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HT08 (ht08-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HT08 (ht08-il)
Reynolds number: 200,000
Max Cl/Cd: 34.19 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht08-il-200000-n5.txt
Download as CSV file: xf-ht08-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT08                                            
  
 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  -0.7631   0.09521   0.09179   0.0327   1.0000   0.0146
  -8.750  -0.7667   0.08966   0.08627   0.0292   1.0000   0.0144
  -8.500  -0.7715   0.08369   0.08035   0.0245   1.0000   0.0143
  -8.250  -0.7733   0.07619   0.07285   0.0166   1.0000   0.0141
  -8.000  -0.7720   0.06716   0.06372   0.0081   1.0000   0.0139
  -7.750  -0.7697   0.05687   0.05320   0.0012   1.0000   0.0137
  -7.500  -0.7669   0.04593   0.04177  -0.0033   1.0000   0.0136
  -7.250  -0.7586   0.03699   0.03209  -0.0052   1.0000   0.0138
  -7.000  -0.7426   0.03102   0.02534  -0.0055   1.0000   0.0142
  -6.750  -0.7216   0.02714   0.02079  -0.0054   1.0000   0.0149
  -6.500  -0.6985   0.02447   0.01760  -0.0052   1.0000   0.0161
  -6.250  -0.6743   0.02280   0.01574  -0.0051   1.0000   0.0173
  -6.000  -0.6491   0.02111   0.01377  -0.0048   1.0000   0.0182
  -5.750  -0.6234   0.01957   0.01196  -0.0044   1.0000   0.0193
  -5.500  -0.5973   0.01819   0.01033  -0.0041   1.0000   0.0207
  -5.250  -0.5707   0.01733   0.00927  -0.0038   1.0000   0.0225
  -5.000  -0.5451   0.01601   0.00790  -0.0036   1.0000   0.0252
  -4.750  -0.5186   0.01516   0.00697  -0.0033   1.0000   0.0276
  -4.500  -0.4918   0.01441   0.00611  -0.0030   1.0000   0.0306
  -4.250  -0.4651   0.01367   0.00531  -0.0029   1.0000   0.0356
  -4.000  -0.4379   0.01311   0.00467  -0.0027   1.0000   0.0413
  -3.750  -0.4108   0.01252   0.00409  -0.0026   1.0000   0.0501
  -3.500  -0.3836   0.01202   0.00361  -0.0025   1.0000   0.0642
  -3.250  -0.3564   0.01154   0.00318  -0.0024   1.0000   0.0851
  -3.000  -0.3293   0.01108   0.00286  -0.0023   1.0000   0.1193
  -2.750  -0.3022   0.01062   0.00258  -0.0023   1.0000   0.1643
  -2.500  -0.2752   0.01015   0.00235  -0.0023   1.0000   0.2229
  -2.250  -0.2484   0.00968   0.00214  -0.0022   1.0000   0.2937
  -2.000  -0.2218   0.00918   0.00198  -0.0021   1.0000   0.3788
  -1.750  -0.1959   0.00863   0.00185  -0.0019   1.0000   0.4829
  -1.500  -0.1715   0.00800   0.00176  -0.0011   1.0000   0.6135
  -1.250  -0.1504   0.00740   0.00172   0.0009   1.0000   0.7585
  -1.000  -0.1262   0.00705   0.00173   0.0028   1.0000   0.8983
  -0.750  -0.0805   0.00693   0.00165  -0.0007   1.0000   0.9960
  -0.500  -0.0529   0.00691   0.00159  -0.0006   1.0000   1.0000
  -0.250  -0.0264   0.00690   0.00154  -0.0003   1.0000   1.0000
   0.000   0.0001   0.00689   0.00153   0.0000   1.0000   1.0000
   0.250   0.0266   0.00690   0.00154   0.0003   1.0000   1.0000
   0.500   0.0530   0.00691   0.00159   0.0006   1.0000   1.0000
   0.750   0.0808   0.00693   0.00165   0.0007   0.9957   1.0000
   1.000   0.1264   0.00705   0.00173  -0.0028   0.8976   1.0000
   1.250   0.1505   0.00740   0.00172  -0.0009   0.7581   1.0000
   1.500   0.1716   0.00800   0.00176   0.0011   0.6126   1.0000
   1.750   0.1960   0.00863   0.00185   0.0019   0.4824   1.0000
   2.000   0.2219   0.00918   0.00198   0.0021   0.3785   1.0000
   2.250   0.2485   0.00968   0.00214   0.0022   0.2934   1.0000
   2.500   0.2753   0.01015   0.00235   0.0023   0.2225   1.0000
   2.750   0.3023   0.01062   0.00258   0.0023   0.1639   1.0000
   3.000   0.3294   0.01108   0.00286   0.0023   0.1191   1.0000
   3.250   0.3565   0.01155   0.00319   0.0024   0.0850   1.0000
   3.500   0.3838   0.01202   0.00362   0.0024   0.0641   1.0000
   3.750   0.4110   0.01252   0.00409   0.0026   0.0501   1.0000
   4.000   0.4381   0.01311   0.00468   0.0027   0.0413   1.0000
   4.250   0.4652   0.01367   0.00531   0.0029   0.0356   1.0000
   4.500   0.4919   0.01442   0.00611   0.0030   0.0305   1.0000
   4.750   0.5187   0.01517   0.00697   0.0033   0.0275   1.0000
   5.000   0.5453   0.01602   0.00790   0.0035   0.0252   1.0000
   5.250   0.5709   0.01734   0.00928   0.0038   0.0225   1.0000
   5.500   0.5975   0.01820   0.01034   0.0041   0.0206   1.0000
   5.750   0.6236   0.01957   0.01197   0.0044   0.0193   1.0000
   6.000   0.6493   0.02112   0.01378   0.0048   0.0181   1.0000
   6.250   0.6745   0.02282   0.01576   0.0050   0.0173   1.0000
   6.500   0.6987   0.02448   0.01761   0.0051   0.0162   1.0000
   6.750   0.7218   0.02717   0.02082   0.0053   0.0149   1.0000
   7.000   0.7428   0.03104   0.02536   0.0055   0.0142   1.0000
   7.250   0.7588   0.03704   0.03214   0.0051   0.0138   1.0000
   7.500   0.7671   0.04599   0.04183   0.0032   0.0136   1.0000
   7.750   0.7700   0.05694   0.05327  -0.0012   0.0137   1.0000
   8.000   0.7723   0.06723   0.06380  -0.0082   0.0139   1.0000
   8.250   0.7736   0.07628   0.07294  -0.0167   0.0141   1.0000
   8.500   0.7719   0.08377   0.08042  -0.0246   0.0143   1.0000
   8.750   0.7671   0.08974   0.08635  -0.0293   0.0144   1.0000
   9.000   0.7636   0.09531   0.09188  -0.0329   0.0146   1.0000
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