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HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il)
Reynolds number: 50,000
Max Cl/Cd: 38.18 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259-il-50000.txt
Download as CSV file: xf-hq259-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4114   0.10941   0.10264  -0.0216   1.0000   0.2017
  -8.750  -0.4177   0.10768   0.10102  -0.0225   1.0000   0.2119
  -8.500  -0.4143   0.10455   0.09796  -0.0221   1.0000   0.2256
  -8.250  -0.4097   0.10142   0.09489  -0.0214   1.0000   0.2392
  -8.000  -0.4127   0.09902   0.09260  -0.0209   1.0000   0.2535
  -7.750  -0.3958   0.09461   0.08821  -0.0189   1.0000   0.2716
  -7.500  -0.3900   0.09172   0.08538  -0.0175   1.0000   0.2885
  -7.250  -0.3950   0.08964   0.08342  -0.0161   1.0000   0.3068
  -7.000  -0.4087   0.08828   0.08222  -0.0141   1.0000   0.3224
  -6.750  -0.3918   0.08446   0.07843  -0.0116   1.0000   0.3479
  -6.500  -0.3823   0.08141   0.07538  -0.0091   1.0000   0.3715
  -6.250  -0.3833   0.07925   0.07333  -0.0060   1.0000   0.3954
  -6.000  -0.3754   0.07650   0.07065  -0.0028   1.0000   0.4258
  -5.750  -0.3735   0.07426   0.06850   0.0009   1.0000   0.4562
  -5.500  -0.3789   0.07275   0.06711   0.0055   1.0000   0.4864
  -4.750  -0.4011   0.04923   0.04253  -0.0426   1.0000   0.1727
  -4.500  -0.3693   0.04331   0.03566  -0.0457   1.0000   0.1288
  -4.250  -0.3435   0.03927   0.03094  -0.0463   1.0000   0.1153
  -4.000  -0.3200   0.03589   0.02720  -0.0462   1.0000   0.1107
  -3.750  -0.2946   0.03316   0.02386  -0.0460   1.0000   0.1112
  -3.500  -0.2680   0.03113   0.02110  -0.0456   1.0000   0.1166
  -3.250  -0.2433   0.02860   0.01848  -0.0449   1.0000   0.1221
  -3.000  -0.2169   0.02673   0.01622  -0.0440   1.0000   0.1320
  -2.750  -0.1927   0.02511   0.01457  -0.0430   1.0000   0.1564
  -2.500  -0.1697   0.02351   0.01313  -0.0417   1.0000   0.1918
  -2.250  -0.1431   0.02174   0.01163  -0.0411   1.0000   0.2512
  -2.000  -0.1428   0.01900   0.01177  -0.0328   1.0000   0.7637
  -1.750  -0.1233   0.01828   0.01099  -0.0280   1.0000   1.0000
  -1.500  -0.0997   0.01839   0.01049  -0.0286   1.0000   1.0000
  -1.250  -0.0760   0.01858   0.01022  -0.0292   1.0000   1.0000
  -1.000  -0.0525   0.01882   0.01007  -0.0296   1.0000   1.0000
  -0.750  -0.0293   0.01911   0.01002  -0.0300   1.0000   1.0000
  -0.500  -0.0066   0.01944   0.01006  -0.0303   1.0000   1.0000
  -0.250   0.0157   0.01982   0.01014  -0.0305   1.0000   1.0000
   0.000   0.0376   0.02023   0.01032  -0.0306   1.0000   1.0000
   0.250   0.0590   0.02069   0.01059  -0.0307   1.0000   1.0000
   0.500   0.0802   0.02119   0.01091  -0.0308   1.0000   1.0000
   0.750   0.1009   0.02173   0.01128  -0.0309   1.0000   1.0000
   1.000   0.1212   0.02232   0.01174  -0.0310   1.0000   1.0000
   1.250   0.1412   0.02296   0.01227  -0.0311   1.0000   1.0000
