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

HT05 (ht05-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HT05 (ht05-il)
Reynolds number: 50,000
Max Cl/Cd: 19.62 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht05-il-50000-n5.txt
Download as CSV file: xf-ht05-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT05                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7287   0.10319   0.09624   0.0188   1.0000   0.0475
  -8.750  -0.7259   0.09879   0.09186   0.0177   1.0000   0.0470
  -8.500  -0.7259   0.09394   0.08705   0.0150   1.0000   0.0468
  -8.250  -0.7248   0.08847   0.08159   0.0111   1.0000   0.0467
  -8.000  -0.7225   0.08238   0.07547   0.0065   1.0000   0.0466
  -7.750  -0.7197   0.07600   0.06900   0.0021   1.0000   0.0467
  -7.500  -0.7160   0.06953   0.06235  -0.0019   1.0000   0.0469
  -7.250  -0.7101   0.06331   0.05581  -0.0050   1.0000   0.0473
  -7.000  -0.7014   0.05757   0.04974  -0.0070   1.0000   0.0478
  -6.750  -0.6892   0.05266   0.04446  -0.0079   1.0000   0.0484
  -6.500  -0.6743   0.04791   0.03929  -0.0085   1.0000   0.0489
  -6.250  -0.6568   0.04378   0.03474  -0.0087   1.0000   0.0498
  -6.000  -0.6372   0.04006   0.03057  -0.0087   1.0000   0.0512
  -5.750  -0.6158   0.03706   0.02711  -0.0085   1.0000   0.0552
  -5.500  -0.5921   0.03374   0.02296  -0.0080   1.0000   0.0596
  -5.250  -0.5684   0.03097   0.01989  -0.0075   1.0000   0.0626
  -5.000  -0.5436   0.02881   0.01743  -0.0069   1.0000   0.0670
  -4.750  -0.5183   0.02698   0.01522  -0.0063   1.0000   0.0758
  -4.500  -0.4927   0.02532   0.01333  -0.0056   1.0000   0.0844
  -4.250  -0.4671   0.02381   0.01172  -0.0049   1.0000   0.0955
  -4.000  -0.4408   0.02257   0.01031  -0.0043   1.0000   0.1141
  -3.750  -0.4140   0.02135   0.00915  -0.0040   1.0000   0.1377
  -3.500  -0.3877   0.02026   0.00807  -0.0036   1.0000   0.1739
  -3.250  -0.3617   0.01925   0.00722  -0.0032   1.0000   0.2272
  -3.000  -0.3366   0.01823   0.00654  -0.0028   1.0000   0.2970
  -2.750  -0.3134   0.01719   0.00604  -0.0018   1.0000   0.3978
  -2.500  -0.2928   0.01624   0.00567   0.0001   1.0000   0.5395
  -2.250  -0.2709   0.01526   0.00508   0.0021   1.0000   0.6656
  -2.000  -0.1971   0.01452   0.00467  -0.0056   1.0000   1.0000
  -1.750  -0.1725   0.01443   0.00431  -0.0050   1.0000   1.0000
  -1.500  -0.1478   0.01436   0.00399  -0.0043   1.0000   1.0000
  -1.250  -0.1232   0.01430   0.00375  -0.0036   1.0000   1.0000
  -1.000  -0.0986   0.01425   0.00356  -0.0029   1.0000   1.0000
  -0.750  -0.0739   0.01422   0.00342  -0.0022   1.0000   1.0000
  -0.500  -0.0493   0.01419   0.00331  -0.0014   1.0000   1.0000
  -0.250  -0.0246   0.01418   0.00325  -0.0007   1.0000   1.0000
   0.000   0.0001   0.01417   0.00323   0.0000   1.0000   1.0000
   0.250   0.0247   0.01418   0.00325   0.0007   1.0000   1.0000
   0.500   0.0494   0.01419   0.00331   0.0014   1.0000   1.0000
   0.750   0.0740   0.01422   0.00342   0.0022   1.0000   1.0000
   1.000   0.0987   0.01425   0.00356   0.0029   1.0000   1.0000
   1.250   0.1233   0.01430   0.00375   0.0036   1.0000   1.0000
   1.500   0.1480   0.01436   0.00399   0.0043   1.0000   1.0000
   1.750   0.1726   0.01443   0.00431   0.0050   1.0000   1.0000
   2.000   0.1972   0.01452   0.00467   0.0056   1.0000   1.0000
   2.250   0.2710   0.01527   0.00508  -0.0021   0.6648   1.0000
   2.500   0.2929   0.01625   0.00567  -0.0001   0.5387   1.0000
   2.750   0.3135   0.01720   0.00604   0.0018   0.3972   1.0000
   3.000   0.3367   0.01823   0.00654   0.0028   0.2967   1.0000
   3.250   0.3619   0.01925   0.00723   0.0032   0.2270   1.0000
   3.500   0.3878   0.02027   0.00808   0.0036   0.1739   1.0000
   3.750   0.4141   0.02135   0.00915   0.0040   0.1374   1.0000
   4.000   0.4409   0.02258   0.01033   0.0043   0.1140   1.0000
   4.250   0.4672   0.02381   0.01173   0.0049   0.0953   1.0000
   4.500   0.4928   0.02533   0.01335   0.0056   0.0844   1.0000
   4.750   0.5184   0.02698   0.01520   0.0063   0.0757   1.0000
   5.000   0.5437   0.02881   0.01744   0.0069   0.0670   1.0000
   5.250   0.5685   0.03099   0.01990   0.0075   0.0626   1.0000
   5.500   0.5923   0.03377   0.02302   0.0080   0.0596   1.0000
   5.750   0.6159   0.03707   0.02712   0.0084   0.0551   1.0000
   6.000   0.6374   0.04004   0.03055   0.0087   0.0512   1.0000
   6.250   0.6570   0.04375   0.03469   0.0087   0.0498   1.0000
   6.500   0.6745   0.04789   0.03926   0.0085   0.0489   1.0000
   6.750   0.6893   0.05264   0.04443   0.0079   0.0484   1.0000
   7.250   0.7100   0.06349   0.05602   0.0048   0.0473   1.0000
   7.500   0.7158   0.06977   0.06260   0.0016   0.0469   1.0000
   7.750   0.7195   0.07626   0.06927  -0.0024   0.0468   1.0000
   8.000   0.7225   0.08263   0.07573  -0.0069   0.0467   1.0000
   8.250   0.7249   0.08868   0.08181  -0.0115   0.0467   1.0000
   8.500   0.7259   0.09411   0.08722  -0.0153   0.0469   1.0000
   8.750   0.7261   0.09892   0.09199  -0.0179   0.0472   1.0000
   9.000   0.7289   0.10332   0.09637  -0.0190   0.0476   1.0000
<< Back to HT05 (ht05-il)

Polar data table (+)

Polar graphs


<< Back to HT05 (ht05-il)