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HQ 2.0/8 AIRFOIL (hq208-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.0/8 AIRFOIL (hq208-il)
Reynolds number: 500,000
Max Cl/Cd: 98.36 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq208-il-500000.txt
Download as CSV file: xf-hq208-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4587   0.08406   0.08186  -0.0262   1.0000   0.0127
  -8.250  -0.4579   0.08067   0.07849  -0.0275   1.0000   0.0130
  -8.000  -0.4591   0.07718   0.07504  -0.0291   1.0000   0.0133
  -7.750  -0.4645   0.07359   0.07150  -0.0307   1.0000   0.0133
  -7.500  -0.4675   0.06884   0.06679  -0.0354   1.0000   0.0133
  -7.250  -0.4674   0.06368   0.06161  -0.0402   1.0000   0.0134
  -7.000  -0.4661   0.05942   0.05730  -0.0424   1.0000   0.0136
  -6.750  -0.4660   0.05558   0.05340  -0.0431   1.0000   0.0139
  -6.500  -0.4691   0.05243   0.05019  -0.0418   1.0000   0.0141
  -6.250  -0.4514   0.04775   0.04535  -0.0451   0.9978   0.0147
  -6.000  -0.4197   0.04223   0.03957  -0.0505   0.9942   0.0159
  -5.750  -0.3875   0.03746   0.03449  -0.0539   0.9900   0.0177
  -5.500  -0.3490   0.03581   0.03249  -0.0557   0.9866   0.0194
  -5.250  -0.3160   0.03182   0.02812  -0.0587   0.9836   0.0195
  -4.750  -0.2570   0.01754   0.01247  -0.0617   0.9753   0.0118
  -4.500  -0.2234   0.01458   0.00910  -0.0629   0.9726   0.0121
  -4.250  -0.1894   0.01277   0.00710  -0.0643   0.9700   0.0133
  -4.000  -0.1570   0.01223   0.00653  -0.0654   0.9656   0.0156
  -3.750  -0.1267   0.01103   0.00519  -0.0658   0.9599   0.0157
  -3.500  -0.0946   0.01011   0.00417  -0.0665   0.9555   0.0167
  -3.250  -0.0668   0.00955   0.00351  -0.0664   0.9478   0.0190
  -3.000  -0.0374   0.00879   0.00267  -0.0665   0.9418   0.0356
  -2.750  -0.0112   0.00831   0.00238  -0.0662   0.9332   0.0865
  -2.500   0.0152   0.00765   0.00212  -0.0662   0.9261   0.2055
  -2.250   0.0372   0.00642   0.00194  -0.0658   0.9172   0.5166
  -2.000   0.0624   0.00615   0.00192  -0.0651   0.9088   0.6147
  -1.750   0.0889   0.00607   0.00184  -0.0645   0.9010   0.6538
  -1.500   0.1151   0.00601   0.00180  -0.0640   0.8916   0.6820
  -1.250   0.1412   0.00597   0.00175  -0.0634   0.8819   0.7081
  -1.000   0.1669   0.00593   0.00172  -0.0626   0.8713   0.7368
  -0.750   0.1921   0.00588   0.00170  -0.0617   0.8605   0.7683
  -0.500   0.2174   0.00583   0.00169  -0.0609   0.8498   0.7953
  -0.250   0.2432   0.00579   0.00166  -0.0602   0.8395   0.8171
   0.000   0.2689   0.00574   0.00163  -0.0595   0.8289   0.8359
   0.250   0.2948   0.00569   0.00158  -0.0589   0.8180   0.8524
   0.500   0.3206   0.00563   0.00155  -0.0582   0.8071   0.8717
   0.750   0.3458   0.00554   0.00152  -0.0574   0.7960   0.8973
   1.000   0.3743   0.00546   0.00150  -0.0573   0.7850   0.9376
   1.250   0.4168   0.00542   0.00145  -0.0604   0.7731   0.9982
   1.500   0.4444   0.00548   0.00146  -0.0604   0.7598   1.0000
   1.750   0.4714   0.00555   0.00148  -0.0602   0.7458   1.0000
   2.000   0.4984   0.00563   0.00151  -0.0600   0.7299   1.0000
   2.250   0.5252   0.00573   0.00155  -0.0598   0.7128   1.0000
   2.500   0.5518   0.00583   0.00160  -0.0595   0.6928   1.0000
   2.750   0.5781   0.00597   0.00169  -0.0591   0.6697   1.0000
   3.000   0.6039   0.00614   0.00177  -0.0586   0.6380   1.0000
   3.250   0.6281   0.00644   0.00185  -0.0579   0.5825   1.0000
   3.500   0.6512   0.00688   0.00199  -0.0570   0.5116   1.0000
   3.750   0.6749   0.00734   0.00220  -0.0563   0.4479   1.0000
   4.000   0.6982   0.00788   0.00248  -0.0557   0.3811   1.0000
   4.250   0.7213   0.00848   0.00278  -0.0551   0.3124   1.0000
   4.500   0.7457   0.00894   0.00307  -0.0547   0.2715   1.0000
   4.750   0.7705   0.00937   0.00337  -0.0543   0.2384   1.0000
   5.000   0.7947   0.00986   0.00367  -0.0539   0.1949   1.0000
   5.250   0.8169   0.01062   0.00409  -0.0532   0.1214   1.0000
   5.500   0.8389   0.01143   0.00460  -0.0525   0.0691   1.0000
   5.750   0.8629   0.01197   0.00505  -0.0521   0.0494   1.0000
   6.000   0.8875   0.01242   0.00548  -0.0516   0.0394   1.0000
   6.250   0.9123   0.01283   0.00590  -0.0512   0.0318   1.0000
   6.500   0.9359   0.01343   0.00647  -0.0506   0.0177   1.0000
   6.750   0.9578   0.01436   0.00744  -0.0495   0.0105   1.0000
   7.000   0.9815   0.01493   0.00812  -0.0489   0.0087   1.0000
   7.250   1.0041   0.01564   0.00888  -0.0481   0.0074   1.0000
   7.500   1.0243   0.01673   0.01011  -0.0469   0.0067   1.0000
   7.750   1.0395   0.01861   0.01221  -0.0449   0.0061   1.0000
   8.000   1.0592   0.01979   0.01355  -0.0436   0.0058   1.0000
   8.250   1.0772   0.02134   0.01529  -0.0421   0.0056   1.0000
   8.500   1.0940   0.02326   0.01744  -0.0405   0.0054   1.0000
   8.750   1.1091   0.02572   0.02019  -0.0386   0.0054   1.0000
   9.000   1.1216   0.02870   0.02351  -0.0366   0.0055   1.0000
   9.250   1.1291   0.03234   0.02756  -0.0341   0.0056   1.0000
   9.500   1.1289   0.03668   0.03234  -0.0312   0.0058   1.0000
   9.750   1.1209   0.04179   0.03782  -0.0279   0.0062   1.0000
  10.000   1.1295   0.04321   0.03938  -0.0260   0.0064   1.0000
  10.250   1.0780   0.05315   0.05014  -0.0186   0.0087   1.0000
  10.500   1.0561   0.05717   0.05437  -0.0165   0.0089   1.0000
  10.750   1.0311   0.06215   0.05954  -0.0163   0.0092   1.0000
  11.000   1.0130   0.06693   0.06447  -0.0177   0.0092   1.0000
  11.250   0.9931   0.07294   0.07062  -0.0209   0.0091   1.0000
  11.500   0.9752   0.07993   0.07775  -0.0259   0.0089   1.0000
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