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

HQ 3.0/12 AIRFOIL (hq3012-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 3.0/12 AIRFOIL (hq3012-il)
Reynolds number: 50,000
Max Cl/Cd: 36.8 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq3012-il-50000-n5.txt
Download as CSV file: xf-hq3012-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.0/12 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  -0.3644   0.10422   0.09714  -0.0452   1.0000   0.0530
  -9.750  -0.3686   0.10014   0.09316  -0.0465   1.0000   0.0519
  -9.500  -0.3750   0.09592   0.08905  -0.0478   1.0000   0.0509
  -9.000  -0.4170   0.08319   0.07657  -0.0555   1.0000   0.0471
  -8.750  -0.4313   0.07990   0.07340  -0.0553   1.0000   0.0468
  -8.500  -0.4505   0.07724   0.07085  -0.0540   1.0000   0.0466
  -8.250  -0.4689   0.07451   0.06819  -0.0528   1.0000   0.0464
  -8.000  -0.4860   0.07170   0.06540  -0.0514   1.0000   0.0461
  -7.750  -0.5022   0.06869   0.06236  -0.0500   1.0000   0.0459
  -7.500  -0.5151   0.06551   0.05910  -0.0489   1.0000   0.0458
  -7.250  -0.5248   0.06205   0.05550  -0.0480   1.0000   0.0457
  -7.000  -0.5304   0.05833   0.05154  -0.0473   1.0000   0.0458
  -6.750  -0.5309   0.05463   0.04754  -0.0468   1.0000   0.0459
  -6.500  -0.5132   0.04963   0.04193  -0.0497   0.9950   0.0465
  -6.250  -0.4894   0.04563   0.03749  -0.0524   0.9891   0.0486
  -6.000  -0.4613   0.04330   0.03496  -0.0547   0.9838   0.0522
  -5.750  -0.4348   0.04018   0.03128  -0.0563   0.9779   0.0557
  -5.500  -0.4031   0.03710   0.02740  -0.0581   0.9730   0.0603
  -5.250  -0.3753   0.03534   0.02554  -0.0593   0.9677   0.0661
  -5.000  -0.3445   0.03344   0.02306  -0.0601   0.9623   0.0737
  -4.750  -0.3122   0.03194   0.02151  -0.0616   0.9579   0.0827
  -4.500  -0.2851   0.03073   0.02015  -0.0617   0.9520   0.0922
  -4.250  -0.2545   0.02968   0.01884  -0.0621   0.9469   0.1021
  -4.000  -0.2218   0.02876   0.01782  -0.0632   0.9425   0.1152
  -3.750  -0.1979   0.02792   0.01699  -0.0627   0.9357   0.1251
  -3.500  -0.1660   0.02716   0.01618  -0.0640   0.9306   0.1456
  -3.000  -0.1097   0.02512   0.01474  -0.0660   0.9188   0.2551
  -2.750  -0.0846   0.02408   0.01510  -0.0654   0.9143   0.5281
  -2.500  -0.0674   0.02433   0.01541  -0.0627   0.9062   0.6233
  -2.250  -0.0416   0.02463   0.01566  -0.0612   0.9006   0.6871
  -2.000  -0.0271   0.02490   0.01589  -0.0577   0.8926   0.7374
  -1.750  -0.0093   0.02505   0.01607  -0.0540   0.8866   0.7905
  -1.500   0.0027   0.02508   0.01609  -0.0498   0.8787   0.8327
  -1.250   0.0285   0.02501   0.01592  -0.0485   0.8728   0.8667
  -1.000   0.0576   0.02499   0.01576  -0.0487   0.8661   0.8905
  -0.750   0.0951   0.02501   0.01562  -0.0508   0.8599   0.9080
  -0.500   0.1448   0.02502   0.01544  -0.0552   0.8559   0.9241
  -0.250   0.1824   0.02516   0.01546  -0.0580   0.8474   0.9454
   0.000   0.2357   0.02516   0.01531  -0.0633   0.8426   0.9672
   0.250   0.2677   0.02530   0.01532  -0.0653   0.8334   1.0000
   0.500   0.3031   0.02531   0.01520  -0.0673   0.8272   1.0000
   0.750   0.3225   0.02557   0.01535  -0.0668   0.8167   1.0000
   1.000   0.3566   0.02568   0.01534  -0.0684   0.8099   1.0000
   1.250   0.3828   0.02592   0.01549  -0.0688   0.8004   1.0000
   1.500   0.4097   0.02617   0.01566  -0.0691   0.7912   1.0000
   1.750   0.4449   0.02623   0.01566  -0.0705   0.7841   1.0000
   2.000   0.4688   0.02653   0.01592  -0.0703   0.7733   1.0000
