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HQ 2.0/9 AIRFOIL (hq209-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.0/9 AIRFOIL (hq209-il)
Reynolds number: 50,000
Max Cl/Cd: 36.9 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq209-il-50000.txt
Download as CSV file: xf-hq209-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/9 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4447   0.10784   0.10099  -0.0129   1.0000   0.2364
  -8.750  -0.4624   0.10705   0.10034  -0.0148   1.0000   0.2463
  -8.500  -0.4529   0.10324   0.09656  -0.0141   1.0000   0.2599
  -8.250  -0.4501   0.10006   0.09346  -0.0136   1.0000   0.2738
  -8.000  -0.4513   0.09734   0.09083  -0.0132   1.0000   0.2881
  -7.750  -0.4273   0.09235   0.08582  -0.0113   1.0000   0.3056
  -7.500  -0.4193   0.08919   0.08271  -0.0102   1.0000   0.3236
  -7.250  -0.4336   0.08788   0.08155  -0.0091   1.0000   0.3433
  -7.000  -0.4225   0.08469   0.07841  -0.0067   1.0000   0.3706
  -6.750  -0.4001   0.08076   0.07448  -0.0040   1.0000   0.4038
  -6.500  -0.3908   0.07787   0.07164  -0.0017   1.0000   0.4325
  -5.750  -0.3850   0.07028   0.06432   0.0055   1.0000   0.4985
  -5.500  -0.4520   0.04941   0.04217  -0.0425   1.0000   0.1338
  -5.250  -0.4357   0.04472   0.03696  -0.0427   1.0000   0.1194
  -5.000  -0.4192   0.04110   0.03301  -0.0422   1.0000   0.1164
  -4.750  -0.4006   0.03790   0.02934  -0.0416   1.0000   0.1160
  -4.500  -0.3797   0.03489   0.02582  -0.0409   1.0000   0.1161
  -4.250  -0.3566   0.03205   0.02246  -0.0401   1.0000   0.1157
  -4.000  -0.3325   0.02961   0.01953  -0.0390   1.0000   0.1183
  -3.750  -0.3099   0.02764   0.01734  -0.0380   1.0000   0.1309
  -3.500  -0.2859   0.02570   0.01522  -0.0367   1.0000   0.1446
  -3.250  -0.2633   0.02403   0.01358  -0.0354   1.0000   0.1725
  -3.000  -0.2404   0.02232   0.01197  -0.0339   1.0000   0.2111
  -2.750  -0.2162   0.01973   0.01030  -0.0333   1.0000   0.3294
  -2.500  -0.2303   0.01827   0.01115  -0.0206   1.0000   0.7763
  -2.250  -0.1254   0.01847   0.01061  -0.0271   1.0000   1.0000
  -2.000  -0.1360   0.01806   0.01010  -0.0229   1.0000   1.0000
  -1.750  -0.1333   0.01781   0.00962  -0.0206   1.0000   1.0000
  -1.500  -0.1186   0.01776   0.00921  -0.0201   1.0000   1.0000
  -1.250  -0.0993   0.01783   0.00897  -0.0201   1.0000   1.0000
  -1.000  -0.0783   0.01798   0.00882  -0.0202   1.0000   1.0000
  -0.750  -0.0568   0.01819   0.00877  -0.0203   1.0000   1.0000
  -0.500  -0.0354   0.01845   0.00879  -0.0204   1.0000   1.0000
  -0.250  -0.0139   0.01876   0.00886  -0.0205   1.0000   1.0000
   0.000   0.0073   0.01911   0.00902  -0.0205   1.0000   1.0000
   0.250   0.0283   0.01950   0.00925  -0.0205   1.0000   1.0000
   0.500   0.0490   0.01994   0.00955  -0.0205   1.0000   1.0000
   0.750   0.0694   0.02041   0.00988  -0.0205   1.0000   1.0000
   1.000   0.0896   0.02093   0.01030  -0.0205   1.0000   1.0000
   1.250   0.1094   0.02150   0.01078  -0.0206   1.0000   1.0000
   1.500   0.1288   0.02211   0.01133  -0.0206   1.0000   1.0000
   1.750   0.1479   0.02278   0.01195  -0.0207   1.0000   1.0000
   2.000   0.1666   0.02350   0.01264  -0.0208   1.0000   1.0000
   2.250   0.1891   0.02436   0.01350  -0.0217   0.9978   1.0000
   2.500   0.2511   0.02599   0.01518  -0.0299   0.9748   1.0000
   2.750   0.3126   0.02732   0.01662  -0.0374   0.9489   1.0000
   3.000   0.3701   0.02835   0.01777  -0.0435   0.9243   1.0000
   3.250   0.4250   0.02914   0.01873  -0.0488   0.9007   1.0000
   3.500   0.4706   0.02972   0.01954  -0.0522   0.8758   1.0000
   3.750   0.5194   0.03010   0.02015  -0.0555   0.8501   1.0000
   4.000   0.5712   0.03016   0.02049  -0.0587   0.8232   1.0000
   4.250   0.6272   0.02973   0.02046  -0.0616   0.7950   1.0000
   4.500   0.6799   0.02882   0.01989  -0.0628   0.7653   1.0000
   4.750   0.7325   0.02733   0.01874  -0.0628   0.7341   1.0000
   5.000   0.7689   0.02620   0.01793  -0.0606   0.6953   1.0000
   5.250   0.8059   0.02483   0.01672  -0.0579   0.6521   1.0000
   5.500   0.8330   0.02407   0.01597  -0.0544   0.6004   1.0000
   5.750   0.8572   0.02362   0.01538  -0.0509   0.5437   1.0000
   6.000   0.8768   0.02376   0.01533  -0.0475   0.4827   1.0000
   6.250   0.8937   0.02448   0.01571  -0.0444   0.4188   1.0000
   6.500   0.9080   0.02577   0.01657  -0.0414   0.3492   1.0000
   6.750   0.9219   0.02775   0.01796  -0.0387   0.2781   1.0000
   7.000   0.9428   0.03033   0.02024  -0.0371   0.2256   1.0000
   7.250   0.9634   0.03242   0.02224  -0.0357   0.1928   1.0000
   7.500   0.9871   0.03468   0.02455  -0.0348   0.1713   1.0000
   7.750   1.0119   0.03737   0.02738  -0.0339   0.1568   1.0000
   8.000   1.0318   0.03981   0.03000  -0.0328   0.1415   1.0000
   8.250   1.0495   0.04311   0.03388  -0.0312   0.1325   1.0000
   8.500   1.0649   0.04730   0.03850  -0.0296   0.1265   1.0000
   8.750   1.0782   0.05058   0.04193  -0.0281   0.1148   1.0000
   9.000   1.0766   0.05513   0.04726  -0.0254   0.1110   1.0000
   9.250   1.0870   0.05883   0.05098  -0.0240   0.1016   1.0000
   9.500   1.0768   0.06341   0.05615  -0.0216   0.0998   1.0000
   9.750   1.0628   0.06821   0.06138  -0.0196   0.0989   1.0000
  10.000   1.0470   0.07308   0.06654  -0.0182   0.0994   1.0000
  10.250   1.0287   0.07782   0.07147  -0.0170   0.1003   1.0000
  10.500   1.0107   0.08268   0.07643  -0.0164   0.1012   1.0000
  10.750   0.9939   0.08807   0.08190  -0.0170   0.1020   1.0000
  11.000   0.9813   0.09401   0.08791  -0.0183   0.1028   1.0000
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