Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/7 AIRFOIL (hq07-il)
Reynolds number: 500,000
Max Cl/Cd: 46.12 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq07-il-500000-n5.txt
Download as CSV file: xf-hq07-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/7 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.6655   0.12429   0.12198   0.0256   1.0000   0.0081
 -10.750  -0.6650   0.11949   0.11719   0.0236   1.0000   0.0071
 -10.500  -0.6652   0.11433   0.11204   0.0213   1.0000   0.0063
 -10.250  -0.6701   0.10759   0.10534   0.0183   1.0000   0.0054
 -10.000  -0.6792   0.09916   0.09694   0.0145   1.0000   0.0045
  -9.750  -0.6806   0.09374   0.09155   0.0118   1.0000   0.0043
  -9.500  -0.6816   0.08872   0.08655   0.0090   1.0000   0.0043
  -9.250  -0.6800   0.08425   0.08210   0.0065   1.0000   0.0041
  -9.000  -0.6830   0.07879   0.07666   0.0028   1.0000   0.0041
  -8.750  -0.6918   0.07131   0.06922  -0.0041   1.0000   0.0041
  -8.500  -0.7084   0.06480   0.06272  -0.0089   1.0000   0.0041
  -8.250  -0.7100   0.06071   0.05854  -0.0101   1.0000   0.0040
  -8.000  -0.7133   0.05628   0.05397  -0.0111   1.0000   0.0039
  -7.750  -0.7197   0.04997   0.04745  -0.0119   1.0000   0.0038
  -7.500  -0.7199   0.04381   0.04100  -0.0118   1.0000   0.0037
  -7.250  -0.7161   0.03760   0.03442  -0.0111   1.0000   0.0036
  -7.000  -0.7077   0.03190   0.02827  -0.0099   1.0000   0.0035
  -6.750  -0.6949   0.02697   0.02276  -0.0086   1.0000   0.0033
  -6.500  -0.6778   0.02306   0.01835  -0.0073   1.0000   0.0032
  -6.250  -0.6578   0.01994   0.01479  -0.0061   1.0000   0.0032
  -6.000  -0.6358   0.01752   0.01199  -0.0050   1.0000   0.0031
  -5.750  -0.6129   0.01560   0.00979  -0.0041   1.0000   0.0031
  -5.500  -0.5896   0.01412   0.00808  -0.0032   1.0000   0.0030
  -5.250  -0.5661   0.01292   0.00670  -0.0024   1.0000   0.0030
  -5.000  -0.5421   0.01196   0.00558  -0.0016   1.0000   0.0030
  -4.750  -0.5174   0.01122   0.00470  -0.0010   1.0000   0.0031
  -4.500  -0.4921   0.01067   0.00402  -0.0005   1.0000   0.0032
  -4.250  -0.4665   0.01026   0.00350   0.0000   1.0000   0.0033
  -4.000  -0.4408   0.00993   0.00310   0.0004   1.0000   0.0037
  -3.750  -0.4156   0.00955   0.00268   0.0009   1.0000   0.0102
  -3.500  -0.3904   0.00930   0.00242   0.0014   1.0000   0.0157
  -3.250  -0.3598   0.00909   0.00222   0.0006   0.9935   0.0213
  -3.000  -0.3271   0.00851   0.00196  -0.0009   0.9811   0.0796
  -2.750  -0.2942   0.00787   0.00171  -0.0025   0.9678   0.1746
  -2.500  -0.2621   0.00726   0.00149  -0.0039   0.9510   0.2795
  -2.250  -0.2336   0.00656   0.00130  -0.0045   0.9289   0.4238
  -2.000  -0.2080   0.00618   0.00112  -0.0041   0.9030   0.5089
  -1.500  -0.1591   0.00577   0.00103  -0.0024   0.8600   0.6496
  -1.250  -0.1333   0.00569   0.00101  -0.0018   0.8438   0.6832
  -1.000  -0.1067   0.00568   0.00097  -0.0014   0.8277   0.6996
  -0.750  -0.0800   0.00568   0.00095  -0.0011   0.8125   0.7153
  -0.500  -0.0532   0.00566   0.00093  -0.0008   0.7991   0.7322
  -0.250  -0.0265   0.00564   0.00091  -0.0004   0.7847   0.7487
