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HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 500,000
Max Cl/Cd: 53.38 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq09-il-500000-n5.txt
Download as CSV file: xf-hq09-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/9 AIRFOIL                                  
  
 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  -0.8622   0.04430   0.04164  -0.0162   1.0000   0.0025
  -9.250  -0.8797   0.03761   0.03449  -0.0145   1.0000   0.0024
  -9.000  -0.8850   0.03199   0.02832  -0.0126   1.0000   0.0024
  -8.750  -0.8796   0.02778   0.02358  -0.0109   1.0000   0.0024
  -8.500  -0.8672   0.02468   0.02005  -0.0095   1.0000   0.0024
  -8.250  -0.8508   0.02228   0.01728  -0.0083   1.0000   0.0024
  -8.000  -0.8323   0.02023   0.01491  -0.0072   1.0000   0.0024
  -7.750  -0.8120   0.01862   0.01304  -0.0062   1.0000   0.0024
  -7.500  -0.7907   0.01727   0.01147  -0.0054   1.0000   0.0024
  -7.250  -0.7687   0.01610   0.01011  -0.0046   1.0000   0.0025
  -7.000  -0.7459   0.01513   0.00898  -0.0039   1.0000   0.0025
  -6.750  -0.7224   0.01431   0.00802  -0.0033   1.0000   0.0026
  -6.500  -0.6990   0.01345   0.00696  -0.0026   1.0000   0.0028
  -6.250  -0.6747   0.01280   0.00618  -0.0020   1.0000   0.0031
  -6.000  -0.6501   0.01223   0.00556  -0.0015   1.0000   0.0041
  -5.750  -0.6256   0.01172   0.00510  -0.0010   1.0000   0.0081
  -5.500  -0.6000   0.01146   0.00480  -0.0007   1.0000   0.0110
  -5.250  -0.5750   0.01107   0.00442  -0.0003   1.0000   0.0145
  -5.000  -0.5498   0.01076   0.00409   0.0001   1.0000   0.0164
  -4.750  -0.5246   0.01050   0.00380   0.0005   1.0000   0.0192
  -4.500  -0.4994   0.01018   0.00347   0.0009   0.9995   0.0248
  -4.250  -0.4662   0.00986   0.00316  -0.0005   0.9869   0.0321
  -4.000  -0.4322   0.00953   0.00287  -0.0020   0.9737   0.0450
  -3.750  -0.3982   0.00920   0.00260  -0.0035   0.9601   0.0652
  -3.500  -0.3660   0.00887   0.00236  -0.0047   0.9444   0.0921
  -3.250  -0.3378   0.00837   0.00210  -0.0051   0.9255   0.1569
  -3.000  -0.3120   0.00796   0.00188  -0.0048   0.9041   0.2253
  -2.750  -0.2875   0.00756   0.00167  -0.0042   0.8841   0.2917
  -2.500  -0.2637   0.00709   0.00149  -0.0036   0.8657   0.3868
  -2.250  -0.2388   0.00679   0.00135  -0.0031   0.8491   0.4522
  -2.000  -0.2145   0.00645   0.00125  -0.0024   0.8332   0.5333
  -1.750  -0.1891   0.00625   0.00121  -0.0019   0.8192   0.5963
  -1.250  -0.1362   0.00610   0.00113  -0.0011   0.7934   0.6624
  -1.000  -0.1089   0.00608   0.00109  -0.0009   0.7812   0.6755
  -0.750  -0.0816   0.00606   0.00106  -0.0007   0.7704   0.6895
  -0.500  -0.0543   0.00604   0.00104  -0.0005   0.7590   0.7038
  -0.250  -0.0270   0.00603   0.00103  -0.0003   0.7479   0.7179
   0.000   0.0000   0.00602   0.00103   0.0000   0.7338   0.7339
   0.250   0.0270   0.00603   0.00103   0.0003   0.7179   0.7479
   0.500   0.0543   0.00604   0.00104   0.0005   0.7037   0.7590
   0.750   0.0816   0.00606   0.00106   0.0007   0.6894   0.7705
   1.000   0.1089   0.00608   0.00109   0.0009   0.6754   0.7812
   1.250   0.1363   0.00610   0.00113   0.0011   0.6624   0.7933
   1.750   0.1892   0.00625   0.00121   0.0019   0.5963   0.8191
   2.000   0.2145   0.00645   0.00125   0.0024   0.5328   0.8332
   2.250   0.2389   0.00679   0.00135   0.0031   0.4524   0.8491
   2.500   0.2637   0.00709   0.00149   0.0036   0.3870   0.8658
   2.750   0.2875   0.00756   0.00167   0.0042   0.2918   0.8842
   3.000   0.3120   0.00796   0.00188   0.0048   0.2251   0.9042
   3.250   0.3378   0.00837   0.00210   0.0050   0.1566   0.9257
   3.500   0.3660   0.00887   0.00236   0.0047   0.0922   0.9445
   3.750   0.3982   0.00920   0.00260   0.0036   0.0656   0.9601
   4.000   0.4321   0.00953   0.00287   0.0020   0.0452   0.9737
   4.250   0.4662   0.00987   0.00317   0.0005   0.0317   0.9870
   4.500   0.4994   0.01019   0.00347  -0.0009   0.0247   0.9996
   4.750   0.5245   0.01050   0.00381  -0.0005   0.0194   1.0000
   5.000   0.5497   0.01076   0.00408  -0.0001   0.0163   1.0000
   5.250   0.5750   0.01106   0.00441   0.0003   0.0144   1.0000
   5.500   0.6000   0.01146   0.00482   0.0007   0.0112   1.0000
   5.750   0.6256   0.01172   0.00510   0.0010   0.0082   1.0000
   6.000   0.6501   0.01224   0.00557   0.0015   0.0040   1.0000
   6.250   0.6747   0.01280   0.00618   0.0020   0.0031   1.0000
   6.500   0.6990   0.01347   0.00698   0.0026   0.0027   1.0000
   6.750   0.7225   0.01431   0.00802   0.0032   0.0026   1.0000
   7.000   0.7460   0.01513   0.00898   0.0038   0.0025   1.0000
   7.250   0.7688   0.01612   0.01013   0.0046   0.0025   1.0000
   7.500   0.7910   0.01727   0.01148   0.0053   0.0024   1.0000
   7.750   0.8124   0.01861   0.01303   0.0062   0.0024   1.0000
   8.000   0.8326   0.02024   0.01492   0.0071   0.0024   1.0000
   8.250   0.8513   0.02222   0.01721   0.0082   0.0024   1.0000
   8.500   0.8673   0.02478   0.02016   0.0094   0.0024   1.0000
   8.750   0.8798   0.02786   0.02368   0.0108   0.0024   1.0000
   9.000   0.8851   0.03212   0.02846   0.0125   0.0024   1.0000
   9.250   0.8806   0.03757   0.03444   0.0144   0.0024   1.0000
   9.500   0.8609   0.04467   0.04204   0.0161   0.0025   1.0000
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