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

HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il)
Reynolds number: 1,000,000
Max Cl/Cd: 80.21 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq159b-il-1000000-n5.txt
Download as CSV file: xf-hq159b-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/9 B AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.8500   0.02863   0.02579  -0.0479   1.0000   0.0022
 -10.000  -0.8495   0.02427   0.02094  -0.0465   1.0000   0.0023
  -9.750  -0.8413   0.02107   0.01731  -0.0450   1.0000   0.0024
  -9.500  -0.8251   0.01936   0.01538  -0.0439   1.0000   0.0025
  -9.250  -0.8065   0.01816   0.01401  -0.0429   1.0000   0.0025
  -9.000  -0.7819   0.01714   0.01285  -0.0430   0.9980   0.0027
  -8.750  -0.7530   0.01626   0.01187  -0.0439   0.9928   0.0028
  -8.500  -0.7232   0.01546   0.01093  -0.0449   0.9882   0.0030
  -8.250  -0.6925   0.01457   0.00991  -0.0460   0.9833   0.0031
  -8.000  -0.6621   0.01380   0.00903  -0.0470   0.9775   0.0033
  -7.750  -0.6309   0.01315   0.00827  -0.0481   0.9713   0.0035
  -7.500  -0.6009   0.01249   0.00751  -0.0489   0.9638   0.0037
  -7.250  -0.5718   0.01191   0.00682  -0.0495   0.9552   0.0038
  -7.000  -0.5451   0.01143   0.00624  -0.0494   0.9440   0.0040
  -6.750  -0.5190   0.01101   0.00574  -0.0491   0.9321   0.0041
  -6.500  -0.4931   0.01066   0.00527  -0.0488   0.9199   0.0042
  -6.250  -0.4681   0.01012   0.00459  -0.0483   0.9070   0.0047
  -6.000  -0.4425   0.00973   0.00412  -0.0479   0.8943   0.0056
  -5.750  -0.4165   0.00944   0.00376  -0.0475   0.8831   0.0064
  -5.500  -0.3902   0.00919   0.00343  -0.0473   0.8733   0.0073
  -5.250  -0.3635   0.00893   0.00312  -0.0470   0.8640   0.0098
  -5.000  -0.3369   0.00871   0.00287  -0.0468   0.8547   0.0138
  -4.750  -0.3101   0.00852   0.00266  -0.0466   0.8451   0.0187
  -4.500  -0.2828   0.00838   0.00245  -0.0465   0.8360   0.0212
  -4.250  -0.2557   0.00820   0.00227  -0.0464   0.8279   0.0253
  -4.000  -0.2283   0.00806   0.00209  -0.0463   0.8193   0.0278
  -3.750  -0.2008   0.00793   0.00193  -0.0463   0.8109   0.0311
  -3.500  -0.1735   0.00778   0.00178  -0.0462   0.8022   0.0375
  -3.250  -0.1459   0.00763   0.00162  -0.0461   0.7935   0.0458
  -3.000  -0.1183   0.00752   0.00149  -0.0461   0.7850   0.0535
  -2.500  -0.0637   0.00724   0.00124  -0.0459   0.7621   0.0860
  -2.000  -0.0096   0.00680   0.00100  -0.0458   0.7383   0.1775
  -1.750   0.0178   0.00664   0.00091  -0.0458   0.7276   0.2147
  -1.500   0.0442   0.00625   0.00080  -0.0458   0.7169   0.3110
  -1.250   0.0704   0.00582   0.00071  -0.0457   0.7061   0.4323
  -1.000   0.0977   0.00565   0.00067  -0.0457   0.6952   0.4846
  -0.750   0.1247   0.00546   0.00065  -0.0456   0.6839   0.5521
  -0.500   0.1520   0.00539   0.00065  -0.0454   0.6709   0.5948
  -0.250   0.1793   0.00536   0.00067  -0.0453   0.6547   0.6289
   0.000   0.2068   0.00538   0.00068  -0.0452   0.6383   0.6491
   0.250   0.2345   0.00540   0.00070  -0.0451   0.6237   0.6661
   0.500   0.2622   0.00544   0.00072  -0.0451   0.6094   0.6787
   0.750   0.2899   0.00550   0.00074  -0.0450   0.5929   0.6890
   1.000   0.3175   0.00556   0.00077  -0.0449   0.5722   0.6983
   1.250   0.3434   0.00582   0.00084  -0.0446   0.5144   0.7074
   1.500   0.3687   0.00619   0.00095  -0.0443   0.4472   0.7153
   1.750   0.3949   0.00645   0.00106  -0.0440   0.4003   0.7244
   2.000   0.4198   0.00688   0.00120  -0.0437   0.3198   0.7351
   2.250   0.4453   0.00725   0.00137  -0.0434   0.2637   0.7450
   2.500   0.4719   0.00745   0.00149  -0.0432   0.2374   0.7530
   2.750   0.4988   0.00762   0.00161  -0.0431   0.2187   0.7613
