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HT23 (ht23-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HT23 (ht23-il)
Reynolds number: 100,000
Max Cl/Cd: 39.77 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht23-il-100000-n5.txt
Download as CSV file: xf-ht23-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5278   0.11026   0.10559   0.0228   1.0000   0.0506
  -9.250  -0.5395   0.10617   0.10155   0.0185   1.0000   0.0527
  -9.000  -0.5433   0.10179   0.09721   0.0162   1.0000   0.0529
  -8.750  -0.5397   0.09328   0.08865   0.0165   1.0000   0.0307
  -8.500  -0.5385   0.08914   0.08453   0.0161   1.0000   0.0297
  -8.250  -0.5402   0.08454   0.07996   0.0149   1.0000   0.0288
  -8.000  -0.5440   0.07962   0.07508   0.0130   1.0000   0.0281
  -7.750  -0.5494   0.07444   0.06995   0.0106   1.0000   0.0275
  -7.250  -0.6347   0.06703   0.06206  -0.0013   1.0000   0.0236
  -7.000  -0.6250   0.06194   0.05686  -0.0040   1.0000   0.0233
  -6.750  -0.6130   0.05619   0.05089  -0.0070   1.0000   0.0231
  -6.500  -0.5984   0.05000   0.04432  -0.0099   1.0000   0.0234
  -6.000  -0.5637   0.04086   0.03447  -0.0126   1.0000   0.0249
  -5.750  -0.5426   0.03662   0.02974  -0.0132   1.0000   0.0250
  -5.500  -0.5195   0.03263   0.02515  -0.0135   1.0000   0.0249
  -5.250  -0.4952   0.02955   0.02161  -0.0134   1.0000   0.0251
  -5.000  -0.4698   0.02699   0.01862  -0.0132   1.0000   0.0256
  -4.750  -0.4437   0.02482   0.01609  -0.0129   1.0000   0.0264
  -4.500  -0.4171   0.02299   0.01397  -0.0126   1.0000   0.0277
  -4.250  -0.3900   0.02158   0.01226  -0.0121   1.0000   0.0305
  -4.000  -0.3634   0.01996   0.01044  -0.0117   1.0000   0.0326
  -3.750  -0.3371   0.01860   0.00906  -0.0113   1.0000   0.0348
  -3.500  -0.3106   0.01754   0.00793  -0.0108   1.0000   0.0385
  -3.250  -0.2843   0.01655   0.00691  -0.0104   1.0000   0.0446
  -3.000  -0.2575   0.01571   0.00600  -0.0100   1.0000   0.0523
  -2.750  -0.2309   0.01484   0.00517  -0.0097   1.0000   0.0682
  -2.500  -0.2045   0.01390   0.00449  -0.0095   1.0000   0.1101
  -2.250  -0.1790   0.01283   0.00406  -0.0093   1.0000   0.2371
  -2.000  -0.1548   0.01175   0.00380  -0.0089   1.0000   0.4270
  -1.750  -0.1349   0.01065   0.00368  -0.0067   1.0000   0.6673
  -1.500  -0.1002   0.00996   0.00358  -0.0061   1.0000   0.9388
  -1.250  -0.0604   0.00984   0.00329  -0.0084   1.0000   1.0000
  -1.000  -0.0344   0.00981   0.00311  -0.0081   1.0000   1.0000
  -0.750  -0.0084   0.00979   0.00298  -0.0077   1.0000   1.0000
  -0.500   0.0176   0.00979   0.00289  -0.0074   1.0000   1.0000
  -0.250   0.0436   0.00980   0.00284  -0.0071   1.0000   1.0000
   0.000   0.0839   0.00989   0.00284  -0.0096   0.9358   1.0000
   0.250   0.1234   0.01007   0.00279  -0.0113   0.8332   1.0000
   0.500   0.1486   0.01037   0.00273  -0.0099   0.7456   1.0000
   0.750   0.1725   0.01070   0.00272  -0.0084   0.6790   1.0000
   1.000   0.1977   0.01101   0.00274  -0.0075   0.6267   1.0000
   1.250   0.2236   0.01130   0.00279  -0.0068   0.5842   1.0000
   1.500   0.2498   0.01158   0.00288  -0.0062   0.5479   1.0000
   1.750   0.2762   0.01186   0.00298  -0.0058   0.5162   1.0000
   2.000   0.3030   0.01214   0.00311  -0.0054   0.4873   1.0000
   2.250   0.3298   0.01241   0.00326  -0.0051   0.4607   1.0000
   2.500   0.3567   0.01268   0.00346  -0.0047   0.4361   1.0000
   3.000   0.4104   0.01325   0.00386  -0.0042   0.3904   1.0000
   3.250   0.4373   0.01356   0.00409  -0.0039   0.3693   1.0000
   3.500   0.4643   0.01385   0.00439  -0.0036   0.3473   1.0000
   3.750   0.4911   0.01418   0.00466  -0.0034   0.3264   1.0000
   4.000   0.5180   0.01449   0.00497  -0.0032   0.3045   1.0000
   4.250   0.5447   0.01485   0.00528  -0.0029   0.2829   1.0000
   4.500   0.5715   0.01520   0.00568  -0.0027   0.2596   1.0000
   4.750   0.5981   0.01559   0.00604  -0.0025   0.2360   1.0000
   5.000   0.6246   0.01603   0.00644  -0.0024   0.2113   1.0000
   5.250   0.6510   0.01649   0.00693  -0.0022   0.1851   1.0000
   5.500   0.6771   0.01704   0.00745  -0.0021   0.1588   1.0000
   5.750   0.7031   0.01768   0.00806  -0.0020   0.1340   1.0000
   6.000   0.7286   0.01845   0.00878  -0.0018   0.1128   1.0000
   6.250   0.7540   0.01929   0.00968  -0.0016   0.0944   1.0000
   6.500   0.7789   0.02024   0.01064  -0.0014   0.0805   1.0000
   6.750   0.8033   0.02132   0.01178  -0.0011   0.0716   1.0000
   7.000   0.8271   0.02242   0.01291  -0.0008   0.0638   1.0000
   7.250   0.8513   0.02353   0.01423  -0.0004   0.0578   1.0000
   7.500   0.8746   0.02479   0.01558   0.0000   0.0537   1.0000
   7.750   0.8970   0.02630   0.01722   0.0004   0.0502   1.0000
   8.000   0.9201   0.02767   0.01884   0.0009   0.0462   1.0000
   8.250   0.9421   0.02919   0.02050   0.0013   0.0434   1.0000
   8.500   0.9625   0.03115   0.02258   0.0017   0.0415   1.0000
   8.750   0.9826   0.03340   0.02518   0.0022   0.0396   1.0000
   9.000   1.0015   0.03564   0.02785   0.0026   0.0372   1.0000
   9.250   1.0187   0.03797   0.03053   0.0029   0.0352   1.0000
   9.500   1.0336   0.04066   0.03354   0.0032   0.0340   1.0000
   9.750   1.0463   0.04353   0.03669   0.0034   0.0330   1.0000
  10.000   1.0556   0.04684   0.04027   0.0035   0.0322   1.0000
  10.250   1.0567   0.05121   0.04502   0.0032   0.0315   1.0000
  10.500   1.0490   0.05615   0.05046   0.0020   0.0310   1.0000
  10.750   1.0327   0.06152   0.05617  -0.0003   0.0309   1.0000
  11.000   1.0121   0.06840   0.06327  -0.0065   0.0311   1.0000
  11.250   0.9904   0.07745   0.07243  -0.0150   0.0314   1.0000
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