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

HQ 3.5/9 AIRFOIL (hq359-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 3.5/9 AIRFOIL (hq359-il)
Reynolds number: 100,000
Max Cl/Cd: 62.19 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq359-il-100000-n5.txt
Download as CSV file: xf-hq359-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.5/9 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3695   0.09477   0.09036  -0.0357   1.0000   0.0418
  -7.500  -0.3834   0.09329   0.08898  -0.0337   1.0000   0.0418
  -7.250  -0.3940   0.09115   0.08693  -0.0335   1.0000   0.0419
  -7.000  -0.3878   0.08703   0.08283  -0.0393   0.9969   0.0421
  -6.750  -0.3650   0.08100   0.07670  -0.0494   0.9900   0.0422
  -6.500  -0.3422   0.07549   0.07106  -0.0567   0.9843   0.0423
  -6.000  -0.3123   0.06446   0.06008  -0.0587   0.9755   0.0260
  -5.750  -0.2888   0.05855   0.05401  -0.0657   0.9687   0.0238
  -5.500  -0.2576   0.05093   0.04604  -0.0742   0.9626   0.0211
  -5.000  -0.1994   0.04131   0.03571  -0.0826   0.9516   0.0210
  -4.750  -0.1669   0.03829   0.03241  -0.0861   0.9479   0.0232
  -4.500  -0.1367   0.03464   0.02831  -0.0882   0.9427   0.0235
  -4.250  -0.1041   0.03094   0.02400  -0.0901   0.9380   0.0230
  -4.000  -0.0677   0.02775   0.02023  -0.0922   0.9350   0.0229
  -3.750  -0.0329   0.02526   0.01721  -0.0936   0.9316   0.0233
  -3.500  -0.0041   0.02344   0.01502  -0.0936   0.9260   0.0240
  -3.250   0.0303   0.02175   0.01302  -0.0946   0.9227   0.0254
  -3.000   0.0664   0.02059   0.01162  -0.0960   0.9201   0.0292
  -2.750   0.0923   0.01956   0.01058  -0.0958   0.9138   0.0343
  -2.500   0.1244   0.01863   0.00949  -0.0965   0.9095   0.0408
  -2.250   0.1594   0.01773   0.00857  -0.0979   0.9064   0.0650
  -2.000   0.1855   0.01719   0.00816  -0.0979   0.8996   0.1117
  -1.750   0.2158   0.01551   0.00773  -0.0994   0.8954   0.4331
  -1.500   0.2432   0.01505   0.00788  -0.0985   0.8916   0.6709
  -1.250   0.2617   0.01500   0.00792  -0.0958   0.8835   0.7513
  -1.000   0.2867   0.01475   0.00773  -0.0941   0.8791   0.8177
  -0.750   0.3048   0.01452   0.00758  -0.0913   0.8711   0.8786
  -0.500   0.3460   0.01425   0.00723  -0.0936   0.8669   0.9344
  -0.250   0.3820   0.01420   0.00701  -0.0955   0.8599   1.0000
   0.000   0.4154   0.01418   0.00682  -0.0967   0.8533   1.0000
   0.250   0.4449   0.01422   0.00673  -0.0971   0.8452   1.0000
   0.500   0.4785   0.01414   0.00652  -0.0981   0.8384   1.0000
   0.750   0.5065   0.01414   0.00642  -0.0981   0.8277   1.0000
   1.000   0.5360   0.01407   0.00626  -0.0981   0.8168   1.0000
   1.250   0.5655   0.01397   0.00608  -0.0981   0.8050   1.0000
   1.500   0.5947   0.01389   0.00594  -0.0980   0.7926   1.0000
   1.750   0.6234   0.01384   0.00582  -0.0979   0.7802   1.0000
   2.000   0.6523   0.01381   0.00575  -0.0978   0.7690   1.0000
   2.250   0.6793   0.01387   0.00580  -0.0975   0.7566   1.0000
   2.500   0.7065   0.01392   0.00587  -0.0972   0.7437   1.0000
   2.750   0.7338   0.01397   0.00592  -0.0969   0.7300   1.0000
   3.000   0.7612   0.01402   0.00597  -0.0966   0.7153   1.0000
