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


Details Polar file
Airfoil: HQ 0/9 AIRFOIL (hq09-il)
Reynolds number: 500,000
Max Cl/Cd: 51.83 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq09-il-500000.txt
Download as CSV file: xf-hq09-il-500000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6093   0.08404   0.08204   0.0034   1.0000   0.0109
  -9.750  -0.6255   0.07515   0.07318  -0.0009   1.0000   0.0104
  -8.750  -0.8099   0.04371   0.04074  -0.0157   1.0000   0.0078
  -8.500  -0.7987   0.04164   0.03850  -0.0148   1.0000   0.0076
  -8.250  -0.8061   0.03538   0.03183  -0.0132   1.0000   0.0071
  -8.000  -0.8044   0.02974   0.02562  -0.0112   1.0000   0.0069
  -7.750  -0.7940   0.02540   0.02075  -0.0095   1.0000   0.0069
  -7.500  -0.7772   0.02237   0.01729  -0.0081   1.0000   0.0071
  -7.250  -0.7562   0.02047   0.01506  -0.0071   1.0000   0.0074
  -7.000  -0.7400   0.01700   0.01118  -0.0054   1.0000   0.0085
  -6.750  -0.7159   0.01635   0.01049  -0.0050   1.0000   0.0098
  -6.500  -0.6919   0.01554   0.00957  -0.0044   1.0000   0.0114
  -6.250  -0.6651   0.01552   0.00951  -0.0042   1.0000   0.0129
  -6.000  -0.6439   0.01408   0.00800  -0.0033   1.0000   0.0182
  -5.750  -0.6181   0.01375   0.00760  -0.0029   1.0000   0.0209
  -5.500  -0.5962   0.01258   0.00631  -0.0020   1.0000   0.0234
  -5.250  -0.5716   0.01213   0.00584  -0.0016   1.0000   0.0278
  -5.000  -0.5471   0.01161   0.00526  -0.0010   1.0000   0.0304
  -4.750  -0.5225   0.01116   0.00474  -0.0003   1.0000   0.0320
  -4.500  -0.4999   0.01037   0.00389   0.0006   1.0000   0.0370
  -4.250  -0.4757   0.00997   0.00347   0.0013   1.0000   0.0439
  -4.000  -0.4533   0.00937   0.00300   0.0022   1.0000   0.0698
  -3.750  -0.4323   0.00868   0.00262   0.0031   1.0000   0.1319
  -3.500  -0.4123   0.00810   0.00236   0.0042   1.0000   0.2178
  -3.250  -0.3807   0.00752   0.00211   0.0028   0.9953   0.2987
  -3.000  -0.3462   0.00668   0.00187   0.0007   0.9870   0.4511
  -2.750  -0.3103   0.00613   0.00177  -0.0014   0.9786   0.5676
  -2.500  -0.2738   0.00590   0.00172  -0.0032   0.9685   0.6355
  -2.250  -0.2396   0.00580   0.00167  -0.0045   0.9551   0.6705
  -2.000  -0.2096   0.00573   0.00164  -0.0047   0.9385   0.6979
  -1.750  -0.1827   0.00570   0.00161  -0.0042   0.9205   0.7219
  -1.500  -0.1574   0.00567   0.00158  -0.0033   0.9034   0.7443
  -1.250  -0.1315   0.00566   0.00154  -0.0026   0.8878   0.7584
  -1.000  -0.1055   0.00567   0.00151  -0.0020   0.8731   0.7697
  -0.750  -0.0790   0.00568   0.00149  -0.0015   0.8600   0.7814
  -0.500  -0.0525   0.00568   0.00148  -0.0011   0.8476   0.7940
  -0.250  -0.0261   0.00569   0.00147  -0.0006   0.8351   0.8069
   0.000   0.0000   0.00569   0.00147   0.0000   0.8212   0.8212
   0.250   0.0261   0.00569   0.00147   0.0006   0.8069   0.8351
   0.500   0.0525   0.00568   0.00148   0.0011   0.7939   0.8476
   0.750   0.0790   0.00568   0.00149   0.0015   0.7814   0.8600
   1.000   0.1054   0.00567   0.00151   0.0020   0.7697   0.8731
   1.250   0.1315   0.00566   0.00154   0.0026   0.7584   0.8878
   1.500   0.1574   0.00567   0.00158   0.0033   0.7443   0.9034
   1.750   0.1827   0.00570   0.00161   0.0042   0.7218   0.9205
   2.000   0.2096   0.00573   0.00164   0.0047   0.6979   0.9384
   2.250   0.2396   0.00580   0.00167   0.0045   0.6704   0.9552
   2.500   0.2738   0.00590   0.00172   0.0032   0.6352   0.9686
   2.750   0.3103   0.00613   0.00177   0.0014   0.5692   0.9788
   3.000   0.3462   0.00668   0.00187  -0.0007   0.4510   0.9871
   3.250   0.3808   0.00752   0.00211  -0.0028   0.2986   0.9955
   3.500   0.4119   0.00809   0.00236  -0.0041   0.2183   1.0000
   3.750   0.4319   0.00867   0.00261  -0.0031   0.1328   1.0000
   4.000   0.4529   0.00937   0.00299  -0.0021   0.0704   1.0000
   4.250   0.4754   0.00997   0.00347  -0.0012   0.0440   1.0000
   4.500   0.4996   0.01036   0.00388  -0.0005   0.0371   1.0000
   4.750   0.5224   0.01116   0.00474   0.0004   0.0321   1.0000
   5.000   0.5470   0.01161   0.00525   0.0010   0.0304   1.0000
   5.250   0.5715   0.01213   0.00584   0.0016   0.0278   1.0000
   5.500   0.5962   0.01259   0.00632   0.0020   0.0235   1.0000
   5.750   0.6183   0.01373   0.00758   0.0029   0.0208   1.0000
   6.000   0.6440   0.01408   0.00800   0.0033   0.0182   1.0000
   6.250   0.6647   0.01573   0.00973   0.0043   0.0130   1.0000
   6.500   0.6920   0.01554   0.00958   0.0044   0.0113   1.0000
   6.750   0.7161   0.01636   0.01049   0.0050   0.0098   1.0000
   7.000   0.7402   0.01703   0.01122   0.0054   0.0086   1.0000
   7.250   0.7559   0.02070   0.01532   0.0071   0.0074   1.0000
   7.500   0.7773   0.02243   0.01736   0.0081   0.0071   1.0000
   7.750   0.7942   0.02544   0.02079   0.0094   0.0069   1.0000
   8.000   0.8049   0.02973   0.02561   0.0111   0.0069   1.0000
   8.250   0.8060   0.03550   0.03195   0.0131   0.0071   1.0000
   8.500   0.7996   0.04157   0.03842   0.0147   0.0076   1.0000
   8.750   0.8108   0.04368   0.04071   0.0156   0.0078   1.0000
  10.250   0.5984   0.09137   0.08935  -0.0067   0.0114   1.0000
  10.500   0.5942   0.09570   0.09370  -0.0078   0.0134   1.0000
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