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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 100,000
Max Cl/Cd: 44.7 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq108-il-100000-n5.txt
Download as CSV file: xf-hq108-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5583   0.09043   0.08558  -0.0111   1.0000   0.0194
  -8.750  -0.5612   0.08521   0.08042  -0.0143   1.0000   0.0189
  -8.500  -0.5658   0.07958   0.07487  -0.0182   1.0000   0.0184
  -8.250  -0.5743   0.07306   0.06842  -0.0241   1.0000   0.0179
  -8.000  -0.5827   0.06696   0.06227  -0.0286   1.0000   0.0175
  -7.750  -0.5891   0.06032   0.05547  -0.0319   1.0000   0.0168
  -7.500  -0.5901   0.05437   0.04925  -0.0336   1.0000   0.0162
  -7.250  -0.5863   0.04927   0.04382  -0.0342   1.0000   0.0159
  -7.000  -0.5782   0.04493   0.03911  -0.0341   1.0000   0.0157
  -6.750  -0.5680   0.04076   0.03452  -0.0337   1.0000   0.0157
  -6.500  -0.5551   0.03691   0.03024  -0.0331   1.0000   0.0157
  -6.250  -0.5401   0.03335   0.02622  -0.0322   1.0000   0.0160
  -6.000  -0.5229   0.03088   0.02346  -0.0315   1.0000   0.0173
  -5.750  -0.5033   0.02914   0.02145  -0.0308   1.0000   0.0196
  -5.500  -0.4827   0.02689   0.01879  -0.0296   1.0000   0.0216
  -5.250  -0.4611   0.02448   0.01596  -0.0282   1.0000   0.0227
  -5.000  -0.4391   0.02246   0.01363  -0.0269   1.0000   0.0239
  -4.750  -0.4175   0.02081   0.01178  -0.0255   1.0000   0.0254
  -4.500  -0.3980   0.01919   0.01013  -0.0242   1.0000   0.0293
  -4.250  -0.3767   0.01838   0.00910  -0.0230   1.0000   0.0369
  -4.000  -0.3573   0.01714   0.00786  -0.0216   1.0000   0.0449
  -3.750  -0.3370   0.01628   0.00695  -0.0204   1.0000   0.0623
  -3.500  -0.3172   0.01540   0.00623  -0.0193   1.0000   0.0934
  -3.250  -0.2974   0.01448   0.00559  -0.0183   1.0000   0.1503
  -3.000  -0.2790   0.01328   0.00512  -0.0174   1.0000   0.3065
  -2.750  -0.2624   0.01231   0.00505  -0.0156   1.0000   0.5194
  -2.500  -0.2350   0.01206   0.00510  -0.0153   0.9942   0.6407
  -2.250  -0.2033   0.01200   0.00505  -0.0157   0.9862   0.7123
  -2.000  -0.1733   0.01197   0.00506  -0.0156   0.9779   0.7705
  -1.750  -0.1416   0.01194   0.00504  -0.0158   0.9707   0.8143
  -1.500  -0.1094   0.01190   0.00494  -0.0162   0.9628   0.8463
  -1.250  -0.0745   0.01186   0.00483  -0.0172   0.9563   0.8799
  -1.000  -0.0354   0.01183   0.00473  -0.0192   0.9497   0.9063
  -0.750   0.0089   0.01180   0.00462  -0.0226   0.9444   0.9241
  -0.500   0.0520   0.01176   0.00452  -0.0258   0.9369   0.9427
  -0.250   0.0983   0.01172   0.00439  -0.0297   0.9305   0.9599
   0.000   0.1437   0.01165   0.00429  -0.0335   0.9219   0.9767
   0.250   0.1885   0.01159   0.00420  -0.0372   0.9129   0.9941
   0.500   0.2229   0.01157   0.00415  -0.0388   0.9017   1.0000
   0.750   0.2518   0.01158   0.00413  -0.0392   0.8884   1.0000
   1.000   0.2789   0.01159   0.00411  -0.0390   0.8725   1.0000
   1.250   0.3054   0.01158   0.00408  -0.0384   0.8541   1.0000
   1.500   0.3303   0.01159   0.00409  -0.0375   0.8338   1.0000
   1.750   0.3549   0.01164   0.00412  -0.0366   0.8145   1.0000
   2.000   0.3802   0.01170   0.00419  -0.0358   0.7978   1.0000
   2.250   0.4051   0.01178   0.00430  -0.0350   0.7801   1.0000
   2.500   0.4301   0.01187   0.00442  -0.0342   0.7611   1.0000
   2.750   0.4552   0.01195   0.00460  -0.0333   0.7409   1.0000
   3.000   0.4799   0.01205   0.00473  -0.0323   0.7166   1.0000
   3.250   0.5042   0.01215   0.00486  -0.0312   0.6868   1.0000
   3.500   0.5281   0.01229   0.00498  -0.0300   0.6483   1.0000
   3.750   0.5514   0.01249   0.00510  -0.0286   0.5956   1.0000
   4.000   0.5735   0.01283   0.00531  -0.0271   0.5222   1.0000
   4.250   0.5944   0.01343   0.00555  -0.0256   0.4350   1.0000
   4.500   0.6150   0.01421   0.00598  -0.0243   0.3492   1.0000
   4.750   0.6358   0.01508   0.00652  -0.0234   0.2742   1.0000
   5.000   0.6573   0.01595   0.00717  -0.0226   0.2134   1.0000
   5.250   0.6789   0.01688   0.00786  -0.0219   0.1484   1.0000
   5.500   0.6997   0.01802   0.00872  -0.0210   0.0971   1.0000
   5.750   0.7203   0.01926   0.00979  -0.0201   0.0657   1.0000
   6.000   0.7410   0.02060   0.01126  -0.0190   0.0466   1.0000
   6.250   0.7616   0.02192   0.01259  -0.0181   0.0321   1.0000
   6.500   0.7821   0.02333   0.01414  -0.0170   0.0242   1.0000
   6.750   0.8013   0.02508   0.01601  -0.0158   0.0208   1.0000
   7.000   0.8202   0.02758   0.01869  -0.0144   0.0191   1.0000
   7.250   0.8417   0.02998   0.02141  -0.0132   0.0178   1.0000
   7.500   0.8622   0.03219   0.02398  -0.0122   0.0157   1.0000
   7.750   0.8807   0.03441   0.02655  -0.0111   0.0141   1.0000
   8.000   0.8965   0.03662   0.02906  -0.0101   0.0127   1.0000
   8.250   0.9090   0.03979   0.03264  -0.0088   0.0123   1.0000
   8.500   0.9174   0.04335   0.03664  -0.0073   0.0121   1.0000
   8.750   0.9211   0.04715   0.04087  -0.0058   0.0119   1.0000
   9.000   0.9200   0.05115   0.04526  -0.0043   0.0118   1.0000
   9.250   0.9142   0.05516   0.04962  -0.0029   0.0118   1.0000
   9.500   0.9025   0.05903   0.05377  -0.0015   0.0119   1.0000
   9.750   0.8859   0.06307   0.05804  -0.0008   0.0120   1.0000
  10.000   0.8666   0.06814   0.06331  -0.0023   0.0122   1.0000
  10.250   0.8455   0.07481   0.07014  -0.0066   0.0124   1.0000
  10.500   0.8220   0.08475   0.08020  -0.0143   0.0128   1.0000
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