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CAP 21 (TraCFoil) (cap21c-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: CAP 21 (TraCFoil) (cap21c-il)
Reynolds number: 50,000
Max Cl/Cd: 9.9 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-cap21c-il-50000.txt
Download as CSV file: xf-cap21c-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CAP 21   (TraCFoil)                             
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5023   0.13312   0.12483   0.0354   1.0005   0.3214
  -9.750  -0.4690   0.12766   0.11935   0.0360   1.0005   0.3274
  -9.500  -0.4891   0.12707   0.11881   0.0347   1.0005   0.3368
  -9.250  -0.4537   0.12169   0.11343   0.0354   1.0005   0.3436
  -9.000  -0.4587   0.11959   0.11139   0.0348   1.0005   0.3539
  -8.750  -0.4565   0.11768   0.10951   0.0349   1.0005   0.3668
  -8.500  -0.4283   0.11315   0.10503   0.0353   1.0005   0.3755
  -8.250  -0.4283   0.11070   0.10263   0.0350   1.0005   0.3859
  -8.000  -0.4406   0.10977   0.10176   0.0350   1.0005   0.3986
  -7.750  -0.4022   0.10475   0.09678   0.0353   1.0005   0.4073
  -7.500  -0.3947   0.10205   0.09414   0.0355   1.0005   0.4193
  -7.000  -0.3768   0.09701   0.08924   0.0364   1.0005   0.4486
  -6.500  -0.3594   0.09218   0.08453   0.0374   1.0005   0.4782
  -6.000  -0.3353   0.08717   0.07971   0.0386   1.0005   0.5091
  -5.750  -0.3136   0.08431   0.07695   0.0391   1.0005   0.5269
  -5.500  -0.2948   0.08188   0.07464   0.0399   1.0005   0.5481
  -5.250  -0.2929   0.08031   0.07323   0.0415   1.0005   0.5703
  -5.000  -0.2657   0.07721   0.07030   0.0415   1.0005   0.5886
  -4.750  -0.2438   0.07470   0.06801   0.0417   1.0005   0.6092
  -4.500  -0.2313   0.07274   0.06633   0.0426   1.0005   0.6334
  -4.250  -0.1927   0.07024   0.06393   0.0403   0.7549   0.6647
  -4.000  -0.1854   0.07008   0.06280   0.0447   0.6237   0.6932
  -3.750  -0.3931   0.04886   0.04170   0.0103   1.0005   0.3813
  -3.500  -0.3786   0.04594   0.03895   0.0092   1.0005   0.3782
  -3.250  -0.2735   0.04133   0.03330  -0.0090   0.6796   0.3809
  -3.000  -0.2533   0.03969   0.03083  -0.0098   0.6155   0.3826
  -2.750  -0.2295   0.03840   0.02911  -0.0102   0.5797   0.3843
  -2.500  -0.2034   0.03733   0.02773  -0.0105   0.5553   0.3863
  -2.250  -0.1764   0.03647   0.02668  -0.0109   0.5364   0.3905
  -2.000  -0.1486   0.03551   0.02545  -0.0118   0.5207   0.3953
  -1.750  -0.1199   0.03457   0.02422  -0.0130   0.5069   0.3990
  -1.500  -0.0920   0.03399   0.02356  -0.0131   0.4950   0.4024
  -1.250  -0.0630   0.03347   0.02297  -0.0136   0.4849   0.4062
  -1.000  -0.0338   0.03307   0.02242  -0.0143   0.4762   0.4114
  -0.750  -0.0056   0.03280   0.02208  -0.0144   0.4674   0.4174
  -0.500   0.0225   0.03268   0.02199  -0.0147   0.4588   0.4241
  -0.250   0.0514   0.03260   0.02170  -0.0150   0.4512   0.4304
   0.000   0.0796   0.03267   0.02195  -0.0153   0.4451   0.4362
   0.250   0.1084   0.03283   0.02206  -0.0157   0.4391   0.4435
   0.500   0.1357   0.03310   0.02238  -0.0155   0.4333   0.4520
   0.750   0.1630   0.03350   0.02291  -0.0159   0.4266   0.4622
   1.000   0.1896   0.03388   0.02330  -0.0155   0.4208   0.4728
   1.250   0.2160   0.03456   0.02417  -0.0156   0.4167   0.4838
   1.500   0.2424   0.03538   0.02523  -0.0161   0.4125   0.4963
   1.750   0.2682   0.03611   0.02610  -0.0160   0.4074   0.5123
   2.000   0.2928   0.03692   0.02707  -0.0156   0.4025   0.5339
   2.250   0.3146   0.03815   0.02876  -0.0159   0.3983   0.5599
   2.500   0.3351   0.03944   0.03055  -0.0159   0.3961   0.5992
   2.750   0.3497   0.04039   0.03229  -0.0141   0.3936   0.6786
   3.000   0.4156   0.04196   0.03436  -0.0197   0.3865   0.9995
   3.250   0.4262   0.04471   0.03708  -0.0203   0.3846   0.9995
   3.500   0.4374   0.04771   0.04007  -0.0210   0.3855   0.9995
   3.750   0.4481   0.05093   0.04329  -0.0218   0.3867   0.9995
   4.000   0.4583   0.05427   0.04662  -0.0228   0.3877   0.9995
   4.250   0.2777   0.07521   0.06861  -0.0564   0.6369   0.9995
   4.500   0.2816   0.07650   0.06967  -0.0545   0.6106   0.9995
   4.750   0.3091   0.07947   0.07243  -0.0547   0.5917   0.9995
   5.000   0.3211   0.08180   0.07460  -0.0538   0.5722   0.9995
   5.250   0.3306   0.08366   0.07633  -0.0525   0.5495   0.9995
   5.500   0.3548   0.08649   0.07903  -0.0521   0.5303   0.9995
   5.750   0.3798   0.08966   0.08209  -0.0517   0.5131   0.9995
   6.000   0.3929   0.09222   0.08458  -0.0511   0.4975   0.9995
   6.250   0.4015   0.09454   0.08681  -0.0504   0.4824   0.9995
   6.750   0.4110   0.09896   0.09108  -0.0493   0.4539   0.9995
   7.000   0.4173   0.10141   0.09346  -0.0489   0.4402   0.9995
   7.250   0.4256   0.10400   0.09598  -0.0485   0.4269   0.9995
   7.500   0.4382   0.10696   0.09888  -0.0483   0.4152   0.9995
   7.750   0.4640   0.11057   0.10245  -0.0480   0.4048   0.9995
   8.000   0.4551   0.11225   0.10408  -0.0481   0.3934   0.9995
   8.250   0.4735   0.11612   0.10789  -0.0481   0.3867   0.9995
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