   0.03037  -0.0720   0.9703   0.0642
  -2.500   0.0309   0.03558   0.02686  -0.0759   0.9660   0.0651
  -2.250   0.0691   0.03311   0.02385  -0.0782   0.9597   0.0652
  -2.000   0.1110   0.03131   0.02130  -0.0809   0.9547   0.0713
  -1.750   0.1480   0.02950   0.01903  -0.0827   0.9489   0.0728
  -1.500   0.1837   0.02801   0.01726  -0.0844   0.9429   0.0751
  -1.250   0.2218   0.02706   0.01599  -0.0864   0.9377   0.0831
  -1.000   0.2534   0.02625   0.01487  -0.0869   0.9300   0.0873
  -0.750   0.2923   0.02538   0.01380  -0.0888   0.9252   0.0907
  -0.500   0.3231   0.02487   0.01311  -0.0893   0.9171   0.0950
  -0.250   0.3615   0.02436   0.01239  -0.0912   0.9114   0.1017
   0.000   0.3929   0.02400   0.01197  -0.0921   0.9029   0.1130
   0.250   0.4329   0.02331   0.01153  -0.0946   0.8970   0.1622
   0.500   0.4574   0.02124   0.01131  -0.0938   0.8879   1.0000
   0.750   0.4919   0.02141   0.01114  -0.0948   0.8793   1.0000
   1.000   0.5257   0.02154   0.01104  -0.0957   0.8696   1.0000
   1.250   0.5565   0.02167   0.01101  -0.0960   0.8579   1.0000
   1.500   0.5882   0.02174   0.01096  -0.0963   0.8458   1.0000
   1.750   0.6198   0.02178   0.01091  -0.0965   0.8334   1.0000
   2.000   0.6513   0.02176   0.01082  -0.0964   0.8198   1.0000
   2.250   0.6824   0.02162   0.01063  -0.0960   0.8036   1.0000
   2.500   0.7138   0.02135   0.01031  -0.0953   0.7850   1.0000
   2.750   0.7408   0.02121   0.01014  -0.0939   0.7638   1.0000
   3.000   0.7675   0.02112   0.01002  -0.0927   0.7431   1.0000
   3.250   0.7947   0.02109   0.01001  -0.0916   0.7249   1.0000
   3.500   0.8181   0.02125   0.01023  -0.0903   0.7047   1.0000
   3.750   0.8430   0.02133   0.01035  -0.0891   0.6840   1.0000
   4.000   0.8667   0.02145   0.01057  -0.0878   0.6610   1.0000
   4.250   0.8907   0.02154   0.01071  -0.0864   0.6355   1.0000
   4.500   0.9135   0.02168   0.01091  -0.0849   0.6049   1.0000
   4.750   0.9361   0.02182   0.01108  -0.0833   0.5693   1.0000
   5.000   0.9587   0.02203   0.01129  -0.0817   0.5285   1.0000
   5.250   0.9809   0.02239   0.01151  -0.0800   0.4849   1.0000
   5.500   1.0021   0.02297   0.01190  -0.0785   0.4435   1.0000
   5.750   1.0223   0.02375   0.01247  -0.0769   0.4061   1.0000
   6.000   1.0421   0.02466   0.01326  -0.0756   0.3740   1.0000
   6.250   1.0619   0.02561   0.01411  -0.0744   0.3473   1.0000
   6.500   1.0825   0.02654   0.01505  -0.0733   0.3250   1.0000
   6.750   1.1033   0.02748   0.01601  -0.0723   0.3063   1.0000
   7.000   1.1242   0.02842   0.01706  -0.0714   0.2895   1.0000
   7.250   1.1446   0.02936   0.01807  -0.0704   0.2736   1.0000
   7.500   1.1646   0.03030   0.01914  -0.0693   0.2579   1.0000
   7.750   1.1840   0.03126   0.02026  -0.0682   0.2429   1.0000
   8.000   1.2039   0.03227   0.02143  -0.0672   0.2299   1.0000
   8.250   1.2231   0.03332   0.02270  -0.0661   0.2168   1.0000
   8.500   1.2378   0.03438   0.02400  -0.0645   0.2000   1.0000
   8.750   1.2487   0.03545   0.02526  -0.0627   0.1804   1.0000
   9.000   1.2596   0.03661   0.02659  -0.0608   0.1629   1.0000
   9.250   1.2707   0.03784   0.02805  -0.0590   0.1469   1.0000
   9.500   1.2801   0.03915   0.02959  -0.0571   0.1299   1.0000
   9.750   1.2847   0.04046   0.03105  -0.0550   0.1017   1.0000
  10.000   1.2865   0.04243   0.03295  -0.0528   0.0820   1.0000
  10.250   1.2839   0.04514   0.03555  -0.0506   0.0681   1.0000
  10.500   1.2805   0.04821   0.03870  -0.0486   0.0577   1.0000
  10.750   1.2744   0.05161   0.04219  -0.0469   0.0508   1.0000
  11.000   1.2704   0.05499   0.04582  -0.0457   0.0445   1.0000
  11.250   1.2628   0.05877   0.04969  -0.0450   0.0413   1.0000
  11.500   1.2583   0.06250   0.05383  -0.0443   0.0382   1.0000
  11.750   1.2516   0.06658   0.05814  -0.0442   0.0363   1.0000
  12.000   1.2438   0.07093   0.06267  -0.0445   0.0349   1.0000
  12.250   1.2350   0.07555   0.06745  -0.0453   0.0340   1.0000
  12.500   1.2265   0.08042   0.07251  -0.0462   0.0330   1.0000
  12.750   1.2176   0.08561   0.07797  -0.0474   0.0322   1.0000
  13.000   1.2075   0.09122   0.08383  -0.0491   0.0315   1.0000
  13.250   1.1963   0.09725   0.09010  -0.0513   0.0310   1.0000
  13.500   1.1844   0.10374   0.09681  -0.0541   0.0306   1.0000
  13.750   1.1714   0.11076   0.10405  -0.0573   0.0305   1.0000
  14.000   1.1573   0.11847   0.11196  -0.0612   0.0306   1.0000
  14.250   1.1420   0.12696   0.12064  -0.0656   0.0309   1.0000
  14.500   1.1260   0.13613   0.12994  -0.0706   0.0314   1.0000
<< Back to GOE 210 (DAIMLER) AIRFOIL (goe210-il)

Polar data table (+)

Polar graphs


<< Back to GOE 210 (DAIMLER) AIRFOIL (goe210-il)