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

HT23 (ht23-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HT23 (ht23-il)
Reynolds number: 50,000
Max Cl/Cd: 24.41 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ht23-il-50000.txt
Download as CSV file: xf-ht23-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT23                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6639   0.13070   0.12398   0.0428   1.0000   0.1801
  -9.250  -0.6498   0.12606   0.11932   0.0437   1.0000   0.1889
  -9.000  -0.6706   0.12557   0.11896   0.0386   1.0000   0.1931
  -8.750  -0.6429   0.11881   0.11216   0.0419   1.0000   0.2034
  -8.500  -0.6477   0.11571   0.10915   0.0396   1.0000   0.2095
  -8.250  -0.6416   0.11219   0.10564   0.0394   1.0000   0.2202
  -8.000  -0.6314   0.10772   0.10120   0.0397   1.0000   0.2286
  -7.750  -0.6451   0.10560   0.09922   0.0360   1.0000   0.2375
  -7.500  -0.6381   0.10194   0.09558   0.0361   1.0000   0.2504
  -7.250  -0.6256   0.09756   0.09123   0.0372   1.0000   0.2633
  -7.000  -0.6221   0.09408   0.08781   0.0368   1.0000   0.2781
  -6.750  -0.6265   0.09121   0.08502   0.0329   1.0000   0.2950
  -6.500  -0.6056   0.08655   0.08038   0.0380   1.0000   0.3155
  -6.000  -0.5899   0.07971   0.07365   0.0395   1.0000   0.3639
  -5.750  -0.5801   0.07613   0.07010   0.0410   1.0000   0.3895
  -5.500  -0.5701   0.07274   0.06676   0.0428   1.0000   0.4190
  -5.250  -0.5651   0.06987   0.06396   0.0448   1.0000   0.4578
  -5.000  -0.4920   0.04701   0.03917  -0.0122   1.0000   0.1473
  -4.750  -0.4636   0.04140   0.03298  -0.0142   1.0000   0.1290
  -4.500  -0.4336   0.03682   0.02755  -0.0153   1.0000   0.1188
  -4.250  -0.4057   0.03345   0.02361  -0.0155   1.0000   0.1187
  -4.000  -0.3764   0.03065   0.02012  -0.0153   1.0000   0.1211
  -3.750  -0.3483   0.02776   0.01703  -0.0150   1.0000   0.1242
  -3.500  -0.3196   0.02566   0.01464  -0.0144   1.0000   0.1334
  -3.250  -0.2912   0.02363   0.01248  -0.0137   1.0000   0.1475
  -3.000  -0.2633   0.02172   0.01063  -0.0129   1.0000   0.1724
  -2.750  -0.2353   0.01976   0.00891  -0.0120   1.0000   0.2255
  -2.500  -0.1859   0.01442   0.00693  -0.0105   1.0000   1.0000
  -2.250  -0.1608   0.01426   0.00616  -0.0102   1.0000   1.0000
  -2.000  -0.1357   0.01413   0.00558  -0.0098   1.0000   1.0000
  -1.750  -0.1104   0.01403   0.00514  -0.0094   1.0000   1.0000
  -1.500  -0.0850   0.01396   0.00478  -0.0090   1.0000   1.0000
  -1.250  -0.0595   0.01390   0.00449  -0.0087   1.0000   1.0000
  -1.000  -0.0339   0.01386   0.00425  -0.0083   1.0000   1.0000
  -0.750  -0.0083   0.01384   0.00407  -0.0080   1.0000   1.0000
  -0.500   0.0173   0.01383   0.00396  -0.0076   1.0000   1.0000
  -0.250   0.0429   0.01384   0.00388  -0.0073   1.0000   1.0000
   0.000   0.0684   0.01387   0.00386  -0.0070   1.0000   1.0000
   0.250   0.0937   0.01392   0.00390  -0.0068   1.0000   1.0000
   0.500   0.1187   0.01399   0.00401  -0.0066   1.0000   1.0000
   0.750   0.1432   0.01410   0.00419  -0.0064   1.0000   1.0000
   1.000   0.1662   0.01429   0.00451  -0.0065   1.0000   1.0000
   1.250   0.1842   0.01476   0.00514  -0.0069   1.0000   1.0000
   1.500   0.2695   0.01556   0.00617  -0.0208   0.9380   1.0000
   1.750   0.3300   0.01594   0.00652  -0.0253   0.8460   1.0000
   2.000   0.3530   0.01652   0.00687  -0.0224   0.7836   1.0000
   2.250   0.3735   0.01717   0.00733  -0.0195   0.7335   1.0000
   2.500   0.3953   0.01785   0.00783  -0.0172   0.6897   1.0000
   2.750   0.4183   0.01856   0.00840  -0.0154   0.6500   1.0000
   3.000   0.4421   0.01930   0.00906  -0.0139   0.6131   1.0000
   3.250   0.4660   0.02007   0.00975  -0.0125   0.5789   1.0000
   3.500   0.4902   0.02087   0.01050  -0.0112   0.5468   1.0000
   3.750   0.5148   0.02170   0.01126  -0.0099   0.5156   1.0000
   4.000   0.5396   0.02258   0.01215  -0.0090   0.4835   1.0000
   4.250   0.5643   0.02348   0.01306  -0.0079   0.4523   1.0000
   4.500   0.5889   0.02440   0.01403  -0.0069   0.4211   1.0000
   4.750   0.6133   0.02532   0.01495  -0.0057   0.3895   1.0000
   5.000   0.6373   0.02619   0.01578  -0.0045   0.3575   1.0000
   5.250   0.6612   0.02709   0.01658  -0.0032   0.3251   1.0000
   5.500   0.6852   0.02820   0.01784  -0.0023   0.2898   1.0000
   5.750   0.7090   0.02930   0.01889  -0.0011   0.2573   1.0000
   6.000   0.7328   0.03064   0.02022  -0.0002   0.2278   1.0000
   6.250   0.7564   0.03243   0.02217   0.0005   0.2022   1.0000
   6.500   0.7803   0.03453   0.02436   0.0012   0.1845   1.0000
   6.750   0.8033   0.03694   0.02694   0.0017   0.1697   1.0000
   7.000   0.8242   0.04021   0.03074   0.0017   0.1579   1.0000
   7.250   0.8426   0.04475   0.03581   0.0013   0.1526   1.0000
   7.500   0.8548   0.05060   0.04238  -0.0003   0.1484   1.0000
   7.750   0.8741   0.05313   0.04489   0.0004   0.1406   1.0000
   8.000   0.8785   0.06017   0.05249  -0.0021   0.1404   1.0000
   8.250   0.8815   0.06721   0.05989  -0.0048   0.1420   1.0000
   8.500   0.8288   0.08766   0.08071  -0.0262   0.1650   1.0000
   8.750   0.8417   0.09256   0.08567  -0.0254   0.1694   1.0000
<< Back to HT23 (ht23-il)

Polar data table (+)

Polar graphs


<< Back to HT23 (ht23-il)