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

HQ 2.5/9 B AIRFOIL (hq259b-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 B AIRFOIL (hq259b-il)
Reynolds number: 100,000
Max Cl/Cd: 58.41 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259b-il-100000.txt
Download as CSV file: xf-hq259b-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 B 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4225   0.09200   0.08709  -0.0335   1.0000   0.1006
  -8.250  -0.4439   0.08966   0.08491  -0.0370   1.0000   0.1022
  -8.000  -0.4686   0.08617   0.08153  -0.0441   1.0000   0.1029
  -7.750  -0.4381   0.08262   0.07800  -0.0341   1.0000   0.1071
  -7.500  -0.4393   0.07999   0.07543  -0.0334   1.0000   0.1109
  -7.250  -0.4510   0.07654   0.07208  -0.0366   1.0000   0.1145
  -7.000  -0.4849   0.07359   0.06893  -0.0447   1.0000   0.1175
  -6.750  -0.4719   0.06928   0.06488  -0.0394   1.0000   0.1204
  -6.500  -0.4684   0.06729   0.06294  -0.0363   1.0000   0.1256
  -6.250  -0.4837   0.06357   0.05890  -0.0418   1.0000   0.1332
  -6.000  -0.4737   0.06125   0.05685  -0.0362   1.0000   0.1388
  -5.750  -0.4733   0.05776   0.05321  -0.0381   1.0000   0.1498
  -5.500  -0.4675   0.05498   0.05028  -0.0384   1.0000   0.1639
  -5.250  -0.4587   0.05228   0.04750  -0.0382   1.0000   0.1785
  -5.000  -0.4483   0.04961   0.04482  -0.0374   1.0000   0.1938
  -4.750  -0.4372   0.04715   0.04240  -0.0359   1.0000   0.2107
  -4.250  -0.3479   0.03271   0.02523  -0.0458   1.0000   0.0900
  -4.000  -0.3228   0.03028   0.02252  -0.0456   1.0000   0.0866
  -3.750  -0.2959   0.02756   0.01946  -0.0456   1.0000   0.0822
  -3.500  -0.2682   0.02544   0.01683  -0.0453   1.0000   0.0791
  -3.250  -0.2425   0.02395   0.01503  -0.0448   1.0000   0.0792
  -3.000  -0.2185   0.02294   0.01387  -0.0443   1.0000   0.0835
  -2.750  -0.1939   0.02197   0.01269  -0.0436   1.0000   0.0865
  -2.500  -0.1696   0.02106   0.01165  -0.0428   1.0000   0.0892
  -2.250  -0.1459   0.02001   0.01073  -0.0424   1.0000   0.0955
  -2.000  -0.1225   0.01933   0.01012  -0.0419   1.0000   0.1074
  -1.750  -0.0982   0.01866   0.00955  -0.0415   1.0000   0.1248
  -1.500  -0.0717   0.01601   0.00928  -0.0416   1.0000   0.6239
  -1.250  -0.0667   0.01598   0.00970  -0.0354   1.0000   0.7910
  -1.000  -0.0613   0.01588   0.00978  -0.0299   1.0000   0.8797
  -0.750  -0.0100   0.01590   0.00970  -0.0349   0.9900   1.0000
  -0.500   0.0437   0.01644   0.00994  -0.0409   0.9794   1.0000
  -0.250   0.0978   0.01697   0.01023  -0.0467   0.9698   1.0000
   0.000   0.1372   0.01723   0.01032  -0.0497   0.9599   1.0000
   0.250   0.1795   0.01757   0.01052  -0.0530   0.9512   1.0000
   0.500   0.2241   0.01789   0.01072  -0.0567   0.9429   1.0000
   0.750   0.2603   0.01814   0.01089  -0.0588   0.9329   1.0000
   1.000   0.3054   0.01838   0.01107  -0.0623   0.9246   1.0000
   1.250   0.3454   0.01853   0.01119  -0.0647   0.9144   1.0000
   1.500   0.3827   0.01866   0.01132  -0.0666   0.9035   1.0000
