Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il)
Reynolds number: 50,000
Max Cl/Cd: 38.56 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq2510-il-50000-n5.txt
Download as CSV file: xf-hq2510-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/10 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4075   0.11308   0.10601  -0.0308   1.0000   0.0638
  -9.750  -0.4093   0.10810   0.10108  -0.0339   1.0000   0.0522
  -9.500  -0.4042   0.10424   0.09726  -0.0341   1.0000   0.0499
  -9.250  -0.4030   0.10035   0.09344  -0.0353   1.0000   0.0481
  -9.000  -0.4040   0.09630   0.08949  -0.0369   1.0000   0.0464
  -8.750  -0.4074   0.09207   0.08535  -0.0388   1.0000   0.0449
  -8.250  -0.4335   0.08123   0.07474  -0.0476   1.0000   0.0408
  -8.000  -0.4381   0.07770   0.07130  -0.0480   1.0000   0.0404
  -7.750  -0.4440   0.07414   0.06779  -0.0485   1.0000   0.0401
  -7.500  -0.4507   0.07057   0.06425  -0.0488   1.0000   0.0398
  -7.250  -0.4571   0.06707   0.06074  -0.0488   1.0000   0.0394
  -7.000  -0.4629   0.06348   0.05709  -0.0485   1.0000   0.0392
  -6.750  -0.4667   0.05992   0.05341  -0.0480   1.0000   0.0390
  -6.500  -0.4678   0.05629   0.04960  -0.0475   1.0000   0.0388
  -6.250  -0.4654   0.05261   0.04566  -0.0470   1.0000   0.0387
  -6.000  -0.4592   0.04898   0.04171  -0.0465   1.0000   0.0388
  -5.750  -0.4494   0.04536   0.03768  -0.0461   1.0000   0.0391
  -5.500  -0.4358   0.04193   0.03371  -0.0456   1.0000   0.0400
  -5.250  -0.4208   0.03894   0.03034  -0.0452   1.0000   0.0427
  -5.000  -0.4043   0.03703   0.02825  -0.0445   1.0000   0.0461
  -4.750  -0.3843   0.03453   0.02522  -0.0439   1.0000   0.0486
  -4.500  -0.3622   0.03209   0.02225  -0.0431   1.0000   0.0511
  -4.250  -0.3398   0.03006   0.01980  -0.0421   1.0000   0.0551
  -4.000  -0.3192   0.02882   0.01843  -0.0414   1.0000   0.0638
  -3.750  -0.2892   0.02718   0.01660  -0.0417   0.9966   0.0726
  -3.500  -0.2567   0.02601   0.01528  -0.0428   0.9914   0.0900
  -3.250  -0.2238   0.02493   0.01405  -0.0439   0.9862   0.1098
  -3.000  -0.1914   0.02389   0.01298  -0.0453   0.9806   0.1318
  -2.750  -0.1573   0.02280   0.01200  -0.0473   0.9750   0.1786
  -2.500  -0.1288   0.02081   0.01169  -0.0486   0.9705   0.4707
  -2.250  -0.1079   0.02074   0.01205  -0.0460   0.9628   0.6609
  -2.000  -0.0851   0.02086   0.01217  -0.0439   0.9556   0.7448
  -1.750  -0.0692   0.02082   0.01224  -0.0398   0.9477   0.8309
  -1.500  -0.0234   0.02068   0.01206  -0.0409   0.9445   0.9401
  -1.250   0.0308   0.02065   0.01169  -0.0469   0.9383   1.0000
  -1.000   0.0669   0.02077   0.01150  -0.0498   0.9312   1.0000
  -0.750   0.0950   0.02088   0.01134  -0.0511   0.9215   1.0000
  -0.500   0.1285   0.02106   0.01128  -0.0531   0.9134   1.0000
  -0.250   0.1642   0.02126   0.01127  -0.0554   0.9056   1.0000
   0.000   0.1950   0.02148   0.01131  -0.0567   0.8965   1.0000
   0.250   0.2356   0.02166   0.01131  -0.0596   0.8900   1.0000
   0.500   0.2640   0.02190   0.01143  -0.0603   0.8797   1.0000
   0.750   0.2965   0.02213   0.01155  -0.0616   0.8706   1.0000
   1.000   0.3345   0.02229   0.01163  -0.0637   0.8631   1.0000
   1.250   0.3626   0.02253   0.01181  -0.0641   0.8523   1.0000
   1.500   0.3932   0.02274   0.01198  -0.0649   0.8424   1.0000
   1.750   0.4314   0.02281   0.01205  -0.0667   0.8346   1.0000
