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

HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il)
Reynolds number: 100,000
Max Cl/Cd: 59.51 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq258-il-100000.txt
Download as CSV file: xf-hq258-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4060   0.09139   0.08665  -0.0291   1.0000   0.0702
  -7.750  -0.4101   0.08863   0.08399  -0.0306   1.0000   0.0727
  -7.500  -0.4209   0.08620   0.08168  -0.0328   1.0000   0.0741
  -7.250  -0.4291   0.08279   0.07833  -0.0403   1.0000   0.0750
  -7.000  -0.4369   0.08001   0.07545  -0.0458   1.0000   0.0755
  -6.750  -0.4242   0.07514   0.07082  -0.0377   1.0000   0.0781
  -6.500  -0.4212   0.07238   0.06810  -0.0365   1.0000   0.0809
  -6.250  -0.4215   0.06912   0.06485  -0.0379   1.0000   0.0847
  -6.000  -0.4262   0.06650   0.06178  -0.0457   1.0000   0.0891
  -5.750  -0.4201   0.06184   0.05745  -0.0413   1.0000   0.0914
  -5.500  -0.4143   0.05916   0.05480  -0.0397   1.0000   0.0958
  -5.250  -0.4067   0.05532   0.05071  -0.0425   1.0000   0.1044
  -5.000  -0.3954   0.05271   0.04782  -0.0439   1.0000   0.1167
  -4.750  -0.3852   0.04969   0.04478  -0.0433   1.0000   0.1306
  -4.500  -0.3754   0.04680   0.04201  -0.0417   1.0000   0.1462
  -4.250  -0.3622   0.04423   0.03934  -0.0417   1.0000   0.1721
  -4.000  -0.3482   0.04171   0.03685  -0.0406   1.0000   0.1921
  -3.500  -0.2604   0.03025   0.02320  -0.0457   1.0000   0.0724
  -3.250  -0.2301   0.02699   0.01929  -0.0453   1.0000   0.0588
  -3.000  -0.2019   0.02448   0.01633  -0.0449   1.0000   0.0547
  -2.750  -0.1743   0.02250   0.01391  -0.0443   1.0000   0.0534
  -2.500  -0.1489   0.02106   0.01220  -0.0435   1.0000   0.0554
  -2.250  -0.1247   0.02000   0.01098  -0.0428   1.0000   0.0642
  -2.000  -0.1003   0.01892   0.00984  -0.0420   1.0000   0.0755
  -1.750  -0.0758   0.01746   0.00867  -0.0417   1.0000   0.1128
  -1.500  -0.0556   0.01442   0.00834  -0.0398   1.0000   0.7194
  -1.250  -0.0438   0.01370   0.00807  -0.0351   1.0000   1.0000
  -1.000  -0.0192   0.01398   0.00797  -0.0355   1.0000   1.0000
  -0.750   0.0166   0.01441   0.00801  -0.0381   0.9957   1.0000
  -0.500   0.0597   0.01489   0.00819  -0.0419   0.9877   1.0000
  -0.250   0.1047   0.01543   0.00848  -0.0461   0.9801   1.0000
   0.000   0.1465   0.01584   0.00870  -0.0495   0.9711   1.0000
   0.250   0.1867   0.01623   0.00890  -0.0526   0.9618   1.0000
   0.500   0.2313   0.01664   0.00919  -0.0564   0.9537   1.0000
   0.750   0.2710   0.01694   0.00940  -0.0592   0.9437   1.0000
   1.000   0.3089   0.01721   0.00962  -0.0616   0.9333   1.0000
   1.250   0.3498   0.01745   0.00981  -0.0644   0.9232   1.0000
   1.500   0.3968   0.01749   0.00986  -0.0680   0.9118   1.0000
   1.750   0.4507   0.01730   0.00970  -0.0725   0.9011   1.0000
   2.000   0.4944   0.01716   0.00964  -0.0752   0.8891   1.0000
   2.250   0.5338   0.01702   0.00956  -0.0768   0.8768   1.0000
   2.500   0.5723   0.01680   0.00942  -0.0782   0.8641   1.0000
   2.750   0.6096   0.01650   0.00921  -0.0790   0.8507   1.0000
   3.000   0.6449   0.01615   0.00903  -0.0793   0.8363   1.0000
   3.250   0.6786   0.01578   0.00876  -0.0791   0.8213   1.0000
   3.500   0.7052   0.01557   0.00866  -0.0778   0.8015   1.0000
   3.750   0.7354   0.01517   0.00835  -0.0768   0.7821   1.0000
   4.000   0.7616   0.01489   0.00818  -0.0752   0.7577   1.0000
   4.250   0.7879   0.01462   0.00806  -0.0736   0.7301   1.0000
   4.500   0.8129   0.01443   0.00795  -0.0718   0.6964   1.0000
   4.750   0.8369   0.01436   0.00789  -0.0699   0.6533   1.0000
   5.000   0.8593   0.01444   0.00790  -0.0679   0.5970   1.0000
   5.250   0.8796   0.01483   0.00804  -0.0656   0.5261   1.0000
   5.500   0.8980   0.01561   0.00847  -0.0634   0.4501   1.0000
   5.750   0.9162   0.01662   0.00915  -0.0615   0.3847   1.0000
   6.000   0.9348   0.01769   0.01008  -0.0599   0.3305   1.0000
   6.250   0.9521   0.01870   0.01088  -0.0583   0.2765   1.0000
   6.500   0.9664   0.01955   0.01133  -0.0568   0.1916   1.0000
   6.750   0.9731   0.02252   0.01320  -0.0540   0.0607   1.0000
   7.000   0.9860   0.02483   0.01550  -0.0513   0.0432   1.0000
   7.250   1.0004   0.02695   0.01764  -0.0492   0.0374   1.0000
   7.500   1.0193   0.02948   0.02027  -0.0475   0.0348   1.0000
   7.750   1.0434   0.03214   0.02316  -0.0463   0.0330   1.0000
   8.000   1.0673   0.03515   0.02646  -0.0452   0.0318   1.0000
   8.250   1.0870   0.03788   0.02952  -0.0439   0.0297   1.0000
   8.500   1.1017   0.04070   0.03252  -0.0427   0.0269   1.0000
   8.750   1.1144   0.04469   0.03699  -0.0411   0.0263   1.0000
   9.000   1.1242   0.04813   0.04094  -0.0389   0.0267   1.0000
   9.250   1.1277   0.05220   0.04563  -0.0361   0.0279   1.0000
   9.500   1.1181   0.05715   0.05125  -0.0328   0.0293   1.0000
   9.750   1.1048   0.06185   0.05639  -0.0300   0.0304   1.0000
  10.000   1.0878   0.06584   0.06066  -0.0272   0.0312   1.0000
  10.250   1.0683   0.06992   0.06497  -0.0253   0.0317   1.0000
  10.500   1.0479   0.07437   0.06961  -0.0250   0.0321   1.0000
  10.750   1.0269   0.07942   0.07483  -0.0261   0.0324   1.0000
  11.000   1.0051   0.08543   0.08088  -0.0290   0.0326   1.0000
  11.250   0.9832   0.09258   0.08816  -0.0337   0.0329   1.0000
<< Back to HQ 2.5/8 AIRFOIL (hq258-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/8 AIRFOIL (hq258-il)