   1.750   0.4271   0.01867   0.01134  -0.0695   0.8936   1.0000
   2.000   0.4759   0.01851   0.01123  -0.0731   0.8840   1.0000
   2.250   0.5159   0.01839   0.01117  -0.0749   0.8717   1.0000
   2.500   0.5571   0.01813   0.01099  -0.0767   0.8593   1.0000
   2.750   0.5980   0.01776   0.01071  -0.0781   0.8465   1.0000
   3.000   0.6369   0.01732   0.01038  -0.0789   0.8322   1.0000
   3.250   0.6732   0.01682   0.00998  -0.0791   0.8159   1.0000
   3.500   0.7085   0.01625   0.00950  -0.0788   0.7977   1.0000
   3.750   0.7390   0.01575   0.00910  -0.0777   0.7745   1.0000
   4.000   0.7681   0.01521   0.00860  -0.0762   0.7454   1.0000
   4.250   0.7952   0.01472   0.00809  -0.0743   0.7078   1.0000
   4.500   0.8185   0.01450   0.00785  -0.0722   0.6573   1.0000
   4.750   0.8419   0.01449   0.00767  -0.0704   0.5993   1.0000
   5.000   0.8650   0.01481   0.00768  -0.0687   0.5439   1.0000
   5.250   0.8872   0.01543   0.00802  -0.0673   0.4971   1.0000
   5.500   0.9090   0.01615   0.00856  -0.0661   0.4565   1.0000
   5.750   0.9299   0.01688   0.00913  -0.0647   0.4169   1.0000
   6.000   0.9497   0.01760   0.00975  -0.0633   0.3745   1.0000
   6.250   0.9667   0.01836   0.01037  -0.0615   0.3204   1.0000
   6.500   0.9804   0.01939   0.01105  -0.0593   0.2464   1.0000
   6.750   0.9934   0.02089   0.01204  -0.0572   0.1824   1.0000
   7.000   1.0086   0.02250   0.01334  -0.0555   0.1444   1.0000
   7.250   1.0253   0.02422   0.01488  -0.0538   0.1190   1.0000
   7.500   1.0440   0.02627   0.01669  -0.0526   0.1018   1.0000
   7.750   1.0657   0.02810   0.01846  -0.0518   0.0889   1.0000
   8.000   1.0906   0.03003   0.02062  -0.0510   0.0811   1.0000
   8.250   1.1172   0.03272   0.02321  -0.0510   0.0748   1.0000
   8.500   1.1386   0.03448   0.02535  -0.0498   0.0696   1.0000
   8.750   1.1614   0.03690   0.02797  -0.0491   0.0661   1.0000
   9.000   1.1840   0.04088   0.03203  -0.0489   0.0631   1.0000
   9.250   1.1956   0.04324   0.03501  -0.0465   0.0610   1.0000
   9.500   1.2053   0.04629   0.03860  -0.0443   0.0591   1.0000
   9.750   1.2115   0.05011   0.04293  -0.0420   0.0586   1.0000
  10.000   1.2122   0.05427   0.04757  -0.0394   0.0585   1.0000
  10.250   1.2072   0.05851   0.05225  -0.0367   0.0586   1.0000
  10.500   1.1969   0.06271   0.05684  -0.0339   0.0586   1.0000
  10.750   1.1808   0.06673   0.06116  -0.0309   0.0588   1.0000
  11.000   1.1601   0.07069   0.06536  -0.0282   0.0592   1.0000
  11.250   1.1378   0.07511   0.06998  -0.0267   0.0598   1.0000
  11.500   1.1146   0.08010   0.07515  -0.0265   0.0604   1.0000
  11.750   1.0921   0.08582   0.08101  -0.0277   0.0611   1.0000
  12.000   1.0718   0.09219   0.08749  -0.0299   0.0618   1.0000
  12.250   1.0574   0.09918   0.09453  -0.0324   0.0626   1.0000
  12.500   0.8397   0.12135   0.11722  -0.0458   0.0747   1.0000
<< Back to HQ 2.5/9 B AIRFOIL (hq259b-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/9 B AIRFOIL (hq259b-il)