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

HQ 3.5/12 AIRFOIL (hq3512-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 3.5/12 AIRFOIL (hq3512-il)
Reynolds number: 100,000
Max Cl/Cd: 57.82 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq3512-il-100000-n5.txt
Download as CSV file: xf-hq3512-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.5/12 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3561   0.09346   0.08880  -0.0460   1.0000   0.0266
  -9.000  -0.3670   0.09098   0.08641  -0.0447   1.0000   0.0263
  -8.750  -0.3810   0.08863   0.08416  -0.0430   1.0000   0.0261
  -8.500  -0.3813   0.08427   0.07985  -0.0461   0.9961   0.0257
  -8.250  -0.3757   0.07814   0.07373  -0.0528   0.9883   0.0252
  -8.000  -0.3751   0.07041   0.06601  -0.0624   0.9778   0.0248
  -7.750  -0.3722   0.06135   0.05679  -0.0744   0.9662   0.0242
  -7.500  -0.3645   0.05343   0.04853  -0.0828   0.9580   0.0239
  -7.250  -0.3581   0.04763   0.04235  -0.0863   0.9479   0.0237
  -7.000  -0.3408   0.04197   0.03615  -0.0901   0.9418   0.0238
  -6.750  -0.3243   0.03777   0.03144  -0.0915   0.9341   0.0242
  -6.500  -0.2965   0.03388   0.02692  -0.0939   0.9302   0.0256
  -6.250  -0.2751   0.03111   0.02349  -0.0939   0.9232   0.0274
  -6.000  -0.2469   0.02883   0.02089  -0.0950   0.9187   0.0290
  -5.750  -0.2146   0.02707   0.01889  -0.0965   0.9157   0.0308
  -5.500  -0.1912   0.02565   0.01723  -0.0960   0.9091   0.0328
  -5.250  -0.1613   0.02406   0.01533  -0.0963   0.9047   0.0364
  -5.000  -0.1291   0.02289   0.01411  -0.0975   0.9015   0.0416
  -4.750  -0.1050   0.02222   0.01330  -0.0970   0.8950   0.0500
  -4.500  -0.0751   0.02174   0.01267  -0.0976   0.8902   0.0611
  -4.250  -0.0433   0.02090   0.01186  -0.0987   0.8869   0.0721
  -4.000  -0.0209   0.02028   0.01121  -0.0978   0.8799   0.0786
  -3.750   0.0084   0.01976   0.01058  -0.0981   0.8752   0.0888
  -3.500   0.0403   0.01907   0.00989  -0.0990   0.8719   0.0995
  -3.250   0.0621   0.01869   0.00950  -0.0981   0.8644   0.1126
  -3.000   0.0917   0.01808   0.00899  -0.0987   0.8600   0.1454
  -2.750   0.1199   0.01702   0.00853  -0.0995   0.8559   0.2806
  -2.500   0.1382   0.01633   0.00876  -0.0979   0.8485   0.5086
  -2.250   0.1666   0.01626   0.00879  -0.0975   0.8443   0.5920
  -2.000   0.1907   0.01636   0.00885  -0.0964   0.8381   0.6361
  -1.750   0.2150   0.01646   0.00896  -0.0951   0.8322   0.6771
  -1.500   0.2412   0.01652   0.00905  -0.0937   0.8282   0.7243
  -1.250   0.2585   0.01669   0.00924  -0.0911   0.8201   0.7540
  -1.000   0.2870   0.01660   0.00905  -0.0908   0.8153   0.7699
  -0.750   0.3120   0.01658   0.00897  -0.0901   0.8086   0.7796
  -0.500   0.3397   0.01650   0.00881  -0.0899   0.8025   0.7893
  -0.250   0.3699   0.01639   0.00860  -0.0901   0.7973   0.7996
   0.000   0.3939   0.01640   0.00857  -0.0893   0.7895   0.8098
   0.250   0.4256   0.01621   0.00832  -0.0897   0.7847   0.8200
   0.500   0.4485   0.01621   0.00830  -0.0887   0.7753   0.8314
   0.750   0.4785   0.01600   0.00803  -0.0887   0.7682   0.8433
   1.000   0.5044   0.01585   0.00787  -0.0880   0.7582   0.8564
   1.250   0.5299   0.01571   0.00773  -0.0872   0.7481   0.8707
   1.500   0.5597   0.01547   0.00748  -0.0871   0.7390   0.8871
   1.750   0.5897   0.01530   0.00732  -0.0872   0.7281   0.9083
   2.000   0.6253   0.01520   0.00726  -0.0887   0.7169   0.9435
