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HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il)
Reynolds number: 50,000
Max Cl/Cd: 35.69 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq158-il-50000.txt
Download as CSV file: xf-hq158-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4916   0.10251   0.09584  -0.0029   1.0000   0.2509
  -8.250  -0.4935   0.09989   0.09330  -0.0030   1.0000   0.2684
  -8.000  -0.5005   0.09767   0.09119  -0.0034   1.0000   0.2838
  -7.500  -0.4716   0.08989   0.08339   0.0003   1.0000   0.3308
  -7.250  -0.4843   0.08844   0.08207   0.0010   1.0000   0.3523
  -7.000  -0.4620   0.08405   0.07767   0.0030   1.0000   0.3768
  -6.750  -0.4513   0.08062   0.07428   0.0046   1.0000   0.4008
  -6.500  -0.4441   0.07754   0.07127   0.0064   1.0000   0.4266
  -5.250  -0.4480   0.04325   0.03540  -0.0389   1.0000   0.1248
  -5.000  -0.4274   0.03923   0.03082  -0.0385   1.0000   0.1140
  -4.750  -0.4045   0.03586   0.02641  -0.0377   1.0000   0.1063
  -4.500  -0.3826   0.03273   0.02290  -0.0366   1.0000   0.1052
  -4.250  -0.3590   0.03017   0.01978  -0.0352   1.0000   0.1066
  -4.000  -0.3374   0.02787   0.01741  -0.0341   1.0000   0.1178
  -3.750  -0.3137   0.02563   0.01501  -0.0328   1.0000   0.1278
  -3.500  -0.2896   0.02370   0.01290  -0.0311   1.0000   0.1470
  -3.250  -0.2669   0.02173   0.01117  -0.0295   1.0000   0.1840
  -3.000  -0.2461   0.01940   0.00945  -0.0280   1.0000   0.2726
  -2.750  -0.0982   0.01738   0.00897  -0.0350   1.0000   1.0000
  -2.500  -0.0987   0.01698   0.00847  -0.0320   1.0000   1.0000
  -2.250  -0.1036   0.01663   0.00804  -0.0282   1.0000   1.0000
  -2.000  -0.1104   0.01633   0.00765  -0.0239   1.0000   1.0000
  -1.750  -0.1114   0.01613   0.00727  -0.0205   1.0000   1.0000
  -1.500  -0.1014   0.01606   0.00694  -0.0186   1.0000   1.0000
  -1.250  -0.0856   0.01607   0.00664  -0.0176   1.0000   1.0000
  -1.000  -0.0673   0.01616   0.00648  -0.0169   1.0000   1.0000
  -0.750  -0.0478   0.01629   0.00639  -0.0164   1.0000   1.0000
  -0.500  -0.0278   0.01646   0.00636  -0.0159   1.0000   1.0000
  -0.250  -0.0075   0.01668   0.00639  -0.0155   1.0000   1.0000
   0.000   0.0129   0.01692   0.00645  -0.0151   1.0000   1.0000
   0.250   0.0332   0.01720   0.00660  -0.0147   1.0000   1.0000
   0.500   0.0536   0.01752   0.00680  -0.0143   1.0000   1.0000
   0.750   0.0737   0.01787   0.00706  -0.0140   1.0000   1.0000
   1.000   0.0938   0.01826   0.00738  -0.0137   1.0000   1.0000
   1.250   0.1136   0.01869   0.00776  -0.0135   1.0000   1.0000
   1.500   0.1333   0.01916   0.00819  -0.0132   1.0000   1.0000
   1.750   0.1526   0.01967   0.00869  -0.0131   1.0000   1.0000
   2.000   0.1716   0.02024   0.00926  -0.0129   1.0000   1.0000
   2.250   0.1902   0.02086   0.00990  -0.0128   1.0000   1.0000
   2.500   0.2083   0.02155   0.01062  -0.0128   1.0000   1.0000
   2.750   0.2264   0.02231   0.01147  -0.0129   0.9998   1.0000
   3.000   0.2986   0.02392   0.01329  -0.0231   0.9717   1.0000
   3.250   0.3693   0.02511   0.01477  -0.0321   0.9422   1.0000
   3.500   0.4323   0.02585   0.01591  -0.0390   0.9120   1.0000
   3.750   0.4928   0.02621   0.01665  -0.0445   0.8788   1.0000
   4.000   0.5660   0.02587   0.01685  -0.0506   0.8425   1.0000
   4.250   0.6266   0.02493   0.01650  -0.0528   0.8016   1.0000
   4.500   0.6732   0.02358   0.01556  -0.0513   0.7538   1.0000
   4.750   0.7073   0.02213   0.01434  -0.0471   0.6972   1.0000
   5.000   0.7306   0.02118   0.01338  -0.0419   0.6262   1.0000
   5.250   0.7485   0.02097   0.01299  -0.0370   0.5383   1.0000
   5.500   0.7633   0.02176   0.01315  -0.0327   0.4376   1.0000
   5.750   0.7749   0.02362   0.01422  -0.0291   0.3266   1.0000
   6.000   0.7924   0.02636   0.01623  -0.0268   0.2380   1.0000
   6.250   0.8125   0.02851   0.01814  -0.0252   0.1858   1.0000
   6.500   0.8393   0.03122   0.02071  -0.0242   0.1568   1.0000
   6.750   0.8613   0.03358   0.02318  -0.0229   0.1316   1.0000
   7.000   0.8863   0.03715   0.02706  -0.0217   0.1184   1.0000
   7.250   0.9057   0.04050   0.03047  -0.0206   0.1029   1.0000
   7.500   0.9207   0.04463   0.03560  -0.0185   0.0984   1.0000
   7.750   0.9390   0.04852   0.03948  -0.0177   0.0905   1.0000
   8.000   0.9444   0.05272   0.04446  -0.0156   0.0879   1.0000
   8.250   0.9480   0.05744   0.04975  -0.0140   0.0873   1.0000
   8.500   0.9504   0.06257   0.05526  -0.0129   0.0884   1.0000
   8.750   0.9554   0.06801   0.06089  -0.0121   0.0898   1.0000
   9.000   0.9234   0.07361   0.06721  -0.0118   0.0951   1.0000
   9.250   0.9046   0.07921   0.07298  -0.0123   0.0983   1.0000
   9.500   0.8933   0.08434   0.07815  -0.0126   0.1010   1.0000
   9.750   0.7427   0.09047   0.08460  -0.0183   0.1237   1.0000
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