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

HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il)
Reynolds number: 1,000,000
Max Cl/Cd: 92.51 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq2510-il-1000000-n5.txt
Download as CSV file: xf-hq2510-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/10 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4490   0.11444   0.11275  -0.0238   1.0000   0.0023
 -10.000  -0.7038   0.02739   0.02474  -0.0744   0.9825   0.0019
  -9.750  -0.6837   0.02363   0.02056  -0.0767   0.9753   0.0019
  -9.500  -0.6586   0.02144   0.01808  -0.0781   0.9688   0.0021
  -9.250  -0.6334   0.01969   0.01604  -0.0791   0.9605   0.0021
  -9.000  -0.6091   0.01836   0.01451  -0.0794   0.9511   0.0021
  -8.750  -0.5886   0.01634   0.01214  -0.0792   0.9398   0.0023
  -8.500  -0.5658   0.01518   0.01077  -0.0788   0.9291   0.0025
  -8.250  -0.5413   0.01454   0.01002  -0.0785   0.9188   0.0027
  -8.000  -0.5165   0.01398   0.00935  -0.0782   0.9095   0.0028
  -7.750  -0.4915   0.01346   0.00872  -0.0779   0.9013   0.0030
  -7.500  -0.4663   0.01295   0.00809  -0.0776   0.8929   0.0032
  -7.250  -0.4410   0.01241   0.00744  -0.0774   0.8853   0.0033
  -7.000  -0.4156   0.01190   0.00679  -0.0771   0.8775   0.0036
  -6.750  -0.3897   0.01144   0.00624  -0.0768   0.8703   0.0038
  -6.500  -0.3637   0.01103   0.00572  -0.0766   0.8633   0.0040
  -6.250  -0.3372   0.01065   0.00525  -0.0764   0.8565   0.0042
  -6.000  -0.3113   0.01016   0.00466  -0.0761   0.8498   0.0046
  -5.750  -0.2847   0.00978   0.00424  -0.0760   0.8434   0.0053
  -5.500  -0.2578   0.00954   0.00394  -0.0759   0.8367   0.0060
  -5.250  -0.2307   0.00927   0.00359  -0.0758   0.8306   0.0067
  -5.000  -0.2036   0.00902   0.00328  -0.0757   0.8238   0.0071
  -4.750  -0.1763   0.00877   0.00296  -0.0756   0.8176   0.0076
  -4.500  -0.1491   0.00849   0.00262  -0.0755   0.8107   0.0095
  -4.250  -0.1217   0.00828   0.00237  -0.0754   0.8042   0.0117
  -4.000  -0.0943   0.00805   0.00216  -0.0754   0.7972   0.0184
  -3.750  -0.0668   0.00788   0.00198  -0.0754   0.7904   0.0276
  -3.500  -0.0390   0.00778   0.00183  -0.0753   0.7827   0.0302
  -3.250  -0.0114   0.00765   0.00169  -0.0753   0.7749   0.0359
  -2.750   0.0438   0.00742   0.00144  -0.0752   0.7565   0.0543
  -2.500   0.0714   0.00727   0.00132  -0.0752   0.7481   0.0727
  -2.250   0.0990   0.00715   0.00122  -0.0752   0.7400   0.0923
  -2.000   0.1267   0.00699   0.00112  -0.0752   0.7319   0.1220
  -1.750   0.1539   0.00679   0.00103  -0.0752   0.7221   0.1720
  -1.500   0.1811   0.00660   0.00095  -0.0752   0.7111   0.2277
  -1.250   0.2082   0.00637   0.00087  -0.0752   0.6993   0.2930
  -1.000   0.2348   0.00604   0.00080  -0.0752   0.6869   0.3985
  -0.750   0.2620   0.00590   0.00077  -0.0752   0.6742   0.4554
  -0.500   0.2891   0.00576   0.00077  -0.0751   0.6626   0.5172
  -0.250   0.3158   0.00560   0.00081  -0.0750   0.6524   0.6004
   0.000   0.3431   0.00559   0.00084  -0.0748   0.6403   0.6339
   0.250   0.3707   0.00563   0.00086  -0.0748   0.6261   0.6488
   0.500   0.3981   0.00568   0.00089  -0.0747   0.6090   0.6628
   0.750   0.4253   0.00576   0.00092  -0.0745   0.5905   0.6744
   1.000   0.4526   0.00586   0.00096  -0.0744   0.5718   0.6831
   1.250   0.4798   0.00595   0.00101  -0.0743   0.5522   0.6906
   1.500   0.5066   0.00610   0.00107  -0.0741   0.5270   0.6979
   1.750   0.5326   0.00631   0.00116  -0.0738   0.4916   0.7057
   2.000   0.5584   0.00655   0.00126  -0.0735   0.4536   0.7132
   2.250   0.5844   0.00679   0.00139  -0.0732   0.4209   0.7214
