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HT23 (ht23-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HT23 (ht23-il)
Reynolds number: 200,000
Max Cl/Cd: 50.9 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ht23-il-200000.txt
Download as CSV file: xf-ht23-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5378   0.08668   0.08350   0.0178   1.0000   0.0448
  -8.000  -0.5375   0.08265   0.07950   0.0169   1.0000   0.0459
  -7.750  -0.6527   0.08378   0.08046   0.0070   1.0000   0.0423
  -7.500  -0.6464   0.08106   0.07779   0.0106   1.0000   0.0433
  -7.250  -0.6370   0.07768   0.07440   0.0101   1.0000   0.0444
  -7.000  -0.6263   0.07347   0.07016   0.0072   1.0000   0.0458
  -6.750  -0.6136   0.06853   0.06516   0.0029   1.0000   0.0477
  -6.500  -0.5970   0.06248   0.05894  -0.0032   1.0000   0.0509
  -6.250  -0.5810   0.05412   0.04999  -0.0107   1.0000   0.0543
  -6.000  -0.5650   0.05064   0.04663  -0.0106   1.0000   0.0558
  -5.750  -0.5464   0.04774   0.04368  -0.0109   1.0000   0.0582
  -4.750  -0.4471   0.02747   0.02128  -0.0134   1.0000   0.0404
  -4.500  -0.4210   0.02427   0.01766  -0.0131   1.0000   0.0398
  -4.250  -0.3940   0.02112   0.01410  -0.0127   1.0000   0.0378
  -4.000  -0.3661   0.01865   0.01120  -0.0120   1.0000   0.0364
  -3.750  -0.3382   0.01691   0.00922  -0.0114   1.0000   0.0367
  -3.500  -0.3107   0.01555   0.00773  -0.0109   1.0000   0.0383
  -3.250  -0.2828   0.01484   0.00690  -0.0104   1.0000   0.0417
  -3.000  -0.2567   0.01327   0.00540  -0.0100   1.0000   0.0472
  -2.750  -0.2294   0.01238   0.00449  -0.0095   1.0000   0.0544
  -2.500  -0.2023   0.01144   0.00367  -0.0093   1.0000   0.0751
  -2.250  -0.1765   0.00984   0.00295  -0.0093   1.0000   0.2439
  -2.000  -0.1530   0.00855   0.00274  -0.0089   1.0000   0.5015
  -1.750  -0.1361   0.00744   0.00270  -0.0059   1.0000   0.7646
  -1.500  -0.0875   0.00703   0.00258  -0.0086   1.0000   1.0000
  -1.250  -0.0613   0.00699   0.00242  -0.0082   1.0000   1.0000
  -1.000  -0.0349   0.00697   0.00228  -0.0079   1.0000   1.0000
  -0.750  -0.0086   0.00695   0.00219  -0.0075   1.0000   1.0000
  -0.500   0.0178   0.00695   0.00213  -0.0072   1.0000   1.0000
  -0.250   0.0442   0.00696   0.00209  -0.0069   1.0000   1.0000
   0.000   0.0706   0.00698   0.00209  -0.0066   1.0000   1.0000
   0.250   0.1110   0.00703   0.00213  -0.0093   0.9633   1.0000
   0.500   0.1520   0.00719   0.00208  -0.0111   0.8458   1.0000
   0.750   0.1725   0.00757   0.00202  -0.0086   0.7420   1.0000
   1.000   0.1960   0.00798   0.00203  -0.0071   0.6669   1.0000
   1.250   0.2214   0.00833   0.00207  -0.0064   0.6123   1.0000
   1.500   0.2476   0.00865   0.00215  -0.0059   0.5694   1.0000
   1.750   0.2743   0.00894   0.00224  -0.0055   0.5335   1.0000
   2.000   0.3013   0.00922   0.00234  -0.0052   0.5026   1.0000
   2.250   0.3284   0.00948   0.00246  -0.0049   0.4744   1.0000
   2.500   0.3556   0.00973   0.00262  -0.0047   0.4485   1.0000
   2.750   0.3829   0.01000   0.00277  -0.0045   0.4247   1.0000
   3.000   0.4101   0.01025   0.00293  -0.0043   0.4013   1.0000
   3.250   0.4375   0.01051   0.00311  -0.0041   0.3788   1.0000
   3.500   0.4647   0.01077   0.00332  -0.0039   0.3568   1.0000
   3.750   0.4921   0.01103   0.00353  -0.0037   0.3341   1.0000
   4.000   0.5192   0.01131   0.00374  -0.0035   0.3119   1.0000
   4.250   0.5465   0.01156   0.00397  -0.0034   0.2880   1.0000
   4.500   0.5736   0.01185   0.00423  -0.0033   0.2638   1.0000
   4.750   0.6006   0.01217   0.00447  -0.0031   0.2387   1.0000
   5.000   0.6276   0.01249   0.00477  -0.0030   0.2130   1.0000
   5.250   0.6545   0.01289   0.00514  -0.0029   0.1865   1.0000
   5.500   0.6811   0.01338   0.00555  -0.0028   0.1577   1.0000
   5.750   0.7072   0.01407   0.00611  -0.0027   0.1278   1.0000
   6.000   0.7328   0.01496   0.00688  -0.0025   0.1014   1.0000
   6.250   0.7587   0.01581   0.00775  -0.0022   0.0840   1.0000
   6.500   0.7836   0.01689   0.00881  -0.0019   0.0737   1.0000
   6.750   0.8091   0.01775   0.00973  -0.0016   0.0662   1.0000
   7.000   0.8336   0.01900   0.01103  -0.0011   0.0604   1.0000
   7.250   0.8587   0.02006   0.01220  -0.0007   0.0561   1.0000
   7.500   0.8819   0.02185   0.01399  -0.0003   0.0523   1.0000
   7.750   0.9064   0.02321   0.01561   0.0002   0.0494   1.0000
   8.000   0.9306   0.02460   0.01721   0.0007   0.0466   1.0000
   8.250   0.9539   0.02623   0.01905   0.0012   0.0444   1.0000
   8.500   0.9758   0.02829   0.02124   0.0016   0.0425   1.0000
   8.750   0.9927   0.03219   0.02553   0.0020   0.0407   1.0000
   9.000   1.0124   0.03416   0.02797   0.0026   0.0393   1.0000
   9.250   1.0271   0.03764   0.03195   0.0031   0.0384   1.0000
   9.500   1.0363   0.04200   0.03686   0.0034   0.0379   1.0000
   9.750   1.0382   0.04732   0.04271   0.0032   0.0378   1.0000
  10.000   1.0311   0.05335   0.04922   0.0022   0.0379   1.0000
  10.250   1.0143   0.05995   0.05619  -0.0003   0.0384   1.0000
  10.500   0.9891   0.06696   0.06342  -0.0057   0.0390   1.0000
  10.750   0.9608   0.07774   0.07432  -0.0169   0.0399   1.0000
  11.000   0.9372   0.08790   0.08447  -0.0238   0.0408   1.0000
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