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GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)
Reynolds number: 200,000
Max Cl/Cd: 71.77 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe417a-il-200000.txt
Download as CSV file: xf-goe417a-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 417A (GEW. PLATTE) AIRFOIL                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.2932   0.09229   0.08908  -0.0125   1.0000   0.0232
  -7.250  -0.2971   0.09041   0.08723  -0.0110   1.0000   0.0235
  -7.000  -0.3029   0.08868   0.08554  -0.0093   1.0000   0.0237
  -6.750  -0.3103   0.08711   0.08402  -0.0072   1.0000   0.0238
  -6.500  -0.3143   0.08529   0.08222  -0.0060   1.0000   0.0240
  -6.250  -0.3156   0.08325   0.08022  -0.0053   1.0000   0.0243
  -6.000  -0.3158   0.08120   0.07820  -0.0048   1.0000   0.0245
  -5.750  -0.3132   0.07898   0.07600  -0.0051   0.9997   0.0248
  -5.500  -0.2800   0.07499   0.07195  -0.0130   0.9946   0.0255
  -5.250  -0.2413   0.07249   0.06936  -0.0235   0.9882   0.0257
  -5.000  -0.3202   0.08166   0.07839  -0.0108   1.0000   0.0253
  -4.500  -0.2269   0.07577   0.07229  -0.0309   0.9910   0.0259
  -4.250  -0.2090   0.07107   0.06759  -0.0322   0.9868   0.0262
  -4.000  -0.1782   0.06727   0.06377  -0.0368   0.9833   0.0266
  -3.750  -0.1492   0.06418   0.06062  -0.0407   0.9770   0.0270
  -3.500  -0.1120   0.06105   0.05744  -0.0464   0.9725   0.0277
  -3.250  -0.0734   0.05811   0.05443  -0.0519   0.9684   0.0286
  -3.000  -0.0382   0.05553   0.05177  -0.0563   0.9616   0.0295
  -2.750   0.0364   0.05529   0.05119  -0.0676   0.9577   0.0306
  -2.500   0.0626   0.05150   0.04740  -0.0698   0.9529   0.0308
  -2.250   0.0822   0.04794   0.04386  -0.0707   0.9464   0.0312
  -2.000   0.1168   0.04506   0.04095  -0.0742   0.9429   0.0318
  -1.750   0.1496   0.04283   0.03865  -0.0768   0.9371   0.0328
  -1.500   0.1840   0.04078   0.03652  -0.0794   0.9307   0.0338
  -1.250   0.2285   0.03891   0.03451  -0.0834   0.9276   0.0356
  -1.000   0.2793   0.03920   0.03445  -0.0864   0.9206   0.0368
  -0.750   0.3001   0.03517   0.03053  -0.0875   0.9150   0.0377
  -0.500   0.3379   0.03282   0.02815  -0.0905   0.9118   0.0394
  -0.250   0.3676   0.03125   0.02649  -0.0910   0.9027   0.0411
   0.000   0.4183   0.03048   0.02548  -0.0941   0.8982   0.0441
   0.250   0.4643   0.02832   0.02315  -0.0973   0.8956   0.0452
   0.500   0.4845   0.02616   0.02103  -0.0965   0.8847   0.0465
   0.750   0.5260   0.02433   0.01914  -0.0992   0.8814   0.0495
   1.000   0.5636   0.02486   0.01936  -0.0988   0.8706   0.0545
   1.250   0.5996   0.02164   0.01618  -0.1012   0.8655   0.0567
   1.500   0.6254   0.02034   0.01486  -0.1006   0.8523   0.0609
   1.750   0.6569   0.01948   0.01383  -0.1003   0.8392   0.0683
   2.000   0.6926   0.01964   0.01376  -0.1007   0.8239   0.0784
   2.250   0.7234   0.01720   0.01139  -0.1018   0.8077   0.0832
   2.500   0.7610   0.01627   0.01030  -0.1032   0.7885   0.0945
   2.750   0.7939   0.01551   0.00940  -0.1039   0.7644   0.1099
   3.000   0.8259   0.01482   0.00854  -0.1046   0.7384   0.1376
   3.250   0.8537   0.01425   0.00780  -0.1046   0.7116   0.1766
   3.500   0.8787   0.01389   0.00736  -0.1037   0.6867   0.1988
   3.750   0.9042   0.01375   0.00706  -0.1026   0.6618   0.2057
   4.000   0.9277   0.01366   0.00683  -0.1010   0.6326   0.2033
   4.250   0.9518   0.01372   0.00666  -0.0991   0.6014   0.1536
   4.500   0.9737   0.01404   0.00668  -0.0966   0.5698   0.0991
   4.750   0.9942   0.01429   0.00668  -0.0944   0.5392   0.0789
   5.000   1.0144   0.01423   0.00654  -0.0926   0.5018   0.0744
   5.250   1.0320   0.01438   0.00655  -0.0903   0.4563   0.0712
   5.500   1.0467   0.01471   0.00662  -0.0875   0.3846   0.0705
   5.750   1.0595   0.01560   0.00699  -0.0848   0.2910   0.0749
   6.000   1.0767   0.01641   0.00745  -0.0831   0.2419   0.0779
   6.250   1.0967   0.01701   0.00786  -0.0819   0.2092   0.0814
   6.500   1.1140   0.01770   0.00828  -0.0799   0.1599   0.0888
   6.750   1.1218   0.01920   0.00903  -0.0764   0.0677   0.1027
   7.000   1.2257   0.02103   0.01193  -0.0946   0.0503   1.0000
   7.250   1.2368   0.02215   0.01308  -0.0915   0.0468   1.0000
   7.500   1.2439   0.02380   0.01471  -0.0877   0.0445   1.0000
   7.750   1.2589   0.02479   0.01576  -0.0852   0.0432   1.0000
   8.000   1.2743   0.02594   0.01695  -0.0829   0.0417   1.0000
   8.250   1.2909   0.02717   0.01821  -0.0808   0.0399   1.0000
   8.500   1.3079   0.02844   0.01951  -0.0791   0.0378   1.0000
   8.750   1.3292   0.03029   0.02134  -0.0781   0.0364   1.0000
   9.000   1.3589   0.03336   0.02445  -0.0788   0.0353   1.0000
   9.250   1.3872   0.03709   0.02837  -0.0791   0.0349   1.0000
   9.500   1.4013   0.03869   0.03021  -0.0766   0.0345   1.0000
   9.750   1.4122   0.03990   0.03173  -0.0736   0.0338   1.0000
  10.000   1.4230   0.04169   0.03381  -0.0707   0.0331   1.0000
  10.250   1.4330   0.04447   0.03689  -0.0679   0.0332   1.0000
  10.500   1.4395   0.04766   0.04040  -0.0648   0.0334   1.0000
  10.750   1.4573   0.05357   0.04649  -0.0642   0.0352   1.0000
  11.000   1.3841   0.04562   0.03916  -0.0498   0.0360   1.0000
  11.250   1.3377   0.04923   0.04357  -0.0387   0.0403   1.0000
  14.000   1.0618   0.15539   0.15185  -0.0701   0.0747   1.0000
  14.250   1.0638   0.15905   0.15550  -0.0710   0.0716   1.0000
  14.500   1.0756   0.15992   0.15640  -0.0688   0.0696   1.0000
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