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HT08 (ht08-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: HT08 (ht08-il)
Reynolds number: 500,000
Max Cl/Cd: 48.86 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ht08-il-500000-n5.txt
Download as CSV file: xf-ht08-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HT08                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.8048   0.09437   0.09223   0.0363   1.0000   0.0068
  -9.250  -0.8161   0.08705   0.08495   0.0316   1.0000   0.0068
  -8.750  -0.9114   0.03370   0.03040  -0.0044   1.0000   0.0064
  -8.500  -0.9010   0.02757   0.02353  -0.0047   1.0000   0.0067
  -8.250  -0.8811   0.02477   0.02027  -0.0045   1.0000   0.0070
  -8.000  -0.8591   0.02270   0.01782  -0.0043   1.0000   0.0073
  -7.750  -0.8374   0.02039   0.01517  -0.0041   1.0000   0.0077
  -7.500  -0.8129   0.01920   0.01382  -0.0039   1.0000   0.0081
  -7.250  -0.7878   0.01814   0.01258  -0.0037   1.0000   0.0084
  -7.000  -0.7624   0.01710   0.01137  -0.0035   1.0000   0.0089
  -6.750  -0.7367   0.01610   0.01017  -0.0032   1.0000   0.0094
  -6.500  -0.7107   0.01518   0.00909  -0.0030   1.0000   0.0100
  -6.250  -0.6844   0.01441   0.00817  -0.0028   1.0000   0.0105
  -6.000  -0.6582   0.01351   0.00715  -0.0026   1.0000   0.0112
  -5.750  -0.6316   0.01293   0.00653  -0.0025   1.0000   0.0124
  -5.500  -0.6046   0.01245   0.00597  -0.0023   1.0000   0.0135
  -5.250  -0.5775   0.01192   0.00536  -0.0022   1.0000   0.0145
  -5.000  -0.5502   0.01147   0.00483  -0.0021   1.0000   0.0155
  -4.750  -0.5231   0.01086   0.00417  -0.0019   1.0000   0.0174
  -4.500  -0.4956   0.01051   0.00381  -0.0019   1.0000   0.0199
  -4.250  -0.4680   0.01020   0.00344  -0.0018   1.0000   0.0222
  -4.000  -0.4404   0.00979   0.00302  -0.0017   1.0000   0.0260
  -3.750  -0.4128   0.00951   0.00273  -0.0016   1.0000   0.0308
  -3.500  -0.3851   0.00920   0.00245  -0.0016   1.0000   0.0387
  -3.250  -0.3573   0.00893   0.00220  -0.0015   1.0000   0.0484
  -3.000  -0.3296   0.00866   0.00199  -0.0015   1.0000   0.0635
  -2.750  -0.3019   0.00838   0.00180  -0.0015   1.0000   0.0846
  -2.500  -0.2741   0.00811   0.00164  -0.0015   1.0000   0.1118
  -2.250  -0.2465   0.00782   0.00148  -0.0015   1.0000   0.1472
  -2.000  -0.2188   0.00753   0.00136  -0.0015   1.0000   0.1886
  -1.750  -0.1911   0.00723   0.00125  -0.0015   1.0000   0.2387
  -1.500  -0.1635   0.00693   0.00117  -0.0015   1.0000   0.2955
  -1.250  -0.1360   0.00663   0.00109  -0.0015   1.0000   0.3599
  -1.000  -0.1087   0.00628   0.00104  -0.0015   1.0000   0.4359
  -0.750  -0.0816   0.00589   0.00100  -0.0014   1.0000   0.5296
  -0.500  -0.0553   0.00546   0.00099  -0.0011   1.0000   0.6385
  -0.250  -0.0213   0.00504   0.00103  -0.0024   0.9536   0.7610
   0.000   0.0001   0.00490   0.00107   0.0000   0.8630   0.8632
   0.250   0.0215   0.00505   0.00103   0.0023   0.7602   0.9541
   0.500   0.0554   0.00546   0.00099   0.0011   0.6385   1.0000
   0.750   0.0818   0.00589   0.00100   0.0014   0.5298   1.0000
   1.000   0.1088   0.00629   0.00104   0.0015   0.4356   1.0000
   1.250   0.1362   0.00663   0.00109   0.0015   0.3598   1.0000
   1.500   0.1637   0.00693   0.00117   0.0015   0.2957   1.0000
   1.750   0.1913   0.00724   0.00125   0.0015   0.2384   1.0000
   2.000   0.2189   0.00753   0.00136   0.0015   0.1886   1.0000
   2.250   0.2466   0.00782   0.00148   0.0015   0.1470   1.0000
   2.500   0.2743   0.00811   0.00164   0.0015   0.1115   1.0000
   2.750   0.3020   0.00838   0.00180   0.0015   0.0846   1.0000
   3.000   0.3297   0.00866   0.00199   0.0015   0.0637   1.0000
   3.250   0.3575   0.00893   0.00220   0.0015   0.0483   1.0000
   3.500   0.3852   0.00920   0.00245   0.0016   0.0386   1.0000
   3.750   0.4129   0.00951   0.00273   0.0016   0.0307   1.0000
   4.000   0.4406   0.00980   0.00302   0.0017   0.0260   1.0000
   4.250   0.4682   0.01020   0.00344   0.0018   0.0222   1.0000
   4.500   0.4958   0.01051   0.00381   0.0019   0.0199   1.0000
   4.750   0.5232   0.01086   0.00417   0.0019   0.0174   1.0000
   5.000   0.5504   0.01147   0.00484   0.0021   0.0155   1.0000
   5.250   0.5777   0.01192   0.00537   0.0022   0.0145   1.0000
   5.500   0.6048   0.01245   0.00598   0.0023   0.0134   1.0000
   5.750   0.6318   0.01293   0.00653   0.0025   0.0124   1.0000
   6.000   0.6584   0.01351   0.00715   0.0026   0.0112   1.0000
   6.250   0.6846   0.01442   0.00818   0.0027   0.0105   1.0000
   6.500   0.7110   0.01519   0.00910   0.0030   0.0100   1.0000
   6.750   0.7370   0.01611   0.01019   0.0032   0.0094   1.0000
   7.000   0.7626   0.01712   0.01139   0.0034   0.0089   1.0000
   7.250   0.7881   0.01816   0.01261   0.0037   0.0084   1.0000
   7.500   0.8131   0.01923   0.01384   0.0039   0.0081   1.0000
   7.750   0.8377   0.02039   0.01518   0.0040   0.0077   1.0000
   8.000   0.8593   0.02276   0.01789   0.0043   0.0072   1.0000
   8.250   0.8815   0.02476   0.02024   0.0045   0.0070   1.0000
   8.500   0.9014   0.02755   0.02351   0.0047   0.0066   1.0000
   8.750   0.9118   0.03372   0.03042   0.0043   0.0064   1.0000
   9.250   0.8163   0.08725   0.08515  -0.0319   0.0068   1.0000
   9.500   0.8052   0.09454   0.09240  -0.0366   0.0068   1.0000
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