   0.000   0.0000   0.00564   0.00091   0.0000   0.7673   0.7674
   0.250   0.0265   0.00564   0.00091   0.0004   0.7488   0.7846
   0.500   0.0532   0.00566   0.00093   0.0008   0.7323   0.7991
   0.750   0.0800   0.00568   0.00095   0.0011   0.7153   0.8125
   1.000   0.1067   0.00568   0.00097   0.0014   0.6997   0.8276
   1.250   0.1333   0.00569   0.00101   0.0018   0.6834   0.8438
   1.500   0.1591   0.00577   0.00103   0.0024   0.6494   0.8603
   1.750   0.1840   0.00589   0.00106   0.0032   0.5976   0.8792
   2.000   0.2080   0.00618   0.00112   0.0041   0.5092   0.9031
   2.250   0.2337   0.00656   0.00130   0.0045   0.4224   0.9291
   2.500   0.2621   0.00727   0.00150   0.0039   0.2767   0.9514
   2.750   0.2944   0.00787   0.00171   0.0025   0.1744   0.9681
   3.000   0.3272   0.00852   0.00196   0.0009   0.0779   0.9813
   3.250   0.3599   0.00908   0.00222  -0.0006   0.0221   0.9938
   3.500   0.3902   0.00930   0.00242  -0.0013   0.0157   1.0000
   3.750   0.4155   0.00956   0.00268  -0.0009   0.0101   1.0000
   4.000   0.4407   0.00993   0.00310  -0.0004   0.0037   1.0000
   4.250   0.4664   0.01026   0.00350   0.0000   0.0033   1.0000
   4.500   0.4921   0.01067   0.00402   0.0005   0.0032   1.0000
   4.750   0.5173   0.01122   0.00470   0.0010   0.0031   1.0000
   5.000   0.5421   0.01196   0.00558   0.0016   0.0030   1.0000
   5.250   0.5662   0.01290   0.00668   0.0023   0.0030   1.0000
   5.500   0.5898   0.01410   0.00806   0.0032   0.0030   1.0000
   5.750   0.6130   0.01562   0.00980   0.0040   0.0031   1.0000
   6.000   0.6359   0.01753   0.01201   0.0050   0.0031   1.0000
   6.250   0.6579   0.02001   0.01486   0.0061   0.0032   1.0000
   6.500   0.6781   0.02307   0.01836   0.0072   0.0032   1.0000
   6.750   0.6951   0.02702   0.02281   0.0085   0.0033   1.0000
   7.000   0.7078   0.03203   0.02841   0.0099   0.0035   1.0000
   7.250   0.7161   0.03780   0.03463   0.0110   0.0036   1.0000
   7.500   0.7206   0.04378   0.04098   0.0117   0.0037   1.0000
   7.750   0.7202   0.04999   0.04747   0.0118   0.0038   1.0000
   8.000   0.7142   0.05630   0.05400   0.0110   0.0039   1.0000
   8.250   0.7114   0.06061   0.05844   0.0100   0.0041   1.0000
   8.500   0.7053   0.06537   0.06330   0.0085   0.0041   1.0000
   8.750   0.6913   0.07170   0.06961   0.0035   0.0041   1.0000
   9.000   0.6834   0.07892   0.07680  -0.0031   0.0041   1.0000
   9.250   0.6812   0.08423   0.08208  -0.0066   0.0041   1.0000
   9.500   0.6805   0.08914   0.08696  -0.0095   0.0042   1.0000
   9.750   0.6811   0.09387   0.09168  -0.0120   0.0043   1.0000
  10.000   0.6797   0.09949   0.09727  -0.0147   0.0046   1.0000
  10.250   0.6693   0.10827   0.10601  -0.0187   0.0055   1.0000
  10.500   0.6674   0.11392   0.11164  -0.0213   0.0060   1.0000
  10.750   0.6660   0.11952   0.11723  -0.0237   0.0070   1.0000
  11.000   0.5803   0.12105   0.11896  -0.0237   0.0058   1.0000
  11.250   0.5781   0.12649   0.12440  -0.0258   0.0064   1.0000
<< Back to HQ 0/7 AIRFOIL (hq07-il)

Polar data table (+)

Polar graphs


<< Back to HQ 0/7 AIRFOIL (hq07-il)