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HT12 (ht12-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HT12 (ht12-il)
Reynolds number: 200,000
Max Cl/Cd: 35.1 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht12-il-200000-n5.txt
Download as CSV file: xf-ht12-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7742   0.09642   0.09301   0.0386   1.0000   0.0147
  -8.750  -0.7777   0.09079   0.08742   0.0352   1.0000   0.0146
  -8.500  -0.7823   0.08477   0.08144   0.0307   1.0000   0.0145
  -8.250  -0.7848   0.07726   0.07396   0.0229   1.0000   0.0143
  -8.000  -0.7833   0.06714   0.06377   0.0125   1.0000   0.0141
  -7.750  -0.7829   0.05305   0.04936   0.0018   1.0000   0.0138
  -7.500  -0.7802   0.04103   0.03662  -0.0035   1.0000   0.0138
  -7.250  -0.7672   0.03368   0.02848  -0.0052   1.0000   0.0141
  -7.000  -0.7479   0.02890   0.02298  -0.0056   1.0000   0.0147
  -6.750  -0.7250   0.02589   0.01938  -0.0055   1.0000   0.0156
  -6.500  -0.7014   0.02345   0.01652  -0.0055   1.0000   0.0169
  -6.250  -0.6765   0.02192   0.01479  -0.0054   1.0000   0.0179
  -6.000  -0.6510   0.02034   0.01297  -0.0051   1.0000   0.0189
  -5.750  -0.6250   0.01891   0.01127  -0.0048   1.0000   0.0201
  -5.500  -0.5988   0.01764   0.00980  -0.0045   1.0000   0.0216
  -5.250  -0.5724   0.01659   0.00859  -0.0042   1.0000   0.0238
  -5.000  -0.5461   0.01562   0.00757  -0.0041   1.0000   0.0265
  -4.750  -0.5193   0.01480   0.00665  -0.0038   1.0000   0.0291
  -4.500  -0.4922   0.01415   0.00585  -0.0036   1.0000   0.0322
  -4.250  -0.4654   0.01339   0.00510  -0.0035   1.0000   0.0384
  -4.000  -0.4380   0.01283   0.00445  -0.0033   1.0000   0.0444
  -3.750  -0.4107   0.01230   0.00392  -0.0031   1.0000   0.0554
  -3.500  -0.3833   0.01176   0.00343  -0.0031   1.0000   0.0721
  -3.250  -0.3560   0.01128   0.00305  -0.0030   1.0000   0.0981
  -3.000  -0.3287   0.01082   0.00273  -0.0029   1.0000   0.1357
  -2.750  -0.3015   0.01035   0.00248  -0.0029   1.0000   0.1864
  -2.500  -0.2743   0.00993   0.00229  -0.0029   1.0000   0.2421
  -2.250  -0.2470   0.00957   0.00210  -0.0028   1.0000   0.2962
  -2.000  -0.2198   0.00924   0.00196  -0.0027   1.0000   0.3493
  -1.750  -0.1927   0.00890   0.00184  -0.0025   1.0000   0.4053
  -1.500  -0.1659   0.00853   0.00173  -0.0023   1.0000   0.4695
  -1.250  -0.1399   0.00806   0.00162  -0.0019   1.0000   0.5569
  -1.000  -0.1172   0.00741   0.00159  -0.0004   1.0000   0.7037
  -0.750  -0.0961   0.00687   0.00160   0.0022   1.0000   0.8869
  -0.500  -0.0542   0.00673   0.00152  -0.0005   1.0000   1.0000
  -0.250  -0.0271   0.00672   0.00148  -0.0002   1.0000   1.0000
   0.000   0.0001   0.00672   0.00147   0.0000   1.0000   1.0000
   0.250   0.0272   0.00672   0.00148   0.0002   1.0000   1.0000
   0.500   0.0544   0.00673   0.00152   0.0005   1.0000   1.0000
   0.750   0.0964   0.00687   0.00160  -0.0022   0.8850   1.0000
   1.000   0.1173   0.00741   0.00159   0.0004   0.7027   1.0000
   1.250   0.1401   0.00806   0.00162   0.0019   0.5564   1.0000
   1.500   0.1660   0.00854   0.00173   0.0023   0.4690   1.0000
   1.750   0.1929   0.00890   0.00184   0.0025   0.4048   1.0000
   2.000   0.2200   0.00924   0.00196   0.0027   0.3490   1.0000
   2.250   0.2472   0.00958   0.00210   0.0028   0.2961   1.0000
   2.500   0.2745   0.00994   0.00229   0.0029   0.2416   1.0000
   2.750   0.3017   0.01035   0.00249   0.0029   0.1861   1.0000
   3.000   0.3289   0.01082   0.00274   0.0029   0.1355   1.0000
   3.250   0.3562   0.01129   0.00305   0.0030   0.0981   1.0000
   3.500   0.3835   0.01177   0.00343   0.0031   0.0719   1.0000
   3.750   0.4108   0.01231   0.00392   0.0031   0.0554   1.0000
   4.000   0.4382   0.01283   0.00445   0.0033   0.0444   1.0000
   4.250   0.4655   0.01339   0.00510   0.0035   0.0383   1.0000
   4.500   0.4924   0.01416   0.00585   0.0036   0.0322   1.0000
   4.750   0.5195   0.01480   0.00666   0.0038   0.0291   1.0000
   5.000   0.5463   0.01563   0.00758   0.0040   0.0264   1.0000
   5.250   0.5726   0.01659   0.00859   0.0042   0.0238   1.0000
   5.500   0.5990   0.01765   0.00981   0.0045   0.0216   1.0000
   5.750   0.6252   0.01892   0.01128   0.0048   0.0201   1.0000
   6.000   0.6512   0.02036   0.01298   0.0051   0.0189   1.0000
   6.250   0.6768   0.02194   0.01482   0.0053   0.0179   1.0000
   6.500   0.7017   0.02346   0.01654   0.0055   0.0168   1.0000
   6.750   0.7252   0.02593   0.01941   0.0055   0.0156   1.0000
   7.000   0.7482   0.02891   0.02298   0.0056   0.0146   1.0000
   7.250   0.7674   0.03369   0.02849   0.0052   0.0141   1.0000
   7.500   0.7805   0.04106   0.03666   0.0035   0.0137   1.0000
   7.750   0.7832   0.05311   0.04942  -0.0019   0.0138   1.0000
   8.000   0.7837   0.06722   0.06385  -0.0126   0.0141   1.0000
   8.250   0.7852   0.07733   0.07403  -0.0231   0.0143   1.0000
   8.500   0.7829   0.08482   0.08150  -0.0308   0.0145   1.0000
   8.750   0.7783   0.09085   0.08748  -0.0353   0.0146   1.0000
   9.000   0.7749   0.09647   0.09307  -0.0388   0.0147   1.0000
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