   2.250   0.6594   0.01516   0.00719  -0.0901   0.7067   1.0000
   2.500   0.6898   0.01516   0.00712  -0.0906   0.6976   1.0000
   2.750   0.7159   0.01527   0.00723  -0.0905   0.6858   1.0000
   3.000   0.7428   0.01537   0.00730  -0.0904   0.6740   1.0000
   3.250   0.7699   0.01545   0.00735  -0.0903   0.6617   1.0000
   3.500   0.7968   0.01554   0.00742  -0.0901   0.6484   1.0000
   3.750   0.8229   0.01565   0.00753  -0.0897   0.6339   1.0000
   4.000   0.8484   0.01578   0.00764  -0.0893   0.6179   1.0000
   4.250   0.8736   0.01591   0.00775  -0.0887   0.6004   1.0000
   4.500   0.8988   0.01605   0.00786  -0.0881   0.5813   1.0000
   4.750   0.9222   0.01625   0.00805  -0.0873   0.5587   1.0000
   5.000   0.9458   0.01647   0.00820  -0.0864   0.5342   1.0000
   5.250   0.9684   0.01676   0.00839  -0.0854   0.5070   1.0000
   5.500   0.9899   0.01712   0.00866  -0.0843   0.4777   1.0000
   5.750   1.0104   0.01757   0.00898  -0.0830   0.4488   1.0000
   6.000   1.0299   0.01810   0.00937  -0.0817   0.4213   1.0000
   6.250   1.0491   0.01867   0.00983  -0.0803   0.3957   1.0000
   6.500   1.0680   0.01927   0.01036  -0.0790   0.3733   1.0000
   6.750   1.0869   0.01987   0.01091  -0.0777   0.3538   1.0000
   7.000   1.1057   0.02050   0.01148  -0.0764   0.3367   1.0000
   7.250   1.1245   0.02113   0.01208  -0.0752   0.3219   1.0000
   7.500   1.1433   0.02177   0.01273  -0.0739   0.3085   1.0000
   7.750   1.1620   0.02242   0.01339  -0.0727   0.2965   1.0000
   8.000   1.1802   0.02310   0.01408  -0.0714   0.2851   1.0000
   8.250   1.1974   0.02381   0.01478  -0.0700   0.2745   1.0000
   8.500   1.2139   0.02450   0.01552  -0.0685   0.2637   1.0000
   8.750   1.2303   0.02520   0.01630  -0.0670   0.2532   1.0000
   9.000   1.2457   0.02595   0.01714  -0.0654   0.2432   1.0000
   9.250   1.2604   0.02673   0.01798  -0.0638   0.2335   1.0000
   9.500   1.2751   0.02748   0.01888  -0.0621   0.2234   1.0000
   9.750   1.2885   0.02831   0.01980  -0.0604   0.2139   1.0000
  10.000   1.3002   0.02919   0.02076  -0.0586   0.2042   1.0000
  10.250   1.3117   0.03008   0.02181  -0.0568   0.1931   1.0000
  10.500   1.3216   0.03107   0.02295  -0.0549   0.1816   1.0000
  10.750   1.3294   0.03217   0.02414  -0.0530   0.1691   1.0000
  11.000   1.3358   0.03340   0.02544  -0.0511   0.1557   1.0000
  11.250   1.3415   0.03478   0.02688  -0.0494   0.1424   1.0000
  11.500   1.3457   0.03636   0.02848  -0.0477   0.1282   1.0000
  11.750   1.3492   0.03811   0.03025  -0.0461   0.1149   1.0000
  12.000   1.3510   0.04010   0.03224  -0.0446   0.1008   1.0000
  12.250   1.3518   0.04229   0.03444  -0.0433   0.0873   1.0000
  12.500   1.3517   0.04467   0.03684  -0.0422   0.0750   1.0000
  12.750   1.3502   0.04730   0.03950  -0.0412   0.0635   1.0000
  13.250   1.3420   0.05345   0.04579  -0.0397   0.0443   1.0000
  13.500   1.3344   0.05712   0.04954  -0.0394   0.0378   1.0000
  13.750   1.3254   0.06115   0.05368  -0.0394   0.0321   1.0000
  14.000   1.3138   0.06575   0.05839  -0.0399   0.0288   1.0000
  14.250   1.3026   0.07058   0.06340  -0.0407   0.0260   1.0000
  14.500   1.2886   0.07612   0.06908  -0.0422   0.0241   1.0000
  14.750   1.2731   0.08222   0.07534  -0.0442   0.0230   1.0000
  15.000   1.2571   0.08877   0.08204  -0.0468   0.0223   1.0000
<< Back to HQ 3.5/12 AIRFOIL (hq3512-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.5/12 AIRFOIL (hq3512-il)