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HQ 0/10 AIRFOIL (hq010-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 0/10 AIRFOIL (hq010-il)
Reynolds number: 100,000
Max Cl/Cd: 39.73 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq010-il-100000.txt
Download as CSV file: xf-hq010-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/10 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6401   0.10723   0.10239   0.0075   1.0000   0.1576
 -10.000  -0.6718   0.10162   0.09693   0.0004   1.0000   0.1608
  -9.250  -0.8257   0.06229   0.05678  -0.0213   1.0000   0.0670
  -9.000  -0.8317   0.05797   0.05208  -0.0201   1.0000   0.0673
  -8.750  -0.8345   0.05300   0.04672  -0.0188   1.0000   0.0678
  -8.500  -0.8262   0.04814   0.04151  -0.0176   1.0000   0.0670
  -8.250  -0.8177   0.04353   0.03648  -0.0162   1.0000   0.0666
  -8.000  -0.8086   0.03959   0.03195  -0.0146   1.0000   0.0684
  -7.750  -0.7899   0.03659   0.02896  -0.0142   1.0000   0.0720
  -7.500  -0.7721   0.03346   0.02536  -0.0128   1.0000   0.0726
  -7.250  -0.7514   0.03086   0.02237  -0.0117   1.0000   0.0741
  -7.000  -0.7287   0.02862   0.01978  -0.0106   1.0000   0.0765
  -6.750  -0.7056   0.02714   0.01785  -0.0095   1.0000   0.0805
  -6.500  -0.6818   0.02490   0.01574  -0.0091   1.0000   0.0872
  -6.250  -0.6569   0.02338   0.01401  -0.0082   1.0000   0.0930
  -6.000  -0.6336   0.02155   0.01233  -0.0074   1.0000   0.1011
  -5.750  -0.6124   0.02009   0.01092  -0.0062   1.0000   0.1140
  -5.500  -0.5932   0.01875   0.00974  -0.0048   1.0000   0.1360
  -5.250  -0.5760   0.01742   0.00875  -0.0031   1.0000   0.1746
  -5.000  -0.5639   0.01572   0.00777  -0.0008   1.0000   0.2686
  -4.750  -0.5537   0.01430   0.00720   0.0019   1.0000   0.4004
  -4.500  -0.5407   0.01361   0.00708   0.0049   1.0000   0.5109
  -4.250  -0.5255   0.01347   0.00719   0.0079   1.0000   0.5919
  -4.000  -0.5100   0.01354   0.00729   0.0109   1.0000   0.6469
  -3.750  -0.4950   0.01368   0.00745   0.0141   1.0000   0.6871
  -3.500  -0.4805   0.01383   0.00760   0.0173   1.0000   0.7194
  -3.250  -0.4669   0.01400   0.00776   0.0207   1.0000   0.7482
  -3.000  -0.4548   0.01422   0.00798   0.0246   1.0000   0.7770
  -2.750  -0.4438   0.01443   0.00819   0.0287   1.0000   0.8049
  -2.500  -0.4320   0.01454   0.00824   0.0323   1.0000   0.8299
  -2.250  -0.4172   0.01460   0.00825   0.0351   1.0000   0.8498
  -2.000  -0.4021   0.01458   0.00816   0.0374   1.0000   0.8699
  -1.750  -0.3816   0.01464   0.00817   0.0389   1.0000   0.8879
  -1.500  -0.3562   0.01475   0.00820   0.0392   1.0000   0.9060
  -1.250  -0.3251   0.01489   0.00825   0.0382   1.0000   0.9244
  -1.000  -0.2740   0.01526   0.00852   0.0335   0.9971   0.9404
  -0.750  -0.2040   0.01579   0.00893   0.0253   0.9941   0.9545
  -0.500  -0.1299   0.01612   0.00919   0.0161   0.9917   0.9651
  -0.250  -0.0632   0.01624   0.00926   0.0078   0.9868   0.9733
   0.000   0.0001   0.01634   0.00935   0.0000   0.9814   0.9814
   0.250   0.0633   0.01624   0.00926  -0.0078   0.9733   0.9868
   0.500   0.1301   0.01611   0.00918  -0.0161   0.9649   0.9916
   0.750   0.2041   0.01578   0.00893  -0.0254   0.9545   0.9941
   1.000   0.2740   0.01526   0.00852  -0.0335   0.9405   0.9972
   1.250   0.3252   0.01489   0.00825  -0.0382   0.9244   1.0000
   1.500   0.3561   0.01475   0.00820  -0.0392   0.9061   1.0000
   1.750   0.3816   0.01464   0.00817  -0.0389   0.8880   1.0000
   2.000   0.4021   0.01458   0.00816  -0.0374   0.8700   1.0000
   2.250   0.4172   0.01460   0.00825  -0.0351   0.8499   1.0000
   2.500   0.4320   0.01454   0.00823  -0.0322   0.8299   1.0000
   2.750   0.4437   0.01443   0.00819  -0.0287   0.8049   1.0000
   3.000   0.4548   0.01422   0.00798  -0.0246   0.7770   1.0000
   3.250   0.4668   0.01400   0.00776  -0.0207   0.7482   1.0000
   3.500   0.4805   0.01383   0.00760  -0.0173   0.7195   1.0000
   3.750   0.4949   0.01368   0.00745  -0.0141   0.6870   1.0000
   4.000   0.5100   0.01354   0.00729  -0.0109   0.6468   1.0000
   4.250   0.5255   0.01347   0.00719  -0.0079   0.5921   1.0000
   4.500   0.5407   0.01361   0.00708  -0.0049   0.5110   1.0000
   4.750   0.5538   0.01430   0.00720  -0.0019   0.4005   1.0000
   5.000   0.5639   0.01571   0.00776   0.0008   0.2691   1.0000
   5.250   0.5760   0.01742   0.00875   0.0031   0.1745   1.0000
   5.500   0.5933   0.01875   0.00974   0.0048   0.1360   1.0000
   5.750   0.6124   0.02010   0.01092   0.0062   0.1139   1.0000
   6.000   0.6336   0.02155   0.01234   0.0074   0.1009   1.0000
   6.250   0.6569   0.02338   0.01401   0.0082   0.0930   1.0000
   6.500   0.6818   0.02489   0.01574   0.0091   0.0870   1.0000
   6.750   0.7056   0.02714   0.01785   0.0095   0.0805   1.0000
   7.000   0.7287   0.02862   0.01978   0.0106   0.0764   1.0000
   7.250   0.7513   0.03086   0.02238   0.0117   0.0740   1.0000
   7.500   0.7721   0.03347   0.02537   0.0128   0.0725   1.0000
   7.750   0.7899   0.03660   0.02897   0.0142   0.0720   1.0000
   8.000   0.8087   0.03952   0.03191   0.0147   0.0686   1.0000
   8.250   0.8177   0.04353   0.03648   0.0162   0.0666   1.0000
   8.500   0.8263   0.04813   0.04149   0.0176   0.0670   1.0000
   8.750   0.8347   0.05290   0.04662   0.0188   0.0679   1.0000
   9.000   0.8327   0.05835   0.05243   0.0199   0.0674   1.0000
   9.250   0.8242   0.06200   0.05653   0.0215   0.0669   1.0000
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