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

HQ 2.5/12 AIRFOIL (hq2512-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/12 AIRFOIL (hq2512-il)
Reynolds number: 500,000
Max Cl/Cd: 100.53 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq2512-il-500000.txt
Download as CSV file: xf-hq2512-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/12 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.4130   0.09691   0.09463  -0.0433   1.0000   0.0236
 -10.500  -0.5896   0.05034   0.04793  -0.0690   1.0000   0.0120
 -10.250  -0.6196   0.04534   0.04281  -0.0708   1.0000   0.0118
 -10.000  -0.6528   0.04324   0.04067  -0.0668   1.0000   0.0117
  -9.750  -0.6799   0.03686   0.03384  -0.0675   0.9950   0.0115
  -9.500  -0.6719   0.02984   0.02610  -0.0717   0.9869   0.0116
  -9.250  -0.6471   0.02616   0.02194  -0.0745   0.9832   0.0119
  -9.000  -0.6216   0.02363   0.01904  -0.0760   0.9780   0.0123
  -8.750  -0.5920   0.02164   0.01675  -0.0777   0.9742   0.0126
  -8.500  -0.5585   0.02041   0.01529  -0.0796   0.9714   0.0130
  -8.250  -0.5321   0.01801   0.01267  -0.0808   0.9665   0.0141
  -8.000  -0.5029   0.01719   0.01178  -0.0816   0.9604   0.0150
  -7.750  -0.4719   0.01625   0.01072  -0.0827   0.9559   0.0160
  -7.500  -0.4468   0.01545   0.00979  -0.0824   0.9475   0.0168
  -7.250  -0.4200   0.01452   0.00872  -0.0825   0.9412   0.0181
  -7.000  -0.3973   0.01376   0.00792  -0.0817   0.9320   0.0199
  -6.750  -0.3707   0.01332   0.00742  -0.0815   0.9254   0.0221
  -6.500  -0.3472   0.01273   0.00678  -0.0808   0.9168   0.0257
  -6.250  -0.3204   0.01253   0.00655  -0.0806   0.9101   0.0302
  -6.000  -0.2963   0.01200   0.00597  -0.0800   0.9020   0.0357
  -5.750  -0.2694   0.01186   0.00576  -0.0798   0.8955   0.0402
  -5.500  -0.2448   0.01140   0.00527  -0.0793   0.8876   0.0453
  -5.250  -0.2181   0.01121   0.00503  -0.0791   0.8813   0.0499
  -5.000  -0.1915   0.01106   0.00481  -0.0788   0.8739   0.0531
  -4.750  -0.1661   0.01062   0.00434  -0.0784   0.8678   0.0601
  -4.500  -0.1394   0.01045   0.00413  -0.0782   0.8605   0.0653
  -4.250  -0.1134   0.01005   0.00371  -0.0779   0.8544   0.0740
  -4.000  -0.0871   0.00975   0.00342  -0.0776   0.8474   0.0867
  -3.750  -0.0608   0.00940   0.00313  -0.0774   0.8411   0.1122
  -3.500  -0.0356   0.00887   0.00288  -0.0772   0.8345   0.1845
  -3.250  -0.0104   0.00832   0.00264  -0.0770   0.8279   0.2783
  -3.000   0.0138   0.00765   0.00244  -0.0768   0.8216   0.4196
  -2.750   0.0395   0.00733   0.00238  -0.0764   0.8146   0.5077
  -2.500   0.0665   0.00723   0.00235  -0.0761   0.8085   0.5551
  -2.250   0.0938   0.00717   0.00232  -0.0759   0.8012   0.5841
  -1.750   0.1490   0.00716   0.00230  -0.0755   0.7882   0.6298
  -1.500   0.1770   0.00717   0.00227  -0.0754   0.7823   0.6450
  -1.250   0.2047   0.00718   0.00228  -0.0752   0.7757   0.6602
  -1.000   0.2320   0.00721   0.00229  -0.0749   0.7685   0.6794
  -0.750   0.2593   0.00723   0.00231  -0.0746   0.7603   0.6937
  -0.500   0.2869   0.00724   0.00228  -0.0744   0.7525   0.7037
  -0.250   0.3143   0.00723   0.00228  -0.0742   0.7436   0.7127
   0.000   0.3420   0.00727   0.00226  -0.0740   0.7355   0.7226
   0.250   0.3693   0.00727   0.00227  -0.0737   0.7265   0.7321
   0.500   0.3969   0.00728   0.00227  -0.0736   0.7190   0.7405
   0.750   0.4246   0.00730   0.00227  -0.0735   0.7110   0.7477
   1.000   0.4523   0.00731   0.00228  -0.0734   0.7028   0.7544
   1.250   0.4797   0.00734   0.00229  -0.0732   0.6943   0.7616
   1.500   0.5072   0.00735   0.00231  -0.0731   0.6848   0.7687
   1.750   0.5345   0.00738   0.00234  -0.0729   0.6754   0.7762
   2.000   0.5616   0.00741   0.00236  -0.0727   0.6655   0.7838
   2.250   0.5887   0.00743   0.00240  -0.0725   0.6546   0.7920
