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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 500,000
Max Cl/Cd: 78.58 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq108-il-500000.txt
Download as CSV file: xf-hq108-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5722   0.08099   0.07886  -0.0167   1.0000   0.0138
  -8.750  -0.5731   0.07740   0.07529  -0.0187   1.0000   0.0142
  -8.500  -0.5862   0.06801   0.06593  -0.0284   1.0000   0.0137
  -8.250  -0.5975   0.06270   0.06054  -0.0317   1.0000   0.0137
  -8.000  -0.6007   0.05858   0.05635  -0.0333   1.0000   0.0141
  -7.750  -0.5968   0.05599   0.05370  -0.0339   1.0000   0.0147
  -7.500  -0.5936   0.05163   0.04920  -0.0349   1.0000   0.0150
  -7.250  -0.5857   0.04827   0.04571  -0.0353   1.0000   0.0158
  -7.000  -0.5756   0.04478   0.04201  -0.0354   1.0000   0.0173
  -6.750  -0.5484   0.04439   0.04134  -0.0344   1.0000   0.0212
  -6.500  -0.5366   0.04117   0.03784  -0.0336   1.0000   0.0213
  -5.500  -0.4918   0.02110   0.01611  -0.0266   1.0000   0.0131
  -5.250  -0.4763   0.01762   0.01230  -0.0244   1.0000   0.0118
  -5.000  -0.4592   0.01555   0.00997  -0.0223   1.0000   0.0115
  -4.750  -0.4289   0.01375   0.00795  -0.0228   0.9979   0.0117
  -4.500  -0.3932   0.01262   0.00668  -0.0244   0.9946   0.0126
  -4.250  -0.3607   0.01090   0.00477  -0.0256   0.9903   0.0147
  -4.000  -0.3248   0.01037   0.00420  -0.0275   0.9861   0.0189
  -3.750  -0.2878   0.00955   0.00330  -0.0294   0.9828   0.0287
  -3.500  -0.2530   0.00894   0.00282  -0.0311   0.9779   0.0660
  -3.250  -0.2184   0.00837   0.00250  -0.0329   0.9727   0.1274
  -3.000  -0.1853   0.00737   0.00217  -0.0349   0.9676   0.2980
  -2.750  -0.1573   0.00648   0.00194  -0.0354   0.9583   0.4881
  -2.500  -0.1282   0.00612   0.00187  -0.0357   0.9491   0.5860
  -2.250  -0.0993   0.00595   0.00180  -0.0358   0.9395   0.6343
  -2.000  -0.0720   0.00586   0.00172  -0.0354   0.9282   0.6649
  -1.750  -0.0459   0.00580   0.00166  -0.0347   0.9160   0.6957
  -1.500  -0.0201   0.00575   0.00163  -0.0340   0.9041   0.7214
  -1.250   0.0058   0.00571   0.00161  -0.0333   0.8929   0.7445
  -1.000   0.0314   0.00569   0.00160  -0.0326   0.8815   0.7686
  -0.750   0.0572   0.00568   0.00157  -0.0319   0.8688   0.7853
  -0.500   0.0831   0.00566   0.00153  -0.0312   0.8557   0.8004
  -0.250   0.1089   0.00565   0.00151  -0.0305   0.8432   0.8177
   0.000   0.1345   0.00563   0.00149  -0.0298   0.8309   0.8355
   0.250   0.1600   0.00559   0.00147  -0.0290   0.8180   0.8529
   0.750   0.2117   0.00553   0.00145  -0.0278   0.7927   0.8862
   1.000   0.2375   0.00549   0.00145  -0.0271   0.7805   0.9059
   1.250   0.2650   0.00547   0.00146  -0.0267   0.7679   0.9302
   1.500   0.2986   0.00547   0.00147  -0.0278   0.7541   0.9565
   1.750   0.3373   0.00551   0.00149  -0.0301   0.7380   0.9779
   2.000   0.3768   0.00555   0.00150  -0.0327   0.7169   0.9968
   2.250   0.4039   0.00564   0.00152  -0.0326   0.6946   1.0000
   2.500   0.4285   0.00577   0.00155  -0.0319   0.6649   1.0000
   2.750   0.4535   0.00592   0.00163  -0.0313   0.6318   1.0000
   3.000   0.4786   0.00611   0.00171  -0.0308   0.5942   1.0000
   3.250   0.5029   0.00640   0.00181  -0.0301   0.5383   1.0000
   3.500   0.5263   0.00685   0.00197  -0.0293   0.4631   1.0000
   3.750   0.5496   0.00739   0.00220  -0.0287   0.3841   1.0000
   4.000   0.5731   0.00798   0.00249  -0.0282   0.3054   1.0000
   4.250   0.5969   0.00855   0.00280  -0.0277   0.2448   1.0000
   4.500   0.6216   0.00902   0.00309  -0.0273   0.1960   1.0000
   4.750   0.6456   0.00960   0.00342  -0.0269   0.1373   1.0000
   5.000   0.6689   0.01031   0.00384  -0.0264   0.0809   1.0000
   5.250   0.6930   0.01091   0.00429  -0.0259   0.0511   1.0000
   5.500   0.7166   0.01166   0.00496  -0.0253   0.0250   1.0000
   5.750   0.7400   0.01249   0.00581  -0.0245   0.0142   1.0000
   6.000   0.7634   0.01329   0.00668  -0.0237   0.0111   1.0000
   6.250   0.7821   0.01504   0.00863  -0.0221   0.0094   1.0000
   6.500   0.8049   0.01605   0.00976  -0.0211   0.0088   1.0000
   6.750   0.8264   0.01746   0.01133  -0.0200   0.0085   1.0000
   7.000   0.8474   0.01928   0.01335  -0.0188   0.0084   1.0000
   7.250   0.8679   0.02159   0.01591  -0.0175   0.0085   1.0000
   7.500   0.8863   0.02457   0.01924  -0.0160   0.0088   1.0000
   7.750   0.8986   0.02955   0.02468  -0.0141   0.0098   1.0000
  12.500   0.6523   0.14015   0.13819  -0.0421   0.0166   1.0000
  12.750   0.6490   0.14404   0.14209  -0.0443   0.0166   1.0000
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