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USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 41 AIRFOIL (usa41-il)
Reynolds number: 50,000
Max Cl/Cd: 41.15 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa41-il-50000.txt
Download as CSV file: xf-usa41-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 41 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3416   0.10157   0.09480  -0.0209   1.0000   0.1614
  -7.500  -0.3436   0.09979   0.09312  -0.0207   1.0000   0.1675
  -7.250  -0.3584   0.10033   0.09385  -0.0224   1.0000   0.1701
  -7.000  -0.3418   0.09451   0.08803  -0.0193   1.0000   0.1769
  -6.750  -0.3473   0.09347   0.08712  -0.0208   1.0000   0.1833
  -6.500  -0.3439   0.09021   0.08394  -0.0205   1.0000   0.1871
  -6.250  -0.3388   0.08730   0.08109  -0.0195   1.0000   0.1945
  -6.000  -0.3418   0.08589   0.07980  -0.0228   1.0000   0.1994
  -5.750  -0.3339   0.08203   0.07599  -0.0194   1.0000   0.2059
  -5.500  -0.3335   0.08031   0.07435  -0.0229   1.0000   0.2133
  -5.250  -0.3276   0.07689   0.07098  -0.0197   1.0000   0.2206
  -5.000  -0.3234   0.07432   0.06846  -0.0212   1.0000   0.2289
  -4.750  -0.3165   0.07194   0.06610  -0.0232   1.0000   0.2412
  -4.500  -0.3096   0.06920   0.06339  -0.0230   1.0000   0.2553
  -4.000  -0.2969   0.06350   0.05777  -0.0198   1.0000   0.2873
  -3.750  -0.2888   0.06091   0.05520  -0.0187   1.0000   0.3070
  -3.500  -0.2804   0.05832   0.05265  -0.0179   1.0000   0.3316
  -3.250  -0.2734   0.05589   0.05027  -0.0159   1.0000   0.3612
  -3.000  -0.2683   0.05363   0.04807  -0.0127   1.0000   0.3955
  -2.250  -0.2754   0.04766   0.04240   0.0057   1.0000   0.5474
  -2.000  -0.0567   0.03745   0.02970  -0.0552   1.0000   0.1899
  -1.750  -0.0201   0.03457   0.02619  -0.0579   1.0000   0.1649
  -1.500   0.0071   0.03283   0.02407  -0.0588   1.0000   0.1630
  -1.250   0.0329   0.03137   0.02223  -0.0594   1.0000   0.1623
  -1.000   0.0585   0.03007   0.02048  -0.0597   1.0000   0.1598
  -0.750   0.0817   0.02913   0.01924  -0.0597   1.0000   0.1603
  -0.500   0.1026   0.02852   0.01839  -0.0595   1.0000   0.1633
  -0.250   0.1229   0.02814   0.01773  -0.0592   1.0000   0.1683
   0.000   0.1514   0.02794   0.01738  -0.0606   0.9966   0.1815
   0.250   0.2152   0.02723   0.01640  -0.0675   0.9827   0.2311
   0.500   0.2774   0.02374   0.01512  -0.0738   0.9710   1.0000
   0.750   0.3304   0.02460   0.01536  -0.0792   0.9505   1.0000
   1.000   0.3849   0.02536   0.01572  -0.0847   0.9309   1.0000
   1.250   0.4350   0.02599   0.01609  -0.0892   0.9117   1.0000
   1.500   0.4787   0.02652   0.01646  -0.0924   0.8914   1.0000
   1.750   0.5270   0.02690   0.01672  -0.0962   0.8732   1.0000
   2.000   0.5777   0.02710   0.01684  -0.0999   0.8562   1.0000
   2.250   0.6131   0.02750   0.01722  -0.1012   0.8367   1.0000
   2.500   0.6515   0.02775   0.01745  -0.1026   0.8186   1.0000
   2.750   0.6906   0.02789   0.01760  -0.1039   0.8017   1.0000
   3.000   0.7290   0.02795   0.01772  -0.1049   0.7855   1.0000
   3.250   0.7656   0.02800   0.01781  -0.1054   0.7696   1.0000
   3.500   0.7927   0.02844   0.01830  -0.1048   0.7518   1.0000
   3.750   0.8202   0.02888   0.01880  -0.1041   0.7345   1.0000
   4.000   0.8489   0.02925   0.01929  -0.1035   0.7181   1.0000
   4.250   0.8773   0.02963   0.01976  -0.1029   0.7021   1.0000
   4.500   0.9053   0.03005   0.02028  -0.1021   0.6863   1.0000
   4.750   0.9325   0.03052   0.02086  -0.1012   0.6707   1.0000
   5.000   0.9592   0.03103   0.02154  -0.1002   0.6552   1.0000
   5.250   0.9853   0.03160   0.02224  -0.0992   0.6399   1.0000
   5.500   1.0107   0.03224   0.02304  -0.0981   0.6246   1.0000
   5.750   1.0359   0.03292   0.02389  -0.0969   0.6094   1.0000
   6.000   1.0559   0.03407   0.02529  -0.0956   0.5936   1.0000
   6.250   1.0759   0.03524   0.02669  -0.0942   0.5777   1.0000
   6.500   1.0958   0.03644   0.02814  -0.0928   0.5618   1.0000
   6.750   1.1170   0.03730   0.02925  -0.0910   0.5437   1.0000
   7.000   1.1565   0.03326   0.02518  -0.0858   0.4977   1.0000
   7.250   1.1775   0.03016   0.02188  -0.0801   0.4411   1.0000
   7.500   1.1873   0.02885   0.02061  -0.0748   0.3873   1.0000
   7.750   1.1757   0.02884   0.02019  -0.0673   0.2736   1.0000
   8.000   1.1634   0.03217   0.02223  -0.0619   0.1689   1.0000
   8.250   1.1662   0.03518   0.02474  -0.0586   0.1366   1.0000
   8.500   1.1771   0.03766   0.02712  -0.0562   0.1176   1.0000
   8.750   1.1987   0.04023   0.02969  -0.0546   0.1059   1.0000
   9.000   1.2335   0.04341   0.03275  -0.0545   0.0973   1.0000
   9.250   1.2582   0.04633   0.03595  -0.0536   0.0904   1.0000
   9.500   1.2891   0.05070   0.04031  -0.0538   0.0859   1.0000
   9.750   1.3045   0.05444   0.04459  -0.0521   0.0851   1.0000
  10.000   1.3155   0.05855   0.04920  -0.0502   0.0849   1.0000
  10.250   1.3206   0.06287   0.05398  -0.0481   0.0849   1.0000
  10.500   1.3196   0.06719   0.05878  -0.0459   0.0849   1.0000
  10.750   1.3132   0.07149   0.06348  -0.0436   0.0850   1.0000
  11.000   1.3030   0.07589   0.06821  -0.0415   0.0852   1.0000
  11.250   1.2892   0.08024   0.07283  -0.0394   0.0855   1.0000
  11.500   1.2756   0.08485   0.07762  -0.0378   0.0859   1.0000
  11.750   1.0747   0.08693   0.08061  -0.0323   0.0917   1.0000
  12.000   1.0295   0.09578   0.08963  -0.0366   0.0937   1.0000
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