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

HT12 (ht12-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HT12 (ht12-il)
Reynolds number: 50,000
Max Cl/Cd: 20.29 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht12-il-50000-n5.txt
Download as CSV file: xf-ht12-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.7379   0.10393   0.09710   0.0348   1.0000   0.0496
  -8.500  -0.7382   0.09898   0.09220   0.0317   1.0000   0.0494
  -8.250  -0.7388   0.09383   0.08710   0.0279   1.0000   0.0493
  -8.000  -0.7377   0.08808   0.08139   0.0227   1.0000   0.0492
  -7.750  -0.7342   0.08169   0.07499   0.0169   1.0000   0.0490
  -7.500  -0.7288   0.07496   0.06818   0.0112   1.0000   0.0489
  -7.250  -0.7218   0.06810   0.06116   0.0060   1.0000   0.0488
  -7.000  -0.7126   0.06136   0.05415   0.0015   1.0000   0.0487
  -6.750  -0.7003   0.05488   0.04726  -0.0020   1.0000   0.0485
  -6.500  -0.6843   0.04899   0.04087  -0.0045   1.0000   0.0487
  -6.250  -0.6651   0.04380   0.03503  -0.0060   1.0000   0.0494
  -6.000  -0.6431   0.03936   0.02994  -0.0069   1.0000   0.0511
  -5.750  -0.6185   0.03557   0.02526  -0.0074   1.0000   0.0551
  -5.500  -0.5946   0.03273   0.02217  -0.0074   1.0000   0.0586
  -5.250  -0.5688   0.03015   0.01919  -0.0071   1.0000   0.0620
  -5.000  -0.5416   0.02791   0.01638  -0.0067   1.0000   0.0682
  -4.750  -0.5163   0.02616   0.01456  -0.0064   1.0000   0.0765
  -4.500  -0.4898   0.02430   0.01246  -0.0058   1.0000   0.0841
  -4.250  -0.4637   0.02294   0.01104  -0.0054   1.0000   0.0994
  -4.000  -0.4369   0.02154   0.00956  -0.0050   1.0000   0.1176
  -3.750  -0.4101   0.02031   0.00839  -0.0048   1.0000   0.1512
  -3.500  -0.3836   0.01910   0.00737  -0.0046   1.0000   0.2047
  -3.250  -0.3579   0.01806   0.00669  -0.0044   1.0000   0.2833
  -3.000  -0.3327   0.01718   0.00613  -0.0039   1.0000   0.3735
  -2.750  -0.3087   0.01630   0.00564  -0.0029   1.0000   0.4677
  -2.500  -0.2874   0.01533   0.00522  -0.0010   1.0000   0.5817
  -2.250  -0.2673   0.01433   0.00492   0.0025   1.0000   0.7674
  -2.000  -0.2112   0.01380   0.00432  -0.0020   1.0000   1.0000
  -1.750  -0.1849   0.01370   0.00393  -0.0018   1.0000   1.0000
  -1.500  -0.1586   0.01361   0.00358  -0.0016   1.0000   1.0000
  -1.250  -0.1322   0.01355   0.00332  -0.0014   1.0000   1.0000
  -1.000  -0.1058   0.01350   0.00312  -0.0011   1.0000   1.0000
  -0.750  -0.0794   0.01346   0.00297  -0.0008   1.0000   1.0000
  -0.500  -0.0529   0.01343   0.00286  -0.0006   1.0000   1.0000
  -0.250  -0.0264   0.01342   0.00278  -0.0003   1.0000   1.0000
   0.000   0.0001   0.01341   0.00276   0.0000   1.0000   1.0000
   0.250   0.0266   0.01342   0.00278   0.0003   1.0000   1.0000
   0.500   0.0531   0.01343   0.00285   0.0006   1.0000   1.0000
   0.750   0.0795   0.01346   0.00297   0.0008   1.0000   1.0000
   1.000   0.1060   0.01350   0.00312   0.0011   1.0000   1.0000
   1.250   0.1324   0.01355   0.00332   0.0014   1.0000   1.0000
   1.500   0.1587   0.01361   0.00358   0.0016   1.0000   1.0000
   1.750   0.1851   0.01370   0.00393   0.0018   1.0000   1.0000
   2.000   0.2114   0.01380   0.00433   0.0020   1.0000   1.0000
   2.250   0.2675   0.01433   0.00492  -0.0025   0.7656   1.0000
   2.500   0.2875   0.01533   0.00523   0.0010   0.5809   1.0000
   2.750   0.3089   0.01631   0.00564   0.0029   0.4671   1.0000
   3.000   0.3329   0.01718   0.00614   0.0039   0.3729   1.0000
   3.250   0.3580   0.01806   0.00669   0.0044   0.2827   1.0000
   3.500   0.3838   0.01911   0.00738   0.0046   0.2043   1.0000
   3.750   0.4103   0.02031   0.00839   0.0047   0.1510   1.0000
   4.000   0.4371   0.02154   0.00957   0.0049   0.1175   1.0000
   4.250   0.4639   0.02295   0.01105   0.0054   0.0994   1.0000
   4.500   0.4900   0.02431   0.01247   0.0058   0.0841   1.0000
   4.750   0.5165   0.02618   0.01457   0.0064   0.0764   1.0000
   5.000   0.5418   0.02793   0.01639   0.0067   0.0682   1.0000
   5.250   0.5690   0.03016   0.01921   0.0071   0.0619   1.0000
   5.500   0.5948   0.03275   0.02219   0.0074   0.0586   1.0000
   5.750   0.6187   0.03559   0.02528   0.0073   0.0550   1.0000
   6.000   0.6433   0.03939   0.02997   0.0069   0.0510   1.0000
   6.250   0.6653   0.04384   0.03507   0.0060   0.0493   1.0000
   6.500   0.6845   0.04904   0.04091   0.0044   0.0486   1.0000
   6.750   0.7005   0.05493   0.04732   0.0019   0.0485   1.0000
   7.000   0.7128   0.06143   0.05422  -0.0016   0.0487   1.0000
   7.250   0.7220   0.06820   0.06126  -0.0061   0.0489   1.0000
   7.500   0.7290   0.07506   0.06829  -0.0113   0.0490   1.0000
   7.750   0.7344   0.08180   0.07510  -0.0170   0.0491   1.0000
   8.000   0.7380   0.08818   0.08149  -0.0229   0.0492   1.0000
   8.250   0.7392   0.09393   0.08720  -0.0280   0.0494   1.0000
   8.500   0.7387   0.09908   0.09230  -0.0318   0.0495   1.0000
   8.750   0.7384   0.10403   0.09719  -0.0349   0.0497   1.0000
   9.000   0.7386   0.10885   0.10197  -0.0376   0.0500   1.0000
   9.250   0.7393   0.11361   0.10668  -0.0400   0.0508   1.0000
   9.500   0.7412   0.11825   0.11129  -0.0419   0.0521   1.0000
   9.750   0.7449   0.12273   0.11575  -0.0430   0.0537   1.0000
  10.500   0.6054   0.12654   0.11979  -0.0329   0.0616   1.0000
  10.750   0.6066   0.13082   0.12406  -0.0338   0.0652   1.0000
<< Back to HT12 (ht12-il)

Polar data table (+)

Polar graphs


<< Back to HT12 (ht12-il)