   3.250   0.7879   0.01409   0.00604  -0.0962   0.6985   1.0000
   3.500   0.8140   0.01419   0.00620  -0.0957   0.6799   1.0000
   3.750   0.8408   0.01427   0.00627  -0.0952   0.6597   1.0000
   4.000   0.8662   0.01442   0.00642  -0.0945   0.6356   1.0000
   4.250   0.8914   0.01459   0.00657  -0.0937   0.6081   1.0000
   4.500   0.9161   0.01481   0.00674  -0.0929   0.5759   1.0000
   4.750   0.9397   0.01511   0.00701  -0.0919   0.5383   1.0000
   5.000   0.9622   0.01553   0.00728  -0.0907   0.4966   1.0000
   5.250   0.9836   0.01606   0.00765  -0.0895   0.4549   1.0000
   5.500   1.0043   0.01667   0.00811  -0.0883   0.4163   1.0000
   5.750   1.0247   0.01732   0.00865  -0.0871   0.3831   1.0000
   6.000   1.0455   0.01797   0.00929  -0.0860   0.3546   1.0000
   6.250   1.0663   0.01863   0.00992  -0.0849   0.3296   1.0000
   6.500   1.0871   0.01930   0.01058  -0.0839   0.3074   1.0000
   6.750   1.1073   0.02001   0.01128  -0.0827   0.2878   1.0000
   7.000   1.1278   0.02071   0.01203  -0.0817   0.2694   1.0000
   7.250   1.1481   0.02141   0.01282  -0.0806   0.2519   1.0000
   7.500   1.1656   0.02216   0.01365  -0.0792   0.2241   1.0000
   7.750   1.1813   0.02297   0.01436  -0.0777   0.1818   1.0000
   8.000   1.1939   0.02419   0.01521  -0.0760   0.1220   1.0000
   8.250   1.2023   0.02610   0.01659  -0.0738   0.0613   1.0000
   8.500   1.2099   0.02820   0.01843  -0.0713   0.0277   1.0000
   8.750   1.2182   0.03009   0.02038  -0.0687   0.0196   1.0000
   9.000   1.2249   0.03180   0.02229  -0.0659   0.0171   1.0000
   9.250   1.2318   0.03340   0.02412  -0.0632   0.0151   1.0000
   9.500   1.2374   0.03507   0.02598  -0.0607   0.0132   1.0000
   9.750   1.2393   0.03705   0.02813  -0.0581   0.0119   1.0000
  10.000   1.2358   0.03954   0.03081  -0.0552   0.0111   1.0000
  10.250   1.2380   0.04165   0.03311  -0.0530   0.0107   1.0000
  10.500   1.2392   0.04398   0.03566  -0.0510   0.0104   1.0000
  10.750   1.2400   0.04653   0.03842  -0.0491   0.0102   1.0000
  11.000   1.2410   0.04923   0.04141  -0.0475   0.0099   1.0000
  11.250   1.2420   0.05211   0.04453  -0.0461   0.0097   1.0000
  11.500   1.2423   0.05522   0.04787  -0.0449   0.0096   1.0000
  11.750   1.2414   0.05860   0.05150  -0.0440   0.0094   1.0000
  12.000   1.2387   0.06225   0.05540  -0.0433   0.0093   1.0000
  12.250   1.2338   0.06626   0.05968  -0.0430   0.0092   1.0000
  12.500   1.2266   0.07066   0.06435  -0.0432   0.0092   1.0000
  12.750   1.2170   0.07552   0.06947  -0.0439   0.0092   1.0000
  13.000   1.2055   0.08085   0.07506  -0.0453   0.0091   1.0000
  13.250   1.1923   0.08670   0.08115  -0.0474   0.0092   1.0000
  13.500   1.1778   0.09314   0.08783  -0.0503   0.0092   1.0000
  13.750   1.1626   0.10018   0.09510  -0.0540   0.0092   1.0000
  14.000   1.1469   0.10788   0.10300  -0.0585   0.0093   1.0000
  14.250   1.1310   0.11624   0.11155  -0.0637   0.0094   1.0000
  14.500   1.1150   0.12525   0.12072  -0.0695   0.0095   1.0000
<< Back to HQ 3.5/9 AIRFOIL (hq359-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.5/9 AIRFOIL (hq359-il)