   2.500   0.6155   0.00746   0.00244  -0.0722   0.6433   0.8001
   3.000   0.6683   0.00754   0.00255  -0.0715   0.6168   0.8184
   3.250   0.6942   0.00759   0.00260  -0.0710   0.6004   0.8285
   3.500   0.7197   0.00767   0.00268  -0.0705   0.5819   0.8399
   3.750   0.7444   0.00776   0.00275  -0.0698   0.5613   0.8527
   4.000   0.7684   0.00785   0.00285  -0.0690   0.5380   0.8681
   4.250   0.7909   0.00796   0.00296  -0.0679   0.5111   0.8899
   4.500   0.8149   0.00811   0.00311  -0.0670   0.4785   0.9327
   4.750   0.8505   0.00846   0.00331  -0.0690   0.4390   1.0000
   5.000   0.8731   0.00890   0.00356  -0.0683   0.3999   1.0000
   5.250   0.8953   0.00935   0.00384  -0.0675   0.3621   1.0000
   5.500   0.9179   0.00979   0.00412  -0.0668   0.3299   1.0000
   5.750   0.9402   0.01023   0.00442  -0.0660   0.3026   1.0000
   6.000   0.9632   0.01063   0.00473  -0.0654   0.2813   1.0000
   6.250   0.9863   0.01100   0.00505  -0.0647   0.2648   1.0000
   6.500   1.0095   0.01136   0.00536  -0.0640   0.2502   1.0000
   6.750   1.0325   0.01172   0.00568  -0.0633   0.2355   1.0000
   7.000   1.0549   0.01211   0.00601  -0.0626   0.2189   1.0000
   7.250   1.0768   0.01252   0.00634  -0.0617   0.2011   1.0000
   7.500   1.0988   0.01291   0.00670  -0.0609   0.1848   1.0000
   7.750   1.1201   0.01334   0.00706  -0.0600   0.1679   1.0000
   8.000   1.1405   0.01381   0.00746  -0.0590   0.1513   1.0000
   8.250   1.1605   0.01429   0.00788  -0.0579   0.1367   1.0000
   8.500   1.1800   0.01478   0.00833  -0.0567   0.1232   1.0000
   8.750   1.1983   0.01533   0.00882  -0.0553   0.1088   1.0000
   9.000   1.2157   0.01590   0.00933  -0.0539   0.0952   1.0000
   9.250   1.2314   0.01647   0.00986  -0.0521   0.0824   1.0000
   9.500   1.2457   0.01708   0.01041  -0.0501   0.0705   1.0000
   9.750   1.2580   0.01781   0.01107  -0.0478   0.0570   1.0000
  10.000   1.2633   0.01900   0.01205  -0.0447   0.0310   1.0000
  10.250   1.2643   0.02050   0.01340  -0.0412   0.0136   1.0000
  10.500   1.2730   0.02158   0.01451  -0.0388   0.0107   1.0000
  10.750   1.2842   0.02250   0.01553  -0.0369   0.0096   1.0000
  11.000   1.2940   0.02354   0.01665  -0.0349   0.0088   1.0000
  11.250   1.3020   0.02474   0.01794  -0.0328   0.0083   1.0000
  11.500   1.3058   0.02629   0.01961  -0.0305   0.0078   1.0000
  11.750   1.3135   0.02759   0.02100  -0.0288   0.0076   1.0000
  12.000   1.3203   0.02899   0.02250  -0.0271   0.0074   1.0000
  12.250   1.3252   0.03059   0.02422  -0.0255   0.0073   1.0000
  12.500   1.3287   0.03238   0.02612  -0.0239   0.0071   1.0000
  12.750   1.3316   0.03429   0.02813  -0.0226   0.0069   1.0000
  13.000   1.3328   0.03643   0.03039  -0.0213   0.0068   1.0000
  13.250   1.3331   0.03875   0.03283  -0.0203   0.0066   1.0000
  13.500   1.3324   0.04129   0.03547  -0.0195   0.0063   1.0000
  13.750   1.3304   0.04405   0.03836  -0.0189   0.0063   1.0000
  14.000   1.3273   0.04707   0.04148  -0.0185   0.0062   1.0000
  14.250   1.3227   0.05041   0.04494  -0.0184   0.0061   1.0000
  14.500   1.3185   0.05386   0.04851  -0.0186   0.0061   1.0000
  14.750   1.3125   0.05773   0.05250  -0.0191   0.0060   1.0000
  15.000   1.3068   0.06173   0.05663  -0.0199   0.0060   1.0000
  15.250   1.2968   0.06654   0.06155  -0.0212   0.0059   1.0000
  15.500   1.2913   0.07090   0.06605  -0.0225   0.0059   1.0000
  15.750   1.2815   0.07608   0.07134  -0.0243   0.0058   1.0000
  16.000   1.2772   0.08060   0.07600  -0.0260   0.0059   1.0000
  16.250   1.2667   0.08622   0.08174  -0.0282   0.0058   1.0000
  16.500   1.2612   0.09124   0.08689  -0.0304   0.0059   1.0000
<< Back to HQ 2.5/12 AIRFOIL (hq2512-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.5/12 AIRFOIL (hq2512-il)