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HQ 2.0/8 AIRFOIL (hq208-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.0/8 AIRFOIL (hq208-il)
Reynolds number: 100,000
Max Cl/Cd: 55.88 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq208-il-100000.txt
Download as CSV file: xf-hq208-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4588   0.09217   0.08743  -0.0253   1.0000   0.0804
  -8.000  -0.4726   0.08919   0.08458  -0.0304   1.0000   0.0816
  -7.750  -0.4815   0.08508   0.08048  -0.0385   1.0000   0.0820
  -7.500  -0.4582   0.08187   0.07734  -0.0275   1.0000   0.0890
  -7.250  -0.4616   0.07792   0.07346  -0.0322   1.0000   0.0928
  -7.000  -0.4673   0.07351   0.06899  -0.0397   1.0000   0.0948
  -6.750  -0.4594   0.07001   0.06562  -0.0363   1.0000   0.1007
  -6.500  -0.4102   0.05592   0.05193  -0.0338   1.0000   0.1155
  -6.250  -0.4252   0.05209   0.04796  -0.0384   1.0000   0.1220
  -5.750  -0.4442   0.05621   0.05171  -0.0395   1.0000   0.1292
  -5.500  -0.4387   0.05286   0.04826  -0.0397   1.0000   0.1405
  -5.250  -0.4311   0.04993   0.04519  -0.0393   1.0000   0.1537
  -5.000  -0.4220   0.04722   0.04244  -0.0382   1.0000   0.1688
  -4.750  -0.4124   0.04469   0.03977  -0.0378   1.0000   0.1929
  -4.250  -0.3413   0.03116   0.02405  -0.0407   1.0000   0.0658
  -4.000  -0.3149   0.02853   0.02056  -0.0392   1.0000   0.0554
  -3.750  -0.2920   0.02539   0.01714  -0.0386   1.0000   0.0537
  -3.500  -0.2678   0.02331   0.01469  -0.0376   1.0000   0.0537
  -3.250  -0.2440   0.02150   0.01263  -0.0369   1.0000   0.0586
  -3.000  -0.2200   0.02012   0.01114  -0.0360   1.0000   0.0634
  -2.750  -0.1954   0.01894   0.00975  -0.0349   1.0000   0.0689
  -2.500  -0.1723   0.01738   0.00841  -0.0341   1.0000   0.0866
  -2.250  -0.1490   0.01599   0.00728  -0.0335   1.0000   0.1411
  -2.000  -0.1396   0.01321   0.00725  -0.0292   1.0000   0.7375
  -1.750  -0.1314   0.01307   0.00725  -0.0241   1.0000   0.8356
  -1.500  -0.0986   0.01270   0.00690  -0.0243   1.0000   1.0000
  -1.250  -0.0744   0.01285   0.00675  -0.0249   1.0000   1.0000
  -1.000  -0.0506   0.01306   0.00669  -0.0253   1.0000   1.0000
  -0.750  -0.0275   0.01331   0.00668  -0.0255   1.0000   1.0000
  -0.500  -0.0030   0.01361   0.00679  -0.0260   0.9993   1.0000
  -0.250   0.0416   0.01411   0.00708  -0.0302   0.9911   1.0000
   0.000   0.0869   0.01462   0.00742  -0.0345   0.9829   1.0000
   0.250   0.1285   0.01502   0.00767  -0.0379   0.9734   1.0000
   0.500   0.1717   0.01543   0.00799  -0.0416   0.9643   1.0000
   0.750   0.2171   0.01581   0.00830  -0.0456   0.9556   1.0000
   1.000   0.2577   0.01607   0.00853  -0.0486   0.9443   1.0000
   1.250   0.3054   0.01620   0.00864  -0.0525   0.9312   1.0000
   1.500   0.3597   0.01608   0.00855  -0.0572   0.9169   1.0000
   1.750   0.4086   0.01595   0.00848  -0.0609   0.9041   1.0000
   2.000   0.4508   0.01589   0.00850  -0.0633   0.8923   1.0000
   2.250   0.4907   0.01579   0.00848  -0.0651   0.8801   1.0000
   2.500   0.5283   0.01565   0.00843  -0.0663   0.8671   1.0000
   2.750   0.5631   0.01549   0.00842  -0.0668   0.8530   1.0000
   3.000   0.5956   0.01529   0.00834  -0.0667   0.8374   1.0000
   3.250   0.6268   0.01505   0.00820  -0.0661   0.8209   1.0000
   3.500   0.6534   0.01487   0.00814  -0.0648   0.8005   1.0000
   3.750   0.6814   0.01457   0.00793  -0.0634   0.7792   1.0000
   4.000   0.7067   0.01432   0.00783  -0.0615   0.7524   1.0000
   4.250   0.7303   0.01409   0.00768  -0.0593   0.7173   1.0000
   4.500   0.7536   0.01384   0.00739  -0.0568   0.6700   1.0000
   4.750   0.7745   0.01386   0.00726  -0.0543   0.6032   1.0000
   5.000   0.7939   0.01424   0.00733  -0.0518   0.5207   1.0000
   5.250   0.8124   0.01499   0.00767  -0.0497   0.4371   1.0000
   5.500   0.8310   0.01589   0.00832  -0.0481   0.3672   1.0000
   5.750   0.8484   0.01696   0.00906  -0.0464   0.2938   1.0000
   6.000   0.8617   0.01864   0.01013  -0.0443   0.1913   1.0000
   6.250   0.8774   0.02050   0.01151  -0.0424   0.1338   1.0000
   6.500   0.8965   0.02225   0.01308  -0.0409   0.1128   1.0000
   6.750   0.9176   0.02379   0.01461  -0.0398   0.0955   1.0000
   7.000   0.9398   0.02572   0.01654  -0.0387   0.0815   1.0000
   7.250   0.9580   0.02770   0.01853  -0.0377   0.0615   1.0000
   7.500   0.9787   0.03042   0.02149  -0.0363   0.0475   1.0000
   7.750   0.9989   0.03394   0.02505  -0.0356   0.0398   1.0000
   8.000   1.0191   0.03692   0.02860  -0.0338   0.0378   1.0000
   8.250   1.0349   0.04063   0.03289  -0.0319   0.0367   1.0000
   8.500   1.0459   0.04471   0.03753  -0.0299   0.0366   1.0000
   8.750   1.0519   0.04915   0.04249  -0.0276   0.0369   1.0000
   9.000   1.0532   0.05375   0.04756  -0.0255   0.0376   1.0000
   9.250   1.0506   0.05850   0.05267  -0.0235   0.0383   1.0000
   9.500   1.0452   0.06350   0.05793  -0.0219   0.0391   1.0000
   9.750   1.0481   0.06770   0.06238  -0.0204   0.0412   1.0000
  10.000   1.0078   0.07173   0.06694  -0.0173   0.0432   1.0000
  10.250   0.9781   0.07691   0.07233  -0.0175   0.0443   1.0000
  10.500   0.9520   0.08292   0.07849  -0.0202   0.0451   1.0000
  10.750   0.9252   0.09072   0.08640  -0.0258   0.0458   1.0000
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