   2.500   0.6103   0.00701   0.00151  -0.0729   0.3903   0.7297
   2.750   0.6365   0.00722   0.00164  -0.0727   0.3645   0.7383
   3.000   0.6623   0.00745   0.00178  -0.0724   0.3342   0.7477
   3.250   0.6879   0.00770   0.00195  -0.0721   0.3040   0.7571
   3.500   0.7135   0.00795   0.00212  -0.0718   0.2784   0.7669
   4.000   0.7649   0.00838   0.00245  -0.0712   0.2380   0.7885
   4.250   0.7910   0.00855   0.00263  -0.0709   0.2244   0.8004
   4.500   0.8154   0.00885   0.00284  -0.0704   0.1943   0.8143
   4.750   0.8380   0.00931   0.00312  -0.0697   0.1505   0.8304
   5.000   0.8614   0.00964   0.00337  -0.0690   0.1238   0.8505
   5.250   0.8834   0.00988   0.00363  -0.0679   0.1025   0.8875
   5.500   0.9116   0.01031   0.00401  -0.0684   0.0666   1.0000
   5.750   0.9340   0.01090   0.00442  -0.0677   0.0367   1.0000
   6.000   0.9541   0.01176   0.00508  -0.0665   0.0043   1.0000
   6.250   0.9788   0.01208   0.00543  -0.0660   0.0031   1.0000
   6.500   1.0036   0.01239   0.00577  -0.0656   0.0029   1.0000
   6.750   1.0281   0.01271   0.00613  -0.0651   0.0027   1.0000
   7.000   1.0523   0.01307   0.00652  -0.0646   0.0025   1.0000
   7.250   1.0760   0.01346   0.00698  -0.0640   0.0023   1.0000
   7.500   1.0992   0.01389   0.00745  -0.0634   0.0021   1.0000
   7.750   1.1216   0.01439   0.00800  -0.0626   0.0019   1.0000
   8.000   1.1437   0.01490   0.00856  -0.0618   0.0018   1.0000
   8.250   1.1646   0.01553   0.00925  -0.0607   0.0016   1.0000
   8.500   1.1837   0.01632   0.01013  -0.0594   0.0015   1.0000
   8.750   1.1973   0.01762   0.01158  -0.0573   0.0014   1.0000
   9.000   1.2148   0.01842   0.01246  -0.0558   0.0014   1.0000
   9.250   1.2317   0.01924   0.01338  -0.0543   0.0014   1.0000
   9.500   1.2478   0.02007   0.01429  -0.0526   0.0013   1.0000
   9.750   1.2619   0.02098   0.01529  -0.0507   0.0013   1.0000
  10.000   1.2724   0.02191   0.01631  -0.0482   0.0013   1.0000
  10.250   1.2799   0.02297   0.01747  -0.0453   0.0013   1.0000
  10.500   1.2870   0.02406   0.01866  -0.0425   0.0013   1.0000
  10.750   1.2931   0.02527   0.01998  -0.0398   0.0013   1.0000
  11.000   1.2979   0.02664   0.02146  -0.0371   0.0013   1.0000
  11.250   1.3020   0.02813   0.02306  -0.0346   0.0013   1.0000
  11.500   1.3049   0.02977   0.02482  -0.0323   0.0013   1.0000
  11.750   1.3060   0.03167   0.02685  -0.0301   0.0013   1.0000
  12.000   1.3082   0.03354   0.02884  -0.0283   0.0013   1.0000
  12.250   1.3076   0.03578   0.03122  -0.0265   0.0013   1.0000
  12.500   1.3052   0.03831   0.03389  -0.0251   0.0013   1.0000
  12.750   1.3021   0.04104   0.03679  -0.0239   0.0013   1.0000
  13.000   1.2967   0.04419   0.04009  -0.0231   0.0013   1.0000
  13.250   1.2965   0.04684   0.04286  -0.0228   0.0012   1.0000
  13.500   1.2896   0.05050   0.04667  -0.0228   0.0012   1.0000
  13.750   1.2811   0.05461   0.05094  -0.0233   0.0012   1.0000
  14.000   1.2717   0.05911   0.05559  -0.0243   0.0012   1.0000
  14.250   1.2617   0.06400   0.06062  -0.0259   0.0012   1.0000
  14.500   1.2471   0.07005   0.06684  -0.0283   0.0012   1.0000
  14.750   1.2376   0.07565   0.07257  -0.0310   0.0012   1.0000
  15.000   1.2295   0.08136   0.07841  -0.0340   0.0012   1.0000
  15.250   1.2148   0.08870   0.08590  -0.0380   0.0012   1.0000
  15.500   1.1947   0.09752   0.09489  -0.0429   0.0012   1.0000
  15.750   1.1841   0.10473   0.10221  -0.0471   0.0012   1.0000
<< Back to HQ 2.5/10 AIRFOIL (hq2510-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/10 AIRFOIL (hq2510-il)