Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HT05 (ht05-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: HT05 (ht05-il)
Reynolds number: 500,000
Max Cl/Cd: 49.29 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht05-il-500000-n5.txt
Download as CSV file: xf-ht05-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT05                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.7687   0.11042   0.10817   0.0372   1.0000   0.0070
 -10.000  -0.7739   0.10444   0.10221   0.0344   1.0000   0.0070
  -9.000  -0.9186   0.03307   0.02973  -0.0101   1.0000   0.0065
  -8.750  -0.9082   0.02815   0.02421  -0.0098   1.0000   0.0068
  -8.500  -0.8914   0.02504   0.02065  -0.0092   1.0000   0.0071
  -8.250  -0.8707   0.02308   0.01835  -0.0087   1.0000   0.0075
  -8.000  -0.8480   0.02169   0.01669  -0.0082   1.0000   0.0078
  -7.750  -0.8269   0.01958   0.01423  -0.0075   1.0000   0.0082
  -7.500  -0.8040   0.01815   0.01259  -0.0070   1.0000   0.0087
  -7.250  -0.7794   0.01725   0.01156  -0.0066   1.0000   0.0091
  -7.000  -0.7545   0.01643   0.01059  -0.0062   1.0000   0.0096
  -6.750  -0.7294   0.01561   0.00962  -0.0058   1.0000   0.0101
  -6.500  -0.7041   0.01482   0.00869  -0.0053   1.0000   0.0108
  -6.250  -0.6785   0.01410   0.00785  -0.0049   1.0000   0.0114
  -6.000  -0.6525   0.01360   0.00724  -0.0045   1.0000   0.0119
  -5.750  -0.6273   0.01273   0.00627  -0.0041   1.0000   0.0135
  -5.500  -0.6011   0.01227   0.00576  -0.0038   1.0000   0.0146
  -5.250  -0.5749   0.01179   0.00522  -0.0034   1.0000   0.0158
  -5.000  -0.5485   0.01137   0.00473  -0.0030   1.0000   0.0170
  -4.750  -0.5222   0.01091   0.00420  -0.0027   1.0000   0.0186
  -4.500  -0.4958   0.01050   0.00378  -0.0023   1.0000   0.0218
  -4.250  -0.4693   0.01020   0.00344  -0.0020   1.0000   0.0246
  -4.000  -0.4428   0.00986   0.00308  -0.0016   1.0000   0.0282
  -3.750  -0.4163   0.00957   0.00280  -0.0012   1.0000   0.0337
  -3.500  -0.3899   0.00929   0.00253  -0.0009   1.0000   0.0419
  -3.250  -0.3634   0.00904   0.00231  -0.0005   1.0000   0.0522
  -3.000  -0.3371   0.00878   0.00211  -0.0001   1.0000   0.0664
  -2.750  -0.3109   0.00853   0.00194   0.0003   1.0000   0.0862
  -2.500  -0.2848   0.00829   0.00180   0.0007   1.0000   0.1078
  -2.250  -0.2587   0.00805   0.00166   0.0011   1.0000   0.1358
  -2.000  -0.2327   0.00783   0.00156   0.0015   1.0000   0.1645
  -1.750  -0.1940   0.00756   0.00144  -0.0010   0.9850   0.2049
  -1.500  -0.1576   0.00730   0.00133  -0.0028   0.9576   0.2505
  -1.250  -0.1291   0.00730   0.00120  -0.0024   0.8195   0.2977
  -1.000  -0.1078   0.00771   0.00116  -0.0009   0.6561   0.3566
  -0.750  -0.0817   0.00753   0.00118  -0.0007   0.6421   0.4434
  -0.500  -0.0554   0.00738   0.00126  -0.0003   0.6326   0.5313
  -0.250  -0.0279   0.00733   0.00133  -0.0001   0.6219   0.5824
   0.000   0.0001   0.00730   0.00135   0.0000   0.6069   0.6073
   0.250   0.0281   0.00733   0.00133   0.0001   0.5824   0.6219
   0.500   0.0556   0.00738   0.00126   0.0003   0.5301   0.6326
   0.750   0.0819   0.00753   0.00118   0.0007   0.4421   0.6420
   1.000   0.1080   0.00771   0.00116   0.0009   0.3564   0.6563
   1.250   0.1293   0.00730   0.00120   0.0024   0.2974   0.8202
   1.500   0.1578   0.00730   0.00133   0.0028   0.2506   0.9576
   1.750   0.1942   0.00756   0.00144   0.0009   0.2045   0.9851
   2.000   0.2328   0.00783   0.00156  -0.0015   0.1640   1.0000
   2.250   0.2589   0.00805   0.00166  -0.0011   0.1358   1.0000
   2.500   0.2849   0.00829   0.00180  -0.0007   0.1080   1.0000
   2.750   0.3110   0.00853   0.00194  -0.0003   0.0862   1.0000
   3.000   0.3373   0.00878   0.00211   0.0001   0.0663   1.0000
   3.250   0.3636   0.00904   0.00231   0.0005   0.0522   1.0000
   3.500   0.3900   0.00929   0.00253   0.0009   0.0417   1.0000
   3.750   0.4164   0.00957   0.00280   0.0012   0.0336   1.0000
   4.000   0.4429   0.00987   0.00308   0.0016   0.0281   1.0000
   4.250   0.4694   0.01020   0.00344   0.0020   0.0245   1.0000
   4.500   0.4960   0.01050   0.00377   0.0023   0.0217   1.0000
   4.750   0.5223   0.01092   0.00422   0.0027   0.0185   1.0000
   5.000   0.5487   0.01137   0.00473   0.0030   0.0170   1.0000
   5.250   0.5750   0.01179   0.00523   0.0034   0.0158   1.0000
   5.500   0.6013   0.01227   0.00577   0.0038   0.0146   1.0000
   5.750   0.6274   0.01273   0.00627   0.0041   0.0134   1.0000
   6.000   0.6526   0.01360   0.00724   0.0045   0.0120   1.0000
   6.250   0.6787   0.01411   0.00786   0.0049   0.0114   1.0000
   6.500   0.7042   0.01483   0.00870   0.0053   0.0108   1.0000
   6.750   0.7295   0.01562   0.00963   0.0058   0.0102   1.0000
   7.000   0.7546   0.01644   0.01059   0.0062   0.0096   1.0000
   7.250   0.7795   0.01726   0.01157   0.0066   0.0091   1.0000
   7.500   0.8040   0.01817   0.01261   0.0070   0.0087   1.0000
   7.750   0.8269   0.01966   0.01431   0.0075   0.0082   1.0000
   8.000   0.8483   0.02163   0.01662   0.0082   0.0078   1.0000
   8.250   0.8709   0.02306   0.01833   0.0087   0.0075   1.0000
   8.500   0.8916   0.02504   0.02064   0.0092   0.0071   1.0000
   8.750   0.9082   0.02820   0.02427   0.0098   0.0068   1.0000
   9.000   0.9186   0.03312   0.02979   0.0100   0.0065   1.0000
  10.000   0.7741   0.10463   0.10240  -0.0346   0.0070   1.0000
  10.250   0.7690   0.11061   0.10836  -0.0374   0.0069   1.0000
<< Back to HT05 (ht05-il)

Polar data table (+)

Polar graphs


<< Back to HT05 (ht05-il)