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HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il)
Reynolds number: 100,000
Max Cl/Cd: 51.93 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq158-il-100000.txt
Download as CSV file: xf-hq158-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4999   0.09357   0.08881  -0.0167   1.0000   0.0889
  -8.250  -0.5107   0.09000   0.08535  -0.0216   1.0000   0.0920
  -8.000  -0.5225   0.08591   0.08136  -0.0273   1.0000   0.0926
  -7.750  -0.5065   0.08242   0.07787  -0.0224   1.0000   0.0984
  -7.500  -0.5096   0.07847   0.07399  -0.0260   1.0000   0.1032
  -7.250  -0.5240   0.07361   0.06900  -0.0359   1.0000   0.1063
  -7.000  -0.5086   0.06984   0.06539  -0.0315   1.0000   0.1123
  -6.750  -0.5217   0.06679   0.06185  -0.0397   1.0000   0.1205
  -6.500  -0.5017   0.06187   0.05734  -0.0351   1.0000   0.1257
  -6.250  -0.4990   0.05793   0.05326  -0.0372   1.0000   0.1364
  -5.750  -0.4827   0.05151   0.04668  -0.0368   1.0000   0.1632
  -5.500  -0.4721   0.04851   0.04363  -0.0355   1.0000   0.1780
  -5.250  -0.4609   0.04582   0.04095  -0.0340   1.0000   0.1947
  -5.000  -0.4213   0.03357   0.02666  -0.0370   1.0000   0.0637
  -4.750  -0.3995   0.03044   0.02260  -0.0348   1.0000   0.0539
  -4.500  -0.3800   0.02728   0.01911  -0.0334   1.0000   0.0524
  -4.250  -0.3590   0.02488   0.01631  -0.0319   1.0000   0.0524
  -4.000  -0.3385   0.02265   0.01385  -0.0307   1.0000   0.0569
  -3.750  -0.3167   0.02111   0.01218  -0.0294   1.0000   0.0619
  -3.500  -0.2935   0.01963   0.01043  -0.0278   1.0000   0.0662
  -3.250  -0.2722   0.01802   0.00890  -0.0266   1.0000   0.0779
  -3.000  -0.2518   0.01660   0.00765  -0.0252   1.0000   0.1042
  -2.750  -0.2316   0.01507   0.00655  -0.0241   1.0000   0.1713
  -2.500  -0.2255   0.01254   0.00656  -0.0195   1.0000   0.6879
  -2.250  -0.2160   0.01251   0.00662  -0.0146   1.0000   0.7865
  -2.000  -0.2051   0.01239   0.00654  -0.0101   1.0000   0.8550
  -1.750  -0.1619   0.01226   0.00637  -0.0117   1.0000   0.9494
  -1.500  -0.1172   0.01222   0.00604  -0.0162   1.0000   1.0000
  -1.250  -0.1001   0.01226   0.00586  -0.0157   1.0000   1.0000
  -1.000  -0.0799   0.01239   0.00574  -0.0155   1.0000   1.0000
  -0.750  -0.0589   0.01257   0.00574  -0.0154   1.0000   1.0000
  -0.500  -0.0377   0.01279   0.00580  -0.0152   1.0000   1.0000
  -0.250  -0.0166   0.01306   0.00592  -0.0150   1.0000   1.0000
   0.000   0.0300   0.01351   0.00621  -0.0196   0.9910   1.0000
   0.250   0.0767   0.01398   0.00657  -0.0242   0.9825   1.0000
   0.500   0.1197   0.01432   0.00684  -0.0279   0.9724   1.0000
   0.750   0.1640   0.01465   0.00713  -0.0317   0.9622   1.0000
   1.000   0.2106   0.01494   0.00739  -0.0359   0.9521   1.0000
   1.250   0.2598   0.01512   0.00759  -0.0403   0.9416   1.0000
   1.500   0.3064   0.01518   0.00770  -0.0441   0.9289   1.0000
   1.750   0.3544   0.01514   0.00776  -0.0478   0.9163   1.0000
   2.000   0.3997   0.01507   0.00778  -0.0509   0.9038   1.0000
   2.250   0.4429   0.01491   0.00772  -0.0533   0.8900   1.0000
   2.500   0.4815   0.01475   0.00768  -0.0546   0.8756   1.0000
   2.750   0.5170   0.01456   0.00766  -0.0551   0.8604   1.0000
   3.000   0.5453   0.01446   0.00767  -0.0542   0.8415   1.0000
   3.250   0.5737   0.01425   0.00757  -0.0530   0.8213   1.0000
   3.500   0.5991   0.01406   0.00749  -0.0512   0.7973   1.0000
   3.750   0.6237   0.01382   0.00738  -0.0491   0.7690   1.0000
   4.000   0.6472   0.01357   0.00718  -0.0466   0.7338   1.0000
   4.250   0.6692   0.01340   0.00703  -0.0441   0.6868   1.0000
   4.500   0.6906   0.01339   0.00692  -0.0415   0.6240   1.0000
   4.750   0.7104   0.01368   0.00694  -0.0390   0.5393   1.0000
   5.000   0.7285   0.01441   0.00721  -0.0367   0.4449   1.0000
   5.250   0.7466   0.01538   0.00784  -0.0349   0.3606   1.0000
   5.500   0.7632   0.01666   0.00868  -0.0331   0.2739   1.0000
   5.750   0.7776   0.01841   0.00974  -0.0312   0.1715   1.0000
   6.000   0.7951   0.02024   0.01120  -0.0295   0.1240   1.0000
   6.250   0.8148   0.02195   0.01277  -0.0281   0.0978   1.0000
   6.500   0.8365   0.02402   0.01475  -0.0269   0.0804   1.0000
   6.750   0.8570   0.02609   0.01684  -0.0258   0.0613   1.0000
   7.000   0.8800   0.02909   0.01992  -0.0247   0.0506   1.0000
   7.250   0.9029   0.03138   0.02263  -0.0235   0.0441   1.0000
   7.500   0.9214   0.03489   0.02635  -0.0225   0.0392   1.0000
   7.750   0.9394   0.03818   0.03020  -0.0208   0.0384   1.0000
   8.000   0.9531   0.04213   0.03470  -0.0190   0.0383   1.0000
   8.250   0.9627   0.04652   0.03957  -0.0171   0.0387   1.0000
   8.500   0.9686   0.05154   0.04497  -0.0156   0.0395   1.0000
   8.750   0.9787   0.05507   0.04892  -0.0138   0.0413   1.0000
   9.000   0.9546   0.06196   0.05674  -0.0107   0.0465   1.0000
   9.250   0.9439   0.06709   0.06212  -0.0096   0.0489   1.0000
   9.500   0.9259   0.07277   0.06810  -0.0085   0.0555   1.0000
   9.750   0.8953   0.07797   0.07346  -0.0092   0.0567   1.0000
  10.000   0.8690   0.08441   0.07998  -0.0130   0.0574   1.0000
  10.250   0.7768   0.08142   0.07724  -0.0111   0.0486   1.0000
  10.500   0.7555   0.08975   0.08561  -0.0160   0.0496   1.0000
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