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HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/9 AIRFOIL (hq159-il)
Reynolds number: 50,000
Max Cl/Cd: 35.14 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq159-il-50000.txt
Download as CSV file: xf-hq159-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/9 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4886   0.11068   0.10385  -0.0046   1.0000   0.2491
  -9.000  -0.5024   0.10934   0.10262  -0.0062   1.0000   0.2603
  -8.750  -0.4966   0.10571   0.09905  -0.0058   1.0000   0.2741
  -8.500  -0.4912   0.10229   0.09568  -0.0055   1.0000   0.2879
  -8.250  -0.4804   0.09815   0.09152  -0.0047   1.0000   0.3030
  -8.000  -0.4654   0.09402   0.08741  -0.0035   1.0000   0.3193
  -7.750  -0.4587   0.09102   0.08445  -0.0024   1.0000   0.3402
  -7.500  -0.4481   0.08767   0.08113  -0.0009   1.0000   0.3642
  -6.250  -0.5171   0.05375   0.04665  -0.0396   1.0000   0.1344
  -6.000  -0.5052   0.04850   0.04094  -0.0398   1.0000   0.1199
  -5.750  -0.4915   0.04406   0.03612  -0.0393   1.0000   0.1129
  -5.500  -0.4757   0.04051   0.03159  -0.0384   1.0000   0.1083
  -5.250  -0.4591   0.03732   0.02841  -0.0373   1.0000   0.1134
  -5.000  -0.4399   0.03440   0.02505  -0.0361   1.0000   0.1159
  -4.750  -0.4185   0.03158   0.02172  -0.0347   1.0000   0.1177
  -4.500  -0.3964   0.02908   0.01876  -0.0333   1.0000   0.1241
  -4.250  -0.3753   0.02707   0.01667  -0.0320   1.0000   0.1397
  -4.000  -0.3528   0.02512   0.01465  -0.0304   1.0000   0.1587
  -3.750  -0.3308   0.02336   0.01292  -0.0287   1.0000   0.1889
  -3.500  -0.3097   0.02154   0.01142  -0.0272   1.0000   0.2388
  -3.250  -0.2938   0.01850   0.00994  -0.0250   1.0000   0.4122
  -3.000  -0.3067   0.01805   0.01094  -0.0123   1.0000   0.7471
  -2.750  -0.0695   0.01906   0.01025  -0.0363   1.0000   1.0000
  -2.500  -0.0711   0.01870   0.00983  -0.0334   1.0000   1.0000
  -2.250  -0.0774   0.01840   0.00948  -0.0297   1.0000   1.0000
  -2.000  -0.0880   0.01811   0.00917  -0.0253   1.0000   1.0000
  -1.750  -0.1010   0.01781   0.00884  -0.0205   1.0000   1.0000
  -1.500  -0.1107   0.01753   0.00848  -0.0162   1.0000   1.0000
  -1.250  -0.1082   0.01740   0.00817  -0.0136   1.0000   1.0000
  -1.000  -0.0954   0.01741   0.00797  -0.0125   1.0000   1.0000
  -0.750  -0.0785   0.01752   0.00783  -0.0120   1.0000   1.0000
  -0.500  -0.0598   0.01770   0.00781  -0.0116   1.0000   1.0000
  -0.250  -0.0403   0.01794   0.00787  -0.0114   1.0000   1.0000
   0.000  -0.0205   0.01822   0.00800  -0.0112   1.0000   1.0000
   0.250  -0.0006   0.01855   0.00818  -0.0110   1.0000   1.0000
   0.500   0.0192   0.01891   0.00841  -0.0108   1.0000   1.0000
   0.750   0.0390   0.01932   0.00872  -0.0107   1.0000   1.0000
   1.000   0.0586   0.01977   0.00909  -0.0106   1.0000   1.0000
   1.250   0.0780   0.02027   0.00952  -0.0105   1.0000   1.0000
   1.500   0.0972   0.02081   0.01002  -0.0104   1.0000   1.0000
   1.750   0.1160   0.02141   0.01059  -0.0104   1.0000   1.0000
   2.000   0.1346   0.02206   0.01122  -0.0104   1.0000   1.0000
   2.250   0.1528   0.02276   0.01192  -0.0105   1.0000   1.0000
   2.500   0.2026   0.02413   0.01335  -0.0166   0.9845   1.0000
   2.750   0.2690   0.02555   0.01493  -0.0252   0.9579   1.0000
   3.000   0.3262   0.02658   0.01612  -0.0316   0.9324   1.0000
   3.250   0.3805   0.02740   0.01714  -0.0369   0.9075   1.0000
   3.500   0.4348   0.02801   0.01804  -0.0418   0.8817   1.0000
   3.750   0.4901   0.02831   0.01864  -0.0461   0.8541   1.0000
   4.000   0.5479   0.02820   0.01889  -0.0499   0.8241   1.0000
   4.250   0.6058   0.02752   0.01866  -0.0524   0.7924   1.0000
   4.500   0.6541   0.02652   0.01801  -0.0523   0.7571   1.0000
   4.750   0.6962   0.02515   0.01700  -0.0502   0.7169   1.0000
   5.000   0.7288   0.02385   0.01588  -0.0466   0.6685   1.0000
   5.250   0.7560   0.02279   0.01482  -0.0424   0.6113   1.0000
   5.500   0.7772   0.02236   0.01422  -0.0382   0.5439   1.0000
   5.750   0.7953   0.02263   0.01409  -0.0344   0.4696   1.0000
   6.000   0.8104   0.02371   0.01466  -0.0311   0.3895   1.0000
   6.250   0.8249   0.02553   0.01598  -0.0282   0.3089   1.0000
   6.500   0.8427   0.02780   0.01778  -0.0262   0.2448   1.0000
   6.750   0.8665   0.03022   0.01999  -0.0251   0.2059   1.0000
   7.000   0.8876   0.03220   0.02198  -0.0238   0.1771   1.0000
   7.250   0.9126   0.03486   0.02484  -0.0228   0.1598   1.0000
   7.500   0.9343   0.03722   0.02725  -0.0218   0.1428   1.0000
   7.750   0.9539   0.04034   0.03078  -0.0204   0.1313   1.0000
   8.000   0.9710   0.04437   0.03526  -0.0189   0.1246   1.0000
   8.250   0.9861   0.04763   0.03875  -0.0176   0.1133   1.0000
   8.500   0.9900   0.05211   0.04399  -0.0154   0.1098   1.0000
   8.750   0.9916   0.05675   0.04916  -0.0135   0.1072   1.0000
   9.000   1.0000   0.06059   0.05306  -0.0123   0.0997   1.0000
   9.250   0.9918   0.06540   0.05833  -0.0107   0.0992   1.0000
   9.500   0.9815   0.07042   0.06367  -0.0096   0.0995   1.0000
   9.750   0.9727   0.07562   0.06905  -0.0089   0.1003   1.0000
  10.000   0.9654   0.08095   0.07454  -0.0085   0.1016   1.0000
  10.250   0.8763   0.09124   0.08501  -0.0154   0.1241   1.0000
  10.500   0.8681   0.09793   0.09168  -0.0181   0.1265   1.0000
  10.750   0.8695   0.10424   0.09798  -0.0194   0.1277   1.0000
  11.000   0.7663   0.13411   0.12742  -0.0544   0.2928   1.0000
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