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HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il)
Reynolds number: 50,000
Max Cl/Cd: 39.21 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq258-il-50000.txt
Download as CSV file: xf-hq258-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4118   0.09885   0.09228  -0.0191   1.0000   0.2174
  -7.750  -0.4114   0.09602   0.08955  -0.0189   1.0000   0.2290
  -7.500  -0.4324   0.09551   0.08922  -0.0196   1.0000   0.2374
  -7.250  -0.4104   0.09054   0.08425  -0.0172   1.0000   0.2558
  -7.000  -0.4088   0.08785   0.08165  -0.0156   1.0000   0.2729
  -6.750  -0.4154   0.08587   0.07980  -0.0139   1.0000   0.2894
  -6.500  -0.4287   0.08440   0.07849  -0.0133   1.0000   0.3052
  -6.250  -0.4224   0.08139   0.07556  -0.0100   1.0000   0.3319
  -6.000  -0.4208   0.07896   0.07323  -0.0069   1.0000   0.3591
  -5.750  -0.4009   0.07532   0.06960  -0.0022   1.0000   0.3942
  -5.500  -0.3932   0.07280   0.06716   0.0022   1.0000   0.4340
  -4.250  -0.3781   0.04975   0.04380  -0.0330   1.0000   0.2692
  -4.000  -0.3075   0.03960   0.03172  -0.0480   1.0000   0.1242
  -3.750  -0.2785   0.03638   0.02771  -0.0487   1.0000   0.1144
  -3.500  -0.2529   0.03354   0.02441  -0.0487   1.0000   0.1139
  -3.250  -0.2263   0.03088   0.02133  -0.0484   1.0000   0.1113
  -3.000  -0.1990   0.02857   0.01857  -0.0478   1.0000   0.1098
  -2.750  -0.1718   0.02659   0.01608  -0.0469   1.0000   0.1124
  -2.500  -0.1466   0.02487   0.01421  -0.0458   1.0000   0.1248
  -2.250  -0.1225   0.02329   0.01261  -0.0445   1.0000   0.1486
  -2.000  -0.0960   0.02149   0.01101  -0.0438   1.0000   0.2056
  -1.750  -0.0855   0.01783   0.01039  -0.0381   1.0000   0.7607
  -1.500  -0.0652   0.01717   0.00937  -0.0345   1.0000   1.0000
  -1.250  -0.0410   0.01734   0.00898  -0.0346   1.0000   1.0000
  -1.000  -0.0173   0.01756   0.00875  -0.0348   1.0000   1.0000
  -0.750   0.0059   0.01781   0.00857  -0.0348   1.0000   1.0000
  -0.500   0.0287   0.01811   0.00855  -0.0349   1.0000   1.0000
  -0.250   0.0511   0.01844   0.00860  -0.0348   1.0000   1.0000
   0.000   0.0732   0.01880   0.00873  -0.0348   1.0000   1.0000
   0.250   0.0949   0.01921   0.00894  -0.0348   1.0000   1.0000
   0.500   0.1162   0.01965   0.00915  -0.0347   1.0000   1.0000
   0.750   0.1372   0.02013   0.00949  -0.0346   1.0000   1.0000
   1.000   0.1578   0.02065   0.00989  -0.0346   1.0000   1.0000
   1.250   0.1781   0.02122   0.01036  -0.0345   1.0000   1.0000
   1.500   0.1979   0.02184   0.01090  -0.0345   1.0000   1.0000
   1.750   0.2174   0.02250   0.01151  -0.0345   1.0000   1.0000
   2.000   0.2364   0.02323   0.01219  -0.0345   1.0000   1.0000
   2.250   0.2549   0.02401   0.01296  -0.0346   1.0000   1.0000
   2.500   0.2730   0.02487   0.01381  -0.0347   1.0000   1.0000
   2.750   0.3315   0.02651   0.01551  -0.0424   0.9791   1.0000
   3.000   0.3929   0.02794   0.01710  -0.0500   0.9553   1.0000
   3.250   0.4457   0.02895   0.01826  -0.0556   0.9304   1.0000
   3.500   0.4975   0.02975   0.01925  -0.0605   0.9049   1.0000
   3.750   0.5498   0.03029   0.02003  -0.0649   0.8783   1.0000
   4.000   0.6054   0.03044   0.02058  -0.0690   0.8504   1.0000
   4.250   0.6695   0.02991   0.02045  -0.0733   0.8214   1.0000
   4.500   0.7197   0.02914   0.02007  -0.0746   0.7890   1.0000
   4.750   0.7769   0.02744   0.01892  -0.0752   0.7555   1.0000
   5.000   0.8218   0.02572   0.01757  -0.0735   0.7154   1.0000
   5.250   0.8577   0.02429   0.01638  -0.0705   0.6656   1.0000
   5.500   0.8890   0.02328   0.01539  -0.0669   0.6048   1.0000
   5.750   0.9124   0.02327   0.01520  -0.0634   0.5351   1.0000
   6.000   0.9332   0.02401   0.01575  -0.0604   0.4669   1.0000
   6.250   0.9535   0.02526   0.01668  -0.0579   0.4051   1.0000
   6.500   0.9726   0.02684   0.01805  -0.0557   0.3483   1.0000
   6.750   0.9889   0.02854   0.01957  -0.0532   0.2911   1.0000
   7.000   0.9971   0.03036   0.02105  -0.0499   0.2215   1.0000
   7.250   0.9947   0.03250   0.02262  -0.0457   0.1420   1.0000
   7.500   1.0081   0.03563   0.02520  -0.0436   0.0975   1.0000
   7.750   1.0350   0.03931   0.02914  -0.0424   0.0818   1.0000
   8.000   1.0568   0.04271   0.03257  -0.0415   0.0717   1.0000
   8.250   1.0732   0.04615   0.03674  -0.0396   0.0667   1.0000
   8.500   1.0878   0.05022   0.04130  -0.0379   0.0642   1.0000
   8.750   1.0968   0.05468   0.04630  -0.0360   0.0638   1.0000
   9.000   1.0984   0.05944   0.05162  -0.0339   0.0642   1.0000
   9.250   1.0937   0.06435   0.05702  -0.0318   0.0651   1.0000
   9.500   1.0841   0.06920   0.06227  -0.0300   0.0662   1.0000
   9.750   1.0699   0.07397   0.06735  -0.0284   0.0672   1.0000
  10.000   1.0516   0.07833   0.07191  -0.0269   0.0682   1.0000
  10.250   1.0317   0.08297   0.07669  -0.0265   0.0690   1.0000
  10.500   1.0127   0.08811   0.08194  -0.0274   0.0698   1.0000
  10.750   0.9975   0.09373   0.08763  -0.0293   0.0707   1.0000
  11.000   0.9910   0.09961   0.09353  -0.0308   0.0717   1.0000
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