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

GOE 400 AIRFOIL (goe400-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 400 AIRFOIL (goe400-il)
Reynolds number: 50,000
Max Cl/Cd: 37.42 at α=8.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe400-il-50000.txt
Download as CSV file: xf-goe400-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 400 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4158   0.12414   0.11717   0.0029   1.0000   0.1149
  -9.000  -0.4175   0.12379   0.11691  -0.0003   1.0000   0.1166
  -8.750  -0.4233   0.12450   0.11774  -0.0044   1.0000   0.1172
  -8.500  -0.3965   0.11436   0.10754  -0.0010   1.0000   0.1207
  -8.250  -0.3878   0.11095   0.10415  -0.0017   1.0000   0.1243
  -8.000  -0.3838   0.10883   0.10208  -0.0037   1.0000   0.1284
  -7.750  -0.3873   0.10920   0.10259  -0.0087   1.0000   0.1306
  -7.500  -0.3727   0.10281   0.09620  -0.0067   1.0000   0.1340
  -7.250  -0.3626   0.09930   0.09272  -0.0070   1.0000   0.1391
  -7.000  -0.3565   0.09780   0.09129  -0.0119   1.0000   0.1440
  -6.750  -0.3475   0.09516   0.08872  -0.0164   1.0000   0.1464
  -6.500  -0.3352   0.09015   0.08374  -0.0137   1.0000   0.1507
  -6.250  -0.3238   0.08748   0.08111  -0.0165   1.0000   0.1570
  -6.000  -0.3108   0.08533   0.07900  -0.0229   1.0000   0.1612
  -5.750  -0.3005   0.08097   0.07470  -0.0199   1.0000   0.1675
  -5.500  -0.2848   0.07916   0.07290  -0.0268   1.0000   0.1760
  -5.250  -0.2761   0.07514   0.06895  -0.0236   1.0000   0.1827
  -5.000  -0.2609   0.07297   0.06681  -0.0286   1.0000   0.1912
  -4.750  -0.2530   0.06969   0.06362  -0.0264   1.0000   0.1978
  -4.500  -0.2407   0.06742   0.06138  -0.0292   1.0000   0.2067
  -4.250  -0.2269   0.06610   0.06003  -0.0325   1.0000   0.2192
  -4.000  -0.2282   0.06280   0.05688  -0.0281   1.0000   0.2237
  -3.750  -0.2211   0.06103   0.05512  -0.0288   1.0000   0.2361
  -3.500  -0.2149   0.05916   0.05326  -0.0287   1.0000   0.2504
  -3.250  -0.2089   0.05710   0.05123  -0.0279   1.0000   0.2661
  -3.000  -0.2015   0.05503   0.04919  -0.0271   1.0000   0.2840
  -2.750  -0.1894   0.05307   0.04722  -0.0279   1.0000   0.3098
  -1.750  -0.1606   0.04409   0.03850  -0.0196   1.0000   0.4774
  -1.500  -0.1435   0.04127   0.03577  -0.0181   0.9960   0.5373
  -1.250  -0.1085   0.03817   0.03268  -0.0198   0.9858   0.6067
  -1.000  -0.0593   0.03526   0.02973  -0.0255   0.9745   0.6405
  -0.750   0.1746   0.03541   0.02711  -0.0779   0.9573   0.2277
  -0.500   0.2307   0.03368   0.02483  -0.0837   0.9450   0.2149
  -0.250   0.2841   0.03205   0.02270  -0.0886   0.9328   0.2089
   0.000   0.3349   0.03080   0.02110  -0.0930   0.9205   0.2233
   0.250   0.3882   0.02953   0.01954  -0.0975   0.9088   0.2516
   0.500   0.4372   0.02864   0.01843  -0.1011   0.8960   0.2960
   0.750   0.4780   0.02812   0.01785  -0.1031   0.8821   0.3291
   1.000   0.5160   0.02784   0.01755  -0.1046   0.8681   0.3540
   1.250   0.5510   0.02766   0.01745  -0.1056   0.8539   0.3753
   1.500   0.5835   0.02754   0.01751  -0.1061   0.8397   0.4047
   1.750   0.6138   0.02632   0.01763  -0.1060   0.8258   1.0000