   2.250   0.5002   0.02661   0.01596  -0.0709   0.7640   1.0000
   2.500   0.5345   0.02653   0.01586  -0.0717   0.7548   1.0000
   2.750   0.5604   0.02665   0.01597  -0.0713   0.7424   1.0000
   3.000   0.5879   0.02671   0.01603  -0.0710   0.7305   1.0000
   3.250   0.6184   0.02667   0.01603  -0.0711   0.7198   1.0000
   3.500   0.6531   0.02645   0.01583  -0.0717   0.7101   1.0000
   3.750   0.6784   0.02655   0.01597  -0.0710   0.6970   1.0000
   4.250   0.7319   0.02656   0.01611  -0.0699   0.6699   1.0000
   4.500   0.7594   0.02647   0.01608  -0.0693   0.6554   1.0000
   4.750   0.7870   0.02634   0.01602  -0.0686   0.6400   1.0000
   5.000   0.8151   0.02616   0.01592  -0.0679   0.6236   1.0000
   5.250   0.8397   0.02613   0.01596  -0.0668   0.6050   1.0000
   5.500   0.8621   0.02620   0.01611  -0.0655   0.5846   1.0000
   5.750   0.8896   0.02605   0.01603  -0.0647   0.5646   1.0000
   6.000   0.9111   0.02623   0.01627  -0.0633   0.5414   1.0000
   6.250   0.9365   0.02626   0.01630  -0.0623   0.5180   1.0000
   6.500   0.9596   0.02645   0.01649  -0.0610   0.4927   1.0000
   6.750   0.9811   0.02679   0.01685  -0.0597   0.4666   1.0000
   7.000   1.0020   0.02723   0.01724  -0.0583   0.4406   1.0000
   7.250   1.0222   0.02779   0.01772  -0.0570   0.4156   1.0000
   7.500   1.0402   0.02851   0.01840  -0.0554   0.3913   1.0000
   7.750   1.0573   0.02933   0.01923  -0.0539   0.3685   1.0000
   8.000   1.0740   0.03020   0.02007  -0.0524   0.3472   1.0000
   8.250   1.0901   0.03113   0.02099  -0.0509   0.3275   1.0000
   8.500   1.1048   0.03213   0.02201  -0.0493   0.3087   1.0000
   8.750   1.1194   0.03316   0.02305  -0.0477   0.2912   1.0000
   9.000   1.1339   0.03425   0.02415  -0.0461   0.2748   1.0000
   9.250   1.1474   0.03539   0.02534  -0.0445   0.2589   1.0000
   9.500   1.1601   0.03658   0.02659  -0.0429   0.2438   1.0000
   9.750   1.1707   0.03785   0.02792  -0.0411   0.2292   1.0000
  10.000   1.1804   0.03919   0.02935  -0.0394   0.2154   1.0000
  10.250   1.1883   0.04063   0.03089  -0.0377   0.2021   1.0000
  10.500   1.1960   0.04218   0.03255  -0.0360   0.1895   1.0000
  10.750   1.2028   0.04383   0.03432  -0.0344   0.1776   1.0000
  11.000   1.2089   0.04558   0.03619  -0.0329   0.1665   1.0000
  11.250   1.2159   0.04733   0.03798  -0.0315   0.1564   1.0000
  11.500   1.2197   0.04937   0.04017  -0.0302   0.1464   1.0000
  11.750   1.2232   0.05162   0.04260  -0.0290   0.1371   1.0000
  12.000   1.2302   0.05363   0.04461  -0.0279   0.1295   1.0000
  12.250   1.2316   0.05630   0.04757  -0.0269   0.1218   1.0000
  12.500   1.2348   0.05867   0.04999  -0.0261   0.1153   1.0000
  12.750   1.2324   0.06178   0.05339  -0.0254   0.1090   1.0000
  13.000   1.2376   0.06404   0.05559  -0.0248   0.1036   1.0000
  13.250   1.2312   0.06807   0.06003  -0.0245   0.0992   1.0000
  13.500   1.2268   0.07168   0.06384  -0.0245   0.0950   1.0000
  13.750   1.2297   0.07432   0.06644  -0.0243   0.0905   1.0000
  14.000   1.2135   0.07981   0.07233  -0.0254   0.0879   1.0000
  14.250   1.1969   0.08560   0.07842  -0.0270   0.0855   1.0000
  14.500   1.1807   0.09159   0.08462  -0.0290   0.0834   1.0000
  14.750   1.1675   0.09743   0.09068  -0.0312   0.0814   1.0000
  15.000   1.1725   0.10028   0.09352  -0.0315   0.0782   1.0000
  15.250   1.1436   0.10957   0.10306  -0.0363   0.0780   1.0000
  15.500   1.1102   0.12080   0.11442  -0.0428   0.0782   1.0000
<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)