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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 200,000
Max Cl/Cd: 62.38 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq108-il-200000.txt
Download as CSV file: xf-hq108-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5547   0.08797   0.08458  -0.0149   1.0000   0.0430
  -8.500  -0.5610   0.08251   0.07917  -0.0212   1.0000   0.0436
  -8.250  -0.5723   0.07707   0.07374  -0.0272   1.0000   0.0438
  -7.250  -0.5802   0.05697   0.05323  -0.0350   1.0000   0.0464
  -7.000  -0.5698   0.05368   0.04993  -0.0348   1.0000   0.0478
  -6.750  -0.5595   0.05031   0.04644  -0.0349   1.0000   0.0497
  -6.500  -0.5481   0.04676   0.04268  -0.0350   1.0000   0.0526
  -6.250  -0.5397   0.04272   0.03798  -0.0349   1.0000   0.0587
  -6.000  -0.5242   0.03916   0.03445  -0.0344   1.0000   0.0605
  -5.750  -0.5033   0.03130   0.02566  -0.0313   1.0000   0.0310
  -5.500  -0.4879   0.02684   0.02078  -0.0298   1.0000   0.0277
  -5.250  -0.4685   0.02306   0.01642  -0.0278   1.0000   0.0254
  -5.000  -0.4474   0.02050   0.01347  -0.0261   1.0000   0.0250
  -4.750  -0.4265   0.01896   0.01172  -0.0247   1.0000   0.0267
  -4.500  -0.4054   0.01812   0.01071  -0.0233   1.0000   0.0307
  -4.250  -0.3846   0.01678   0.00920  -0.0216   1.0000   0.0319
  -4.000  -0.3660   0.01488   0.00722  -0.0199   1.0000   0.0348
  -3.750  -0.3470   0.01394   0.00627  -0.0183   1.0000   0.0408
  -3.500  -0.3277   0.01298   0.00533  -0.0169   1.0000   0.0597
  -3.250  -0.3080   0.01202   0.00454  -0.0158   1.0000   0.1010
  -3.000  -0.2896   0.01047   0.00401  -0.0152   1.0000   0.2953
  -2.750  -0.2749   0.00937   0.00406  -0.0133   1.0000   0.5741
  -2.500  -0.2524   0.00924   0.00420  -0.0122   0.9981   0.6741
  -2.250  -0.2156   0.00927   0.00423  -0.0139   0.9915   0.7297
  -2.000  -0.1793   0.00930   0.00429  -0.0154   0.9846   0.7745
  -1.750  -0.1443   0.00930   0.00431  -0.0165   0.9777   0.8106
  -1.500  -0.1093   0.00927   0.00431  -0.0176   0.9708   0.8467
  -1.250  -0.0769   0.00922   0.00428  -0.0179   0.9637   0.8796
  -1.000  -0.0386   0.00918   0.00423  -0.0196   0.9577   0.9044
  -0.750   0.0059   0.00916   0.00419  -0.0226   0.9538   0.9297
  -0.500   0.0522   0.00915   0.00414  -0.0261   0.9481   0.9507
  -0.250   0.1034   0.00909   0.00403  -0.0308   0.9440   0.9622
   0.000   0.1530   0.00900   0.00391  -0.0353   0.9382   0.9720
   0.250   0.2018   0.00883   0.00373  -0.0394   0.9292   0.9806
   0.500   0.2489   0.00864   0.00353  -0.0432   0.9171   0.9889
   0.750   0.2933   0.00847   0.00334  -0.0465   0.9035   0.9998
   1.000   0.3121   0.00839   0.00322  -0.0448   0.8868   1.0000
   1.250   0.3293   0.00840   0.00320  -0.0428   0.8702   1.0000
   1.500   0.3492   0.00844   0.00323  -0.0412   0.8550   1.0000
   1.750   0.3708   0.00849   0.00325  -0.0398   0.8398   1.0000
   2.000   0.3935   0.00855   0.00329  -0.0385   0.8242   1.0000
   2.250   0.4169   0.00860   0.00333  -0.0373   0.8077   1.0000
   2.500   0.4405   0.00865   0.00336  -0.0361   0.7891   1.0000
   2.750   0.4643   0.00869   0.00343  -0.0349   0.7671   1.0000
   3.000   0.4881   0.00873   0.00346  -0.0337   0.7414   1.0000
   3.250   0.5120   0.00880   0.00348  -0.0324   0.7108   1.0000
   3.500   0.5360   0.00890   0.00355  -0.0313   0.6750   1.0000
   3.750   0.5598   0.00907   0.00364  -0.0301   0.6314   1.0000
   4.000   0.5826   0.00934   0.00378  -0.0288   0.5683   1.0000
   4.250   0.6036   0.00987   0.00396  -0.0273   0.4770   1.0000
   4.500   0.6235   0.01066   0.00430  -0.0260   0.3750   1.0000
   4.750   0.6439   0.01156   0.00477  -0.0250   0.2841   1.0000
   5.000   0.6650   0.01247   0.00534  -0.0242   0.2046   1.0000
   5.250   0.6848   0.01369   0.00604  -0.0233   0.1147   1.0000
   5.500   0.7049   0.01504   0.00717  -0.0222   0.0665   1.0000
   5.750   0.7255   0.01634   0.00837  -0.0211   0.0402   1.0000
   6.000   0.7458   0.01788   0.00998  -0.0197   0.0308   1.0000
   6.250   0.7640   0.02055   0.01266  -0.0181   0.0264   1.0000
   6.500   0.7886   0.02145   0.01374  -0.0172   0.0230   1.0000
   6.750   0.8117   0.02297   0.01542  -0.0164   0.0204   1.0000
   7.000   0.8341   0.02511   0.01778  -0.0153   0.0192   1.0000
   7.250   0.8555   0.02779   0.02079  -0.0141   0.0189   1.0000
   7.500   0.8746   0.03131   0.02481  -0.0125   0.0195   1.0000
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