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HT22 (ht22-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HT22 (ht22-il)
Reynolds number: 200,000
Max Cl/Cd: 45.97 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ht22-il-200000.txt
Download as CSV file: xf-ht22-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT22                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6537   0.10333   0.09990   0.0363   1.0001   0.0334
  -8.000  -0.6507   0.09937   0.09597   0.0334   1.0001   0.0345
  -7.750  -0.6481   0.09524   0.09188   0.0282   1.0001   0.0356
  -7.500  -0.6372   0.08997   0.08661   0.0169   1.0001   0.0364
  -7.250  -0.6215   0.08435   0.08090   0.0077   1.0001   0.0368
  -7.000  -0.6134   0.07700   0.07348   0.0021   1.0001   0.0373
  -6.750  -0.6116   0.07422   0.07083   0.0077   1.0001   0.0390
  -6.500  -0.5988   0.07082   0.06742   0.0066   1.0001   0.0408
  -6.250  -0.5821   0.06624   0.06278   0.0028   1.0001   0.0427
  -6.000  -0.5610   0.06106   0.05745  -0.0023   1.0001   0.0455
  -5.750  -0.5254   0.05683   0.05252  -0.0097   1.0001   0.0486
  -5.500  -0.5150   0.04924   0.04506  -0.0111   1.0001   0.0503
  -5.250  -0.4964   0.04623   0.04205  -0.0114   1.0001   0.0522
  -5.000  -0.4737   0.04300   0.03866  -0.0125   1.0001   0.0556
  -4.750  -0.4445   0.03850   0.03349  -0.0147   1.0001   0.0630
  -4.500  -0.4232   0.03538   0.03045  -0.0149   1.0001   0.0656
  -4.250  -0.3948   0.03299   0.02746  -0.0156   1.0001   0.0765
  -4.000  -0.3724   0.03009   0.02474  -0.0158   1.0001   0.0816
  -3.750  -0.3362   0.02330   0.01675  -0.0150   1.0001   0.0512
  -3.500  -0.3065   0.01965   0.01260  -0.0142   1.0001   0.0417
  -3.250  -0.2777   0.01727   0.00976  -0.0136   1.0001   0.0407
  -3.000  -0.2499   0.01601   0.00836  -0.0132   1.0001   0.0436
  -2.750  -0.2219   0.01495   0.00712  -0.0126   1.0001   0.0471
  -2.500  -0.1944   0.01353   0.00560  -0.0121   1.0001   0.0498
  -2.250  -0.1675   0.01257   0.00472  -0.0118   1.0001   0.0574
  -2.000  -0.1404   0.01161   0.00377  -0.0114   1.0001   0.0658
  -1.750  -0.1132   0.01090   0.00315  -0.0111   1.0001   0.0822
  -1.500  -0.0859   0.01004   0.00253  -0.0109   1.0001   0.1292
  -1.250  -0.0714   0.00722   0.00230  -0.0086   1.0001   0.7557
  -1.000  -0.0316   0.00676   0.00209  -0.0099   1.0001   0.9999
  -0.750  -0.0048   0.00675   0.00195  -0.0095   1.0001   0.9999
  -0.500   0.0221   0.00675   0.00185  -0.0092   1.0001   0.9999
  -0.250   0.0489   0.00675   0.00178  -0.0089   1.0001   0.9999
   0.000   0.0757   0.00676   0.00175  -0.0086   1.0001   0.9999
   0.250   0.1024   0.00678   0.00175  -0.0083   1.0001   0.9999
   0.500   0.1290   0.00680   0.00177  -0.0080   1.0001   0.9999
   0.750   0.1556   0.00684   0.00183  -0.0077   1.0001   0.9999
   1.000   0.1939   0.00690   0.00192  -0.0099   0.9725   0.9999
   1.250   0.2383   0.00707   0.00193  -0.0124   0.8471   0.9999
   1.500   0.2582   0.00753   0.00192  -0.0096   0.7247   0.9999
   1.750   0.2812   0.00803   0.00197  -0.0080   0.6299   0.9999
   2.000   0.3063   0.00848   0.00207  -0.0073   0.5607   0.9999
   2.250   0.3324   0.00887   0.00219  -0.0067   0.5075   0.9999
   2.500   0.3589   0.00922   0.00233  -0.0064   0.4627   0.9999
   2.750   0.3856   0.00955   0.00250  -0.0061   0.4224   0.9999
   3.000   0.4125   0.00985   0.00267  -0.0058   0.3841   0.9999
   3.250   0.4394   0.01017   0.00285  -0.0056   0.3460   0.9999
   3.500   0.4663   0.01050   0.00304  -0.0054   0.3062   0.9999
   3.750   0.4932   0.01087   0.00329  -0.0052   0.2635   0.9999
   4.000   0.5199   0.01131   0.00356  -0.0051   0.2165   0.9999
   4.250   0.5464   0.01189   0.00394  -0.0050   0.1702   0.9999
   4.500   0.5728   0.01260   0.00445  -0.0048   0.1317   0.9999
   4.750   0.5990   0.01341   0.00514  -0.0046   0.1067   0.9999
   5.000   0.6254   0.01419   0.00588  -0.0044   0.0904   0.9999
   5.250   0.6514   0.01503   0.00671  -0.0041   0.0790   0.9999
   5.500   0.6769   0.01610   0.00771  -0.0038   0.0711   0.9999
   5.750   0.7031   0.01709   0.00885  -0.0033   0.0661   0.9999
   6.000   0.7291   0.01796   0.00975  -0.0030   0.0609   0.9999
   6.250   0.7539   0.01966   0.01153  -0.0026   0.0570   0.9999
   6.500   0.7800   0.02091   0.01302  -0.0021   0.0544   0.9999
   6.750   0.8055   0.02243   0.01478  -0.0016   0.0520   0.9999
   7.000   0.8304   0.02375   0.01624  -0.0013   0.0492   0.9999
   7.250   0.8521   0.02694   0.01964  -0.0011   0.0465   0.9999
   7.500   0.8747   0.02950   0.02268  -0.0007   0.0455   0.9999
   7.750   0.8956   0.03253   0.02625  -0.0002   0.0448   0.9999
   8.000   0.9137   0.03621   0.03053   0.0000   0.0438   0.9999
   8.250   0.9279   0.04072   0.03563  -0.0001   0.0425   0.9999
   8.500   0.9333   0.04759   0.04311  -0.0010   0.0431   0.9999
   8.750   0.9320   0.05512   0.05108  -0.0030   0.0444   0.9999
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