   3.000   0.5257   0.00777   0.00174  -0.0430   0.2018   0.7709
   3.250   0.5525   0.00791   0.00187  -0.0428   0.1862   0.7812
   3.500   0.5782   0.00820   0.00203  -0.0425   0.1490   0.7936
   3.750   0.6029   0.00858   0.00227  -0.0421   0.1047   0.8096
   4.000   0.6280   0.00886   0.00251  -0.0417   0.0790   0.8287
   4.250   0.6527   0.00912   0.00276  -0.0411   0.0546   0.8508
   4.500   0.6759   0.00947   0.00305  -0.0403   0.0280   0.8798
   4.750   0.6996   0.00953   0.00326  -0.0393   0.0252   0.9318
   5.000   0.7362   0.00973   0.00351  -0.0414   0.0239   1.0000
   5.250   0.7626   0.00997   0.00377  -0.0412   0.0231   1.0000
   5.500   0.7887   0.01024   0.00407  -0.0409   0.0219   1.0000
   5.750   0.8145   0.01054   0.00439  -0.0407   0.0202   1.0000
   6.000   0.8400   0.01089   0.00478  -0.0403   0.0192   1.0000
   6.250   0.8663   0.01109   0.00499  -0.0402   0.0189   1.0000
   6.500   0.8926   0.01129   0.00520  -0.0400   0.0183   1.0000
   6.750   0.9186   0.01152   0.00546  -0.0398   0.0173   1.0000
   7.000   0.9443   0.01178   0.00574  -0.0396   0.0155   1.0000
   7.250   0.9697   0.01209   0.00604  -0.0393   0.0130   1.0000
   7.500   0.9945   0.01245   0.00641  -0.0389   0.0106   1.0000
   7.750   1.0192   0.01283   0.00680  -0.0385   0.0081   1.0000
   8.000   1.0432   0.01329   0.00725  -0.0380   0.0058   1.0000
   8.250   1.0673   0.01372   0.00773  -0.0376   0.0053   1.0000
   8.500   1.0910   0.01417   0.00823  -0.0370   0.0048   1.0000
   8.750   1.1142   0.01469   0.00878  -0.0365   0.0043   1.0000
   9.000   1.1368   0.01526   0.00940  -0.0358   0.0038   1.0000
   9.250   1.1586   0.01593   0.01015  -0.0350   0.0034   1.0000
   9.500   1.1792   0.01671   0.01105  -0.0340   0.0032   1.0000
   9.750   1.2011   0.01728   0.01169  -0.0333   0.0031   1.0000
  10.000   1.2219   0.01796   0.01246  -0.0324   0.0030   1.0000
  10.250   1.2424   0.01863   0.01322  -0.0315   0.0029   1.0000
  10.500   1.2621   0.01936   0.01403  -0.0306   0.0027   1.0000
  10.750   1.2812   0.02009   0.01485  -0.0295   0.0025   1.0000
  11.000   1.2998   0.02085   0.01570  -0.0285   0.0024   1.0000
  11.250   1.3169   0.02169   0.01664  -0.0272   0.0023   1.0000
  11.500   1.3341   0.02247   0.01749  -0.0260   0.0021   1.0000
  11.750   1.3484   0.02344   0.01856  -0.0244   0.0020   1.0000
  12.000   1.3623   0.02436   0.01956  -0.0228   0.0020   1.0000
  12.250   1.3693   0.02550   0.02083  -0.0202   0.0019   1.0000
  12.500   1.3715   0.02686   0.02232  -0.0170   0.0018   1.0000
  12.750   1.3704   0.02850   0.02412  -0.0139   0.0017   1.0000
  13.000   1.3681   0.03034   0.02612  -0.0112   0.0017   1.0000
  13.250   1.3616   0.03269   0.02865  -0.0087   0.0016   1.0000
  13.500   1.3609   0.03471   0.03080  -0.0071   0.0016   1.0000
  13.750   1.3598   0.03693   0.03317  -0.0059   0.0016   1.0000
  14.000   1.3578   0.03938   0.03576  -0.0051   0.0016   1.0000
  14.250   1.3526   0.04238   0.03890  -0.0048   0.0016   1.0000
  14.500   1.3449   0.04594   0.04262  -0.0050   0.0016   1.0000
  14.750   1.3262   0.05126   0.04814  -0.0062   0.0016   1.0000
  15.000   1.3175   0.05572   0.05273  -0.0079   0.0016   1.0000
  15.250   1.3011   0.06184   0.05901  -0.0109   0.0016   1.0000
  15.500   1.2864   0.06836   0.06568  -0.0145   0.0016   1.0000
  15.750   1.2671   0.07640   0.07388  -0.0194   0.0016   1.0000
  16.000   1.2325   0.08844   0.08612  -0.0268   0.0016   1.0000
  16.250   1.1808   0.10508   0.10296  -0.0367   0.0016   1.0000
<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)

Polar data table (+)

Polar graphs


<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)