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NLR-7223-62 AIRFOIL (nl722362-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NLR-7223-62 AIRFOIL (nl722362-il)
Reynolds number: 100,000
Max Cl/Cd: 42.22 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nl722362-il-100000.txt
Download as CSV file: xf-nl722362-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-7223-62 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4888   0.10225   0.09719  -0.0249   1.0000   0.0886
  -9.250  -0.5072   0.09897   0.09402  -0.0308   1.0000   0.0903
  -9.000  -0.5248   0.09539   0.09053  -0.0356   1.0000   0.0907
  -8.750  -0.4972   0.09070   0.08583  -0.0293   1.0000   0.0946
  -8.500  -0.4944   0.08764   0.08282  -0.0291   1.0000   0.0977
  -8.250  -0.5038   0.08419   0.07944  -0.0307   1.0000   0.1003
  -8.000  -0.5228   0.08132   0.07664  -0.0307   1.0000   0.1022
  -7.750  -0.5497   0.07898   0.07430  -0.0303   1.0000   0.1043
  -7.500  -0.5830   0.07820   0.07336  -0.0275   1.0000   0.1054
  -7.250  -0.5819   0.07329   0.06856  -0.0255   1.0000   0.1074
  -7.000  -0.5704   0.07032   0.06570  -0.0226   1.0000   0.1113
  -6.750  -0.5769   0.06798   0.06332  -0.0199   1.0000   0.1156
  -6.500  -0.6055   0.06710   0.06196  -0.0169   1.0000   0.1214
  -6.250  -0.5890   0.06225   0.05743  -0.0151   1.0000   0.1254
  -6.000  -0.6027   0.06162   0.05630  -0.0120   1.0000   0.1365
  -5.750  -0.5884   0.05689   0.05189  -0.0103   1.0000   0.1406
  -5.500  -0.5890   0.05455   0.04933  -0.0078   1.0000   0.1534
  -3.750  -0.4883   0.03304   0.02486   0.0086   1.0000   0.0921
  -3.500  -0.4629   0.02981   0.02126   0.0105   1.0000   0.0744
  -3.250  -0.4392   0.02710   0.01811   0.0123   1.0000   0.0674
  -3.000  -0.4147   0.02631   0.01674   0.0145   1.0000   0.0633
  -2.750  -0.3901   0.02440   0.01465   0.0155   1.0000   0.0625
  -2.500  -0.3656   0.02261   0.01277   0.0165   1.0000   0.0653
  -2.250  -0.3419   0.02146   0.01158   0.0175   1.0000   0.0679
  -2.000  -0.3184   0.02042   0.01053   0.0188   1.0000   0.0707
  -1.750  -0.1195   0.01555   0.00913  -0.0119   1.0000   1.0000
  -1.500  -0.1062   0.01553   0.00886  -0.0094   1.0000   1.0000
  -1.250  -0.0922   0.01555   0.00868  -0.0072   1.0000   1.0000
  -1.000  -0.0774   0.01559   0.00853  -0.0050   1.0000   1.0000
  -0.750  -0.0620   0.01565   0.00844  -0.0029   1.0000   1.0000
  -0.500  -0.0463   0.01574   0.00840  -0.0010   1.0000   1.0000
  -0.250  -0.0302   0.01585   0.00839   0.0009   1.0000   1.0000
   0.000  -0.0138   0.01598   0.00843   0.0027   1.0000   1.0000
   0.250   0.0026   0.01614   0.00849   0.0045   1.0000   1.0000
   0.500   0.0192   0.01632   0.00861   0.0062   1.0000   1.0000
   0.750   0.0358   0.01653   0.00876   0.0078   1.0000   1.0000
   1.000   0.0524   0.01677   0.00896   0.0094   1.0000   1.0000
   1.250   0.1042   0.01726   0.00944   0.0037   0.9902   1.0000
   1.500   0.1835   0.01775   0.00995  -0.0070   0.9721   1.0000
   1.750   0.2458   0.01787   0.01014  -0.0140   0.9554   1.0000
   2.000   0.3111   0.01776   0.01013  -0.0212   0.9376   1.0000
   2.250   0.3843   0.01734   0.00988  -0.0294   0.9191   1.0000
   2.500   0.4735   0.01630   0.00907  -0.0400   0.8984   1.0000
   2.750   0.5218   0.01530   0.00827  -0.0419   0.8665   1.0000
   3.000   0.5474   0.01442   0.00749  -0.0392   0.8164   1.0000
   3.250   0.5763   0.01365   0.00562  -0.0351   0.5429   1.0000
   3.500   0.5876   0.01507   0.00601  -0.0315   0.4029   1.0000
   3.750   0.6067   0.01597   0.00649  -0.0298   0.3554   1.0000
   4.000   0.6281   0.01671   0.00701  -0.0286   0.3291   1.0000
   4.250   0.6506   0.01738   0.00755  -0.0276   0.3103   1.0000
   4.500   0.6735   0.01804   0.00815  -0.0266   0.2950   1.0000
   4.750   0.6967   0.01870   0.00878  -0.0258   0.2821   1.0000
   5.000   0.7201   0.01941   0.00944  -0.0250   0.2704   1.0000
   5.250   0.7436   0.02012   0.01011  -0.0242   0.2593   1.0000
   5.500   0.7662   0.02078   0.01091  -0.0232   0.2480   1.0000
   5.750   0.7891   0.02157   0.01176  -0.0223   0.2373   1.0000
   6.000   0.8117   0.02240   0.01257  -0.0215   0.2259   1.0000
   6.250   0.8334   0.02319   0.01343  -0.0204   0.2138   1.0000
   6.500   0.8537   0.02397   0.01437  -0.0190   0.2007   1.0000
   6.750   0.8724   0.02464   0.01510  -0.0175   0.1855   1.0000
   7.000   0.8898   0.02516   0.01561  -0.0159   0.1690   1.0000
   7.250   0.9066   0.02579   0.01624  -0.0141   0.1522   1.0000
   7.500   0.9201   0.02628   0.01698  -0.0116   0.1336   1.0000
   7.750   0.9331   0.02669   0.01744  -0.0090   0.1142   1.0000
   8.000   0.9461   0.02732   0.01810  -0.0064   0.0969   1.0000
   8.250   0.9614   0.02846   0.01934  -0.0041   0.0844   1.0000
   8.500   0.9770   0.02981   0.02068  -0.0021   0.0757   1.0000
   8.750   0.9931   0.03145   0.02250  -0.0002   0.0693   1.0000
   9.000   1.0098   0.03397   0.02513   0.0014   0.0656   1.0000
   9.250   1.0210   0.03636   0.02800   0.0041   0.0625   1.0000
   9.500   1.0319   0.03831   0.03018   0.0065   0.0592   1.0000
   9.750   1.0442   0.04068   0.03260   0.0082   0.0567   1.0000
  10.000   1.0470   0.04424   0.03652   0.0112   0.0558   1.0000
  10.250   1.0438   0.04746   0.04019   0.0149   0.0555   1.0000
  10.500   1.0358   0.05087   0.04400   0.0188   0.0554   1.0000
  10.750   1.0226   0.05439   0.04789   0.0229   0.0554   1.0000
  11.000   1.0065   0.05792   0.05168   0.0268   0.0555   1.0000
  11.250   0.9865   0.06123   0.05518   0.0309   0.0557   1.0000
  11.500   0.9656   0.06495   0.05907   0.0338   0.0559   1.0000
  11.750   0.9422   0.06893   0.06325   0.0355   0.0562   1.0000
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