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HT12 (ht12-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: HT12 (ht12-il)
Reynolds number: 1,000,000
Max Cl/Cd: 58.34 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ht12-il-1000000.txt
Download as CSV file: xf-ht12-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.7997   0.16223   0.16062   0.0711   1.0000   0.0085
 -12.250  -0.7948   0.15850   0.15689   0.0696   1.0000   0.0088
  -8.750  -0.9061   0.03752   0.03531  -0.0021   1.0000   0.0073
  -8.500  -0.8944   0.03199   0.02929  -0.0037   1.0000   0.0074
  -8.250  -0.8853   0.02474   0.02123  -0.0045   1.0000   0.0076
  -8.000  -0.8686   0.02000   0.01589  -0.0044   1.0000   0.0079
  -7.750  -0.8455   0.01803   0.01366  -0.0042   1.0000   0.0082
  -7.500  -0.8205   0.01687   0.01236  -0.0041   1.0000   0.0085
  -7.250  -0.7948   0.01594   0.01129  -0.0039   1.0000   0.0089
  -7.000  -0.7686   0.01521   0.01046  -0.0037   1.0000   0.0094
  -6.750  -0.7421   0.01457   0.00972  -0.0035   1.0000   0.0101
  -6.500  -0.7155   0.01390   0.00893  -0.0033   1.0000   0.0106
  -6.250  -0.6885   0.01338   0.00833  -0.0031   1.0000   0.0110
  -6.000  -0.6631   0.01161   0.00632  -0.0028   1.0000   0.0119
  -5.750  -0.6362   0.01090   0.00554  -0.0027   1.0000   0.0128
  -5.500  -0.6089   0.01047   0.00507  -0.0026   1.0000   0.0138
  -5.250  -0.5814   0.01007   0.00462  -0.0024   1.0000   0.0150
  -5.000  -0.5537   0.00984   0.00437  -0.0023   1.0000   0.0159
  -4.750  -0.5263   0.00906   0.00348  -0.0022   1.0000   0.0182
  -4.500  -0.4985   0.00872   0.00313  -0.0021   1.0000   0.0205
  -4.250  -0.4706   0.00847   0.00285  -0.0020   1.0000   0.0226
  -4.000  -0.4427   0.00805   0.00240  -0.0019   1.0000   0.0273
  -3.750  -0.4148   0.00782   0.00218  -0.0018   1.0000   0.0319
  -3.500  -0.3868   0.00750   0.00189  -0.0017   1.0000   0.0421
  -3.250  -0.3588   0.00723   0.00167  -0.0017   1.0000   0.0568
  -3.000  -0.3308   0.00694   0.00149  -0.0017   1.0000   0.0812
  -2.750  -0.3029   0.00666   0.00134  -0.0017   1.0000   0.1106
  -2.500  -0.2749   0.00636   0.00120  -0.0017   1.0000   0.1497
  -2.250  -0.2469   0.00607   0.00107  -0.0017   1.0000   0.1956
  -2.000  -0.2189   0.00581   0.00098  -0.0017   1.0000   0.2412
  -1.750  -0.1909   0.00558   0.00091  -0.0017   1.0000   0.2859
  -1.500  -0.1628   0.00536   0.00086  -0.0017   1.0000   0.3312
  -1.250  -0.1348   0.00515   0.00081  -0.0017   1.0000   0.3761
  -1.000  -0.1068   0.00493   0.00077  -0.0017   1.0000   0.4297
  -0.750  -0.0789   0.00467   0.00074  -0.0017   1.0000   0.4917
  -0.500  -0.0514   0.00429   0.00072  -0.0017   1.0000   0.5938
  -0.250  -0.0251   0.00375   0.00072  -0.0013   1.0000   0.7452
   0.000   0.0001   0.00338   0.00080   0.0000   0.9090   0.9093
   0.250   0.0252   0.00376   0.00072   0.0013   0.7429   1.0000
   0.500   0.0515   0.00429   0.00072   0.0017   0.5925   1.0000
   0.750   0.0791   0.00468   0.00074   0.0017   0.4914   1.0000
   1.000   0.1070   0.00493   0.00077   0.0017   0.4294   1.0000
   1.250   0.1350   0.00515   0.00081   0.0017   0.3760   1.0000
   1.500   0.1630   0.00536   0.00086   0.0017   0.3309   1.0000
   1.750   0.1911   0.00558   0.00092   0.0017   0.2856   1.0000
   2.000   0.2191   0.00580   0.00099   0.0017   0.2422   1.0000
   2.250   0.2471   0.00607   0.00107   0.0017   0.1953   1.0000
   2.500   0.2751   0.00636   0.00120   0.0017   0.1498   1.0000
   2.750   0.3030   0.00666   0.00134   0.0017   0.1104   1.0000
   3.000   0.3310   0.00694   0.00150   0.0017   0.0809   1.0000
   3.250   0.3590   0.00723   0.00167   0.0017   0.0570   1.0000
   3.500   0.3870   0.00750   0.00190   0.0017   0.0420   1.0000
   3.750   0.4150   0.00783   0.00218   0.0018   0.0319   1.0000
   4.000   0.4429   0.00805   0.00240   0.0019   0.0273   1.0000
   4.250   0.4708   0.00847   0.00284   0.0020   0.0225   1.0000
   4.500   0.4987   0.00872   0.00313   0.0021   0.0205   1.0000
   4.750   0.5265   0.00907   0.00349   0.0022   0.0182   1.0000
   5.000   0.5539   0.00985   0.00438   0.0023   0.0159   1.0000
   5.250   0.5816   0.01008   0.00463   0.0024   0.0149   1.0000
   5.500   0.6091   0.01048   0.00507   0.0025   0.0138   1.0000
   5.750   0.6365   0.01091   0.00556   0.0027   0.0128   1.0000
   6.000   0.6634   0.01161   0.00631   0.0028   0.0118   1.0000
   6.250   0.6888   0.01334   0.00828   0.0031   0.0110   1.0000
   6.500   0.7158   0.01388   0.00891   0.0033   0.0106   1.0000
   6.750   0.7424   0.01460   0.00975   0.0035   0.0101   1.0000
   7.000   0.7689   0.01525   0.01051   0.0037   0.0094   1.0000
   7.250   0.7951   0.01599   0.01136   0.0038   0.0089   1.0000
   7.500   0.8207   0.01692   0.01241   0.0040   0.0085   1.0000
   7.750   0.8458   0.01801   0.01364   0.0042   0.0082   1.0000
   8.000   0.8691   0.01993   0.01581   0.0044   0.0079   1.0000
   8.250   0.8865   0.02439   0.02084   0.0045   0.0076   1.0000
   8.500   0.8922   0.03299   0.03036   0.0034   0.0074   1.0000
   8.750   0.9049   0.03816   0.03598   0.0018   0.0073   1.0000
  13.250   0.6212   0.15622   0.15469  -0.0490   0.0097   1.0000
  13.500   0.6207   0.15866   0.15714  -0.0497   0.0092   1.0000
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