   2.000   0.6453   0.02694   0.01785  -0.1059   0.8113   1.0000
   2.250   0.6742   0.02755   0.01818  -0.1054   0.7967   1.0000
   2.500   0.7013   0.02819   0.01864  -0.1047   0.7823   1.0000
   2.750   0.7273   0.02886   0.01918  -0.1039   0.7677   1.0000
   3.000   0.7526   0.02956   0.01984  -0.1030   0.7533   1.0000
   3.250   0.7774   0.03028   0.02051  -0.1020   0.7389   1.0000
   3.500   0.8017   0.03102   0.02123  -0.1010   0.7245   1.0000
   3.750   0.8258   0.03176   0.02198  -0.0998   0.7102   1.0000
   4.000   0.8497   0.03251   0.02277  -0.0986   0.6960   1.0000
   4.250   0.8736   0.03322   0.02352  -0.0973   0.6818   1.0000
   4.500   0.8979   0.03387   0.02420  -0.0959   0.6678   1.0000
   4.750   0.9207   0.03465   0.02505  -0.0945   0.6531   1.0000
   5.000   0.9429   0.03550   0.02602  -0.0932   0.6381   1.0000
   5.250   0.9650   0.03634   0.02695  -0.0917   0.6231   1.0000
   5.500   0.9873   0.03711   0.02783  -0.0902   0.6079   1.0000
   5.750   1.0068   0.03811   0.02897  -0.0887   0.5908   1.0000
   6.000   1.0263   0.03893   0.02992  -0.0870   0.5722   1.0000
   6.250   1.0511   0.03865   0.02978  -0.0843   0.5531   1.0000
   6.500   1.0814   0.03722   0.02840  -0.0808   0.5341   1.0000
   6.750   1.1021   0.03735   0.02868  -0.0785   0.5119   1.0000
   7.000   1.1282   0.03686   0.02834  -0.0759   0.4922   1.0000
   7.250   1.1545   0.03633   0.02791  -0.0734   0.4714   1.0000
   7.500   1.1787   0.03595   0.02764  -0.0709   0.4467   1.0000
   7.750   1.2019   0.03579   0.02760  -0.0684   0.4194   1.0000
   8.000   1.2271   0.03529   0.02713  -0.0656   0.3873   1.0000
   8.250   1.2475   0.03466   0.02636  -0.0623   0.3448   1.0000
   8.500   1.2611   0.03370   0.02504  -0.0589   0.2947   1.0000
   8.750   1.2688   0.03452   0.02566  -0.0559   0.2468   1.0000
   9.000   1.2730   0.03636   0.02709  -0.0529   0.1990   1.0000
   9.250   1.2768   0.03933   0.02965  -0.0501   0.1576   1.0000
   9.500   1.2885   0.04243   0.03265  -0.0480   0.1327   1.0000
   9.750   1.3043   0.04531   0.03544  -0.0464   0.1172   1.0000
  10.000   1.3236   0.04888   0.03922  -0.0451   0.1085   1.0000
  10.250   1.3457   0.05252   0.04272  -0.0443   0.1008   1.0000
  10.500   1.3479   0.05617   0.04703  -0.0423   0.0973   1.0000
  10.750   1.3510   0.05999   0.05127  -0.0406   0.0939   1.0000
  11.000   1.3537   0.06408   0.05568  -0.0391   0.0919   1.0000
  11.250   1.3505   0.06848   0.06042  -0.0376   0.0909   1.0000
  11.500   1.3354   0.07301   0.06532  -0.0359   0.0910   1.0000
  11.750   1.3060   0.07777   0.07044  -0.0347   0.0920   1.0000
  12.000   1.2707   0.08378   0.07675  -0.0359   0.0935   1.0000
  12.250   1.2342   0.09130   0.08451  -0.0395   0.0953   1.0000
  12.500   1.2001   0.10006   0.09342  -0.0444   0.0974   1.0000
  12.750   1.1734   0.10904   0.10248  -0.0492   0.0993   1.0000
  13.000   1.1536   0.11777   0.11123  -0.0535   0.1005   1.0000
  13.250   1.0540   0.15125   0.14448  -0.0790   0.1247   1.0000
<< Back to GOE 400 AIRFOIL (goe400-il)

Polar data table (+)

Polar graphs


<< Back to GOE 400 AIRFOIL (goe400-il)