   2.000   0.4594   0.02300   0.01223  -0.0668   0.8225   1.0000
   2.250   0.4900   0.02306   0.01230  -0.0670   0.8096   1.0000
   2.500   0.5214   0.02302   0.01230  -0.0672   0.7960   1.0000
   2.750   0.5525   0.02296   0.01231  -0.0672   0.7821   1.0000
   3.000   0.5827   0.02292   0.01232  -0.0671   0.7682   1.0000
   3.250   0.6122   0.02289   0.01236  -0.0669   0.7541   1.0000
   3.500   0.6416   0.02283   0.01239  -0.0666   0.7395   1.0000
   3.750   0.6713   0.02272   0.01241  -0.0661   0.7240   1.0000
   4.000   0.6954   0.02277   0.01258  -0.0650   0.7047   1.0000
   4.250   0.7236   0.02264   0.01256  -0.0642   0.6857   1.0000
   4.500   0.7496   0.02258   0.01260  -0.0631   0.6640   1.0000
   4.750   0.7774   0.02243   0.01260  -0.0621   0.6411   1.0000
   5.000   0.8020   0.02243   0.01271  -0.0608   0.6139   1.0000
   5.250   0.8255   0.02250   0.01287  -0.0594   0.5828   1.0000
   5.500   0.8494   0.02260   0.01300  -0.0580   0.5480   1.0000
   5.750   0.8724   0.02281   0.01319  -0.0565   0.5098   1.0000
   6.000   0.8943   0.02319   0.01356  -0.0549   0.4698   1.0000
   6.250   0.9148   0.02377   0.01403  -0.0533   0.4305   1.0000
   6.500   0.9340   0.02453   0.01468  -0.0518   0.3936   1.0000
   6.750   0.9524   0.02541   0.01549  -0.0503   0.3595   1.0000
   7.000   0.9702   0.02637   0.01640  -0.0488   0.3284   1.0000
   7.250   0.9879   0.02740   0.01741  -0.0474   0.3006   1.0000
   7.500   1.0054   0.02849   0.01850  -0.0461   0.2751   1.0000
   7.750   1.0233   0.02964   0.01974  -0.0448   0.2525   1.0000
   8.000   1.0415   0.03081   0.02104  -0.0436   0.2308   1.0000
   8.250   1.0593   0.03204   0.02232  -0.0424   0.2118   1.0000
   8.500   1.0759   0.03330   0.02378  -0.0410   0.1924   1.0000
   8.750   1.0908   0.03463   0.02523  -0.0395   0.1743   1.0000
   9.000   1.1009   0.03601   0.02673  -0.0375   0.1547   1.0000
   9.250   1.0995   0.03730   0.02804  -0.0347   0.1320   1.0000
   9.500   1.0939   0.03884   0.02952  -0.0320   0.1053   1.0000
   9.750   1.0935   0.04091   0.03150  -0.0302   0.0745   1.0000
  10.000   1.0922   0.04357   0.03404  -0.0285   0.0556   1.0000
  10.250   1.0901   0.04664   0.03714  -0.0268   0.0447   1.0000
  10.500   1.0898   0.04966   0.04030  -0.0253   0.0396   1.0000
  10.750   1.0900   0.05268   0.04350  -0.0241   0.0360   1.0000
  11.000   1.0871   0.05596   0.04687  -0.0234   0.0335   1.0000
  11.250   1.0859   0.05938   0.05056  -0.0228   0.0316   1.0000
  11.500   1.0864   0.06287   0.05438  -0.0223   0.0297   1.0000
  11.750   1.0834   0.06672   0.05851  -0.0223   0.0278   1.0000
  12.000   1.0792   0.07083   0.06288  -0.0227   0.0269   1.0000
  12.250   1.0721   0.07538   0.06765  -0.0238   0.0259   1.0000
  12.500   1.0636   0.08033   0.07280  -0.0254   0.0251   1.0000
  12.750   1.0539   0.08572   0.07838  -0.0274   0.0246   1.0000
  13.000   1.0431   0.09166   0.08454  -0.0300   0.0243   1.0000
  13.250   1.0306   0.09841   0.09155  -0.0333   0.0245   1.0000
  13.500   1.0154   0.10620   0.09958  -0.0375   0.0247   1.0000
  13.750   0.9947   0.11624   0.10990  -0.0433   0.0257   1.0000
  14.000   0.9697   0.12853   0.12238  -0.0504   0.0276   1.0000
<< Back to HQ 2.5/10 AIRFOIL (hq2510-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/10 AIRFOIL (hq2510-il)