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

HQ 2.0/10 AIRFOIL (hq2010-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.0/10 AIRFOIL (hq2010-il)
Reynolds number: 1,000,000
Max Cl/Cd: 102.17 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq2010-il-1000000.txt
Download as CSV file: xf-hq2010-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/10 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6413   0.04310   0.04128  -0.0573   1.0000   0.0063
  -9.250  -0.6751   0.03636   0.03417  -0.0560   1.0000   0.0064
  -9.000  -0.6772   0.03390   0.03152  -0.0537   1.0000   0.0063
  -8.750  -0.6637   0.02793   0.02502  -0.0569   0.9954   0.0064
  -8.500  -0.6409   0.02401   0.02065  -0.0591   0.9901   0.0064
  -8.250  -0.6140   0.02105   0.01732  -0.0610   0.9859   0.0064
  -8.000  -0.5848   0.01815   0.01403  -0.0629   0.9829   0.0065
  -7.750  -0.5580   0.01611   0.01175  -0.0638   0.9761   0.0067
  -7.500  -0.5267   0.01506   0.01059  -0.0651   0.9709   0.0071
  -7.250  -0.4984   0.01412   0.00951  -0.0656   0.9626   0.0074
  -7.000  -0.4709   0.01316   0.00841  -0.0658   0.9532   0.0076
  -6.750  -0.4458   0.01238   0.00749  -0.0654   0.9416   0.0080
  -6.500  -0.4213   0.01172   0.00672  -0.0648   0.9296   0.0085
  -6.250  -0.3965   0.01117   0.00605  -0.0642   0.9187   0.0089
  -6.000  -0.3707   0.01083   0.00563  -0.0638   0.9095   0.0093
  -5.750  -0.3469   0.00999   0.00460  -0.0631   0.9005   0.0102
  -5.500  -0.3213   0.00950   0.00403  -0.0627   0.8927   0.0112
  -5.250  -0.2950   0.00920   0.00366  -0.0624   0.8856   0.0124
  -5.000  -0.2680   0.00896   0.00338  -0.0622   0.8789   0.0138
  -4.750  -0.2417   0.00856   0.00294  -0.0619   0.8723   0.0186
  -4.500  -0.2147   0.00828   0.00268  -0.0617   0.8661   0.0281
  -4.250  -0.1877   0.00807   0.00247  -0.0616   0.8598   0.0367
  -4.000  -0.1602   0.00794   0.00229  -0.0615   0.8538   0.0422
  -3.750  -0.1328   0.00776   0.00213  -0.0614   0.8473   0.0508
  -3.500  -0.1056   0.00758   0.00195  -0.0613   0.8411   0.0633
  -3.250  -0.0783   0.00733   0.00178  -0.0612   0.8341   0.0887
  -3.000  -0.0519   0.00697   0.00161  -0.0611   0.8277   0.1484
  -2.750  -0.0251   0.00659   0.00146  -0.0611   0.8208   0.2192
  -2.250   0.0268   0.00560   0.00116  -0.0609   0.8058   0.4468
  -2.000   0.0536   0.00538   0.00109  -0.0607   0.7963   0.5129
  -1.750   0.0809   0.00527   0.00104  -0.0605   0.7867   0.5530
  -1.500   0.1087   0.00520   0.00101  -0.0605   0.7777   0.5817
  -1.250   0.1363   0.00515   0.00099  -0.0603   0.7688   0.6102
  -1.000   0.1640   0.00513   0.00097  -0.0602   0.7593   0.6324
  -0.750   0.1918   0.00511   0.00096  -0.0601   0.7492   0.6523
  -0.500   0.2195   0.00510   0.00096  -0.0600   0.7394   0.6729
  -0.250   0.2471   0.00512   0.00096  -0.0599   0.7295   0.6881
   0.000   0.2751   0.00513   0.00097  -0.0598   0.7185   0.7008
   0.250   0.3029   0.00514   0.00098  -0.0597   0.7080   0.7121
   0.500   0.3306   0.00515   0.00099  -0.0596   0.6982   0.7241
   0.750   0.3582   0.00519   0.00102  -0.0595   0.6876   0.7368
   1.000   0.3859   0.00521   0.00104  -0.0594   0.6755   0.7479
   1.500   0.4412   0.00529   0.00111  -0.0592   0.6506   0.7660
   1.750   0.4686   0.00535   0.00115  -0.0591   0.6358   0.7751
   2.000   0.4959   0.00543   0.00119  -0.0589   0.6195   0.7851
   2.250   0.5229   0.00550   0.00125  -0.0587   0.6011   0.7951
   2.500   0.5497   0.00559   0.00132  -0.0585   0.5792   0.8057
   2.750   0.5760   0.00572   0.00139  -0.0582   0.5525   0.8173
   3.000   0.6015   0.00590   0.00150  -0.0577   0.5194   0.8302
   3.250   0.6263   0.00613   0.00162  -0.0572   0.4744   0.8449
   3.500   0.6484   0.00658   0.00181  -0.0562   0.4008   0.8633
   3.750   0.6706   0.00691   0.00200  -0.0551   0.3448   0.8911