   1.500   0.1607   0.02364   0.01287  -0.0312   1.0000   1.0000
   1.750   0.1798   0.02438   0.01354  -0.0313   1.0000   1.0000
   2.000   0.1985   0.02518   0.01430  -0.0314   1.0000   1.0000
   2.250   0.2440   0.02660   0.01569  -0.0367   0.9860   1.0000
   2.500   0.2989   0.02813   0.01723  -0.0434   0.9664   1.0000
   2.750   0.3502   0.02935   0.01850  -0.0490   0.9445   1.0000
   3.000   0.3988   0.03038   0.01963  -0.0537   0.9221   1.0000
   3.250   0.4544   0.03131   0.02068  -0.0592   0.9006   1.0000
   3.500   0.4951   0.03196   0.02147  -0.0619   0.8768   1.0000
   3.750   0.5400   0.03246   0.02220  -0.0648   0.8526   1.0000
   4.000   0.5890   0.03267   0.02263  -0.0678   0.8278   1.0000
   4.250   0.6423   0.03241   0.02265  -0.0706   0.8021   1.0000
   4.500   0.6987   0.03159   0.02223  -0.0730   0.7752   1.0000
   4.750   0.7480   0.03044   0.02141  -0.0734   0.7467   1.0000
   5.000   0.7952   0.02892   0.02022  -0.0728   0.7155   1.0000
   5.250   0.8419   0.02705   0.01870  -0.0714   0.6805   1.0000
   5.500   0.8776   0.02573   0.01755  -0.0688   0.6371   1.0000
   5.750   0.9110   0.02471   0.01652  -0.0661   0.5873   1.0000
   6.000   0.9363   0.02461   0.01628  -0.0633   0.5324   1.0000
   6.250   0.9594   0.02513   0.01657  -0.0607   0.4781   1.0000
   6.500   0.9799   0.02618   0.01745  -0.0585   0.4275   1.0000
   6.750   1.0019   0.02745   0.01856  -0.0567   0.3824   1.0000
   7.000   1.0228   0.02895   0.01996  -0.0550   0.3415   1.0000
   7.250   1.0423   0.03065   0.02168  -0.0533   0.3039   1.0000
   7.500   1.0606   0.03245   0.02347  -0.0514   0.2664   1.0000
   7.750   1.0750   0.03428   0.02523  -0.0491   0.2244   1.0000
   8.000   1.0769   0.03551   0.02618  -0.0454   0.1725   1.0000
   8.250   1.0764   0.03750   0.02772  -0.0418   0.1236   1.0000
   8.500   1.0863   0.04037   0.03059  -0.0393   0.0946   1.0000
   8.750   1.1075   0.04419   0.03462  -0.0379   0.0823   1.0000
   9.000   1.1270   0.04780   0.03840  -0.0368   0.0752   1.0000
   9.250   1.1368   0.05197   0.04308  -0.0348   0.0715   1.0000
   9.500   1.1398   0.05595   0.04765  -0.0325   0.0689   1.0000
   9.750   1.1400   0.05990   0.05213  -0.0303   0.0667   1.0000
  10.000   1.1371   0.06388   0.05646  -0.0282   0.0653   1.0000
  10.250   1.1276   0.06809   0.06100  -0.0261   0.0649   1.0000
  10.500   1.1053   0.07247   0.06575  -0.0236   0.0660   1.0000
  10.750   1.0804   0.07714   0.07068  -0.0222   0.0674   1.0000
  11.000   1.0555   0.08243   0.07616  -0.0225   0.0688   1.0000
  11.250   1.0337   0.08826   0.08204  -0.0241   0.0705   1.0000
  11.500   1.0132   0.09466   0.08853  -0.0268   0.0717   1.0000
  11.750   0.9990   0.10120   0.09510  -0.0296   0.0728   1.0000
  12.000   0.9854   0.10824   0.10218  -0.0330   0.0735   1.0000
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