   4.000   0.6987   0.00711   0.00220  -0.0553   0.3027   0.9616
   4.250   0.7313   0.00741   0.00238  -0.0566   0.2750   1.0000
   4.500   0.7573   0.00766   0.00256  -0.0564   0.2571   1.0000
   4.750   0.7832   0.00793   0.00274  -0.0561   0.2390   1.0000
   5.000   0.8091   0.00818   0.00293  -0.0559   0.2204   1.0000
   5.250   0.8350   0.00842   0.00312  -0.0556   0.2037   1.0000
   5.500   0.8603   0.00872   0.00333  -0.0553   0.1827   1.0000
   5.750   0.8843   0.00915   0.00359  -0.0548   0.1473   1.0000
   6.000   0.9068   0.00974   0.00396  -0.0541   0.1066   1.0000
   6.250   0.9294   0.01031   0.00436  -0.0534   0.0764   1.0000
   6.500   0.9533   0.01075   0.00471  -0.0529   0.0592   1.0000
   6.750   0.9775   0.01114   0.00504  -0.0524   0.0483   1.0000
   7.000   1.0015   0.01154   0.00540  -0.0519   0.0399   1.0000
   7.250   1.0263   0.01185   0.00571  -0.0515   0.0357   1.0000
   7.500   1.0501   0.01226   0.00610  -0.0509   0.0296   1.0000
   7.750   1.0747   0.01255   0.00640  -0.0505   0.0263   1.0000
   8.000   1.0979   0.01299   0.00679  -0.0499   0.0193   1.0000
   8.250   1.1180   0.01375   0.00743  -0.0488   0.0075   1.0000
   8.500   1.1401   0.01429   0.00798  -0.0480   0.0056   1.0000
   8.750   1.1621   0.01481   0.00855  -0.0472   0.0051   1.0000
   9.000   1.1837   0.01536   0.00916  -0.0463   0.0047   1.0000
   9.250   1.2045   0.01597   0.00985  -0.0453   0.0043   1.0000
   9.500   1.2240   0.01669   0.01066  -0.0441   0.0041   1.0000
   9.750   1.2426   0.01745   0.01151  -0.0427   0.0040   1.0000
  10.000   1.2614   0.01813   0.01228  -0.0415   0.0039   1.0000
  10.250   1.2793   0.01885   0.01307  -0.0402   0.0037   1.0000
  10.500   1.2953   0.01968   0.01400  -0.0385   0.0037   1.0000
  10.750   1.3102   0.02053   0.01494  -0.0368   0.0034   1.0000
  11.000   1.3218   0.02144   0.01595  -0.0345   0.0034   1.0000
  11.250   1.3299   0.02240   0.01701  -0.0317   0.0032   1.0000
  11.500   1.3370   0.02342   0.01815  -0.0289   0.0032   1.0000
  11.750   1.3421   0.02462   0.01946  -0.0261   0.0030   1.0000
  12.000   1.3457   0.02597   0.02092  -0.0234   0.0030   1.0000
  12.250   1.3489   0.02742   0.02249  -0.0210   0.0029   1.0000
  12.500   1.3498   0.02913   0.02432  -0.0186   0.0028   1.0000
  12.750   1.3512   0.03090   0.02622  -0.0166   0.0029   1.0000
  13.000   1.3494   0.03306   0.02850  -0.0147   0.0028   1.0000
  13.250   1.3457   0.03552   0.03110  -0.0132   0.0028   1.0000
  13.500   1.3449   0.03787   0.03359  -0.0121   0.0028   1.0000
  13.750   1.3395   0.04082   0.03668  -0.0112   0.0028   1.0000
  14.000   1.3310   0.04435   0.04036  -0.0109   0.0027   1.0000
  14.250   1.3233   0.04803   0.04419  -0.0110   0.0027   1.0000
  14.500   1.3153   0.05200   0.04830  -0.0116   0.0027   1.0000
  14.750   1.3045   0.05665   0.05311  -0.0128   0.0028   1.0000
  15.000   1.2909   0.06207   0.05869  -0.0148   0.0028   1.0000
  15.250   1.2771   0.06797   0.06475  -0.0174   0.0027   1.0000
  15.500   1.2651   0.07401   0.07095  -0.0204   0.0028   1.0000
  15.750   1.2468   0.08166   0.07876  -0.0245   0.0028   1.0000
  16.000   1.2301   0.08943   0.08668  -0.0289   0.0028   1.0000
  16.250   1.2096   0.09842   0.09583  -0.0340   0.0028   1.0000
  16.500   1.1897   0.10754   0.10511  -0.0393   0.0029   1.0000
  16.750   1.1703   0.11680   0.11452  -0.0446   0.0029   1.0000
  17.000   1.1492   0.12670   0.12454  -0.0504   0.0029   1.0000
  17.250   1.1297   0.13646   0.13443  -0.0561   0.0030   1.0000
  17.500   1.1095   0.14682   0.14491  -0.0621   0.0030   1.0000
  17.750   1.0883   0.15789   0.15611  -0.0686   0.0031   1.0000
  18.000   1.0628   0.17092   0.16927  -0.0760   0.0032   1.0000
<< Back to HQ 2.0/10 AIRFOIL (hq2010-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.0/10 AIRFOIL (hq2010-il)