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HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/9 AIRFOIL (hq109-il)
Reynolds number: 500,000
Max Cl/Cd: 79.53 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq109-il-500000.txt
Download as CSV file: xf-hq109-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/9 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.000  -0.6038   0.16479   0.16236   0.0194   1.0000   0.0114
 -13.750  -0.5981   0.16084   0.15841   0.0176   1.0000   0.0121
  -7.250  -0.6322   0.03037   0.02636  -0.0338   1.0000   0.0151
  -7.000  -0.6251   0.02451   0.02002  -0.0312   1.0000   0.0113
  -6.750  -0.6135   0.01980   0.01466  -0.0284   1.0000   0.0104
  -6.500  -0.5959   0.01794   0.01256  -0.0266   1.0000   0.0106
  -6.250  -0.5780   0.01672   0.01116  -0.0249   1.0000   0.0112
  -6.000  -0.5613   0.01567   0.00999  -0.0230   1.0000   0.0119
  -5.750  -0.5359   0.01455   0.00874  -0.0228   0.9985   0.0126
  -5.500  -0.5003   0.01354   0.00759  -0.0246   0.9950   0.0136
  -5.250  -0.4670   0.01229   0.00621  -0.0261   0.9906   0.0157
  -5.000  -0.4318   0.01148   0.00535  -0.0279   0.9862   0.0188
  -4.750  -0.3947   0.01084   0.00460  -0.0299   0.9830   0.0226
  -4.500  -0.3607   0.01007   0.00388  -0.0314   0.9778   0.0404
  -4.250  -0.3252   0.00970   0.00352  -0.0332   0.9726   0.0563
  -4.000  -0.2899   0.00922   0.00313  -0.0350   0.9678   0.0789
  -3.750  -0.2596   0.00866   0.00281  -0.0358   0.9587   0.1342
  -3.500  -0.2303   0.00802   0.00250  -0.0365   0.9495   0.2258
  -3.250  -0.2034   0.00730   0.00222  -0.0368   0.9396   0.3444
  -3.000  -0.1790   0.00670   0.00204  -0.0363   0.9282   0.4730
  -2.750  -0.1541   0.00640   0.00195  -0.0356   0.9168   0.5524
  -2.500  -0.1283   0.00626   0.00188  -0.0350   0.9060   0.5969
  -2.250  -0.1022   0.00619   0.00183  -0.0344   0.8952   0.6300
  -2.000  -0.0762   0.00614   0.00179  -0.0338   0.8845   0.6596
  -1.750  -0.0500   0.00611   0.00176  -0.0332   0.8738   0.6830
  -1.500  -0.0234   0.00609   0.00173  -0.0328   0.8639   0.7059
  -1.250   0.0030   0.00608   0.00173  -0.0322   0.8550   0.7269
  -1.000   0.0296   0.00608   0.00171  -0.0317   0.8450   0.7430
  -0.750   0.0561   0.00607   0.00169  -0.0313   0.8335   0.7583
  -0.500   0.0828   0.00606   0.00167  -0.0308   0.8223   0.7722
  -0.250   0.1095   0.00606   0.00165  -0.0304   0.8118   0.7851
   0.000   0.1361   0.00605   0.00164  -0.0300   0.8005   0.7978
   0.250   0.1626   0.00604   0.00162  -0.0295   0.7889   0.8110
   0.500   0.1892   0.00602   0.00163  -0.0290   0.7774   0.8250
   0.750   0.2158   0.00600   0.00163  -0.0286   0.7660   0.8379
   1.000   0.2425   0.00599   0.00163  -0.0282   0.7545   0.8501
   1.250   0.2689   0.00598   0.00163  -0.0277   0.7429   0.8632
   1.500   0.2951   0.00596   0.00165  -0.0272   0.7305   0.8779
   1.750   0.3210   0.00595   0.00167  -0.0266   0.7171   0.8951
   2.000   0.3469   0.00593   0.00169  -0.0259   0.7018   0.9152
   2.250   0.3756   0.00594   0.00172  -0.0259   0.6835   0.9401
   2.500   0.4111   0.00599   0.00177  -0.0274   0.6629   0.9635
   2.750   0.4495   0.00609   0.00180  -0.0297   0.6337   0.9831
   3.000   0.4880   0.00624   0.00184  -0.0321   0.5935   1.0000
   3.250   0.5114   0.00643   0.00192  -0.0313   0.5572   1.0000
   3.500   0.5348   0.00673   0.00204  -0.0305   0.5068   1.0000
   3.750   0.5578   0.00715   0.00221  -0.0298   0.4441   1.0000
   4.000   0.5806   0.00766   0.00243  -0.0290   0.3737   1.0000
   4.250   0.6038   0.00820   0.00269  -0.0284   0.3072   1.0000
   4.500   0.6277   0.00868   0.00298  -0.0279   0.2622   1.0000
   4.750   0.6520   0.00913   0.00328  -0.0275   0.2223   1.0000
   5.000   0.6768   0.00953   0.00355  -0.0271   0.1851   1.0000
   5.250   0.7009   0.01002   0.00387  -0.0266   0.1433   1.0000
   5.500   0.7243   0.01062   0.00427  -0.0261   0.1024   1.0000
   5.750   0.7479   0.01120   0.00470  -0.0256   0.0717   1.0000
   6.000   0.7715   0.01179   0.00518  -0.0251   0.0500   1.0000
   6.250   0.7956   0.01231   0.00568  -0.0245   0.0358   1.0000
   6.500   0.8190   0.01298   0.00633  -0.0238   0.0243   1.0000
   6.750   0.8411   0.01386   0.00723  -0.0229   0.0159   1.0000
   7.000   0.8628   0.01475   0.00816  -0.0220   0.0122   1.0000
   7.250   0.8827   0.01592   0.00946  -0.0207   0.0109   1.0000
   7.500   0.9047   0.01675   0.01040  -0.0198   0.0101   1.0000
   7.750   0.9258   0.01773   0.01152  -0.0188   0.0092   1.0000
   8.000   0.9464   0.01876   0.01264  -0.0178   0.0084   1.0000
   8.250   0.9658   0.02002   0.01402  -0.0166   0.0081   1.0000
   8.500   0.9836   0.02165   0.01578  -0.0153   0.0077   1.0000
   8.750   1.0007   0.02355   0.01789  -0.0139   0.0076   1.0000
   9.000   1.0168   0.02572   0.02029  -0.0124   0.0076   1.0000
   9.250   1.0310   0.02813   0.02300  -0.0108   0.0076   1.0000
   9.500   1.0404   0.03151   0.02683  -0.0085   0.0080   1.0000
   9.750   1.0213   0.03975   0.03594  -0.0037   0.0094   1.0000
  10.000   1.0065   0.04468   0.04126  -0.0005   0.0101   1.0000
  10.250   0.9839   0.04865   0.04548   0.0034   0.0105   1.0000
  10.500   0.9612   0.05250   0.04952   0.0056   0.0107   1.0000
  10.750   0.8854   0.04699   0.04432   0.0064   0.0103   1.0000
  11.000   0.8624   0.05277   0.05026   0.0046   0.0104   1.0000
  11.250   0.8409   0.05933   0.05696   0.0016   0.0103   1.0000
  11.500   0.8175   0.06746   0.06523  -0.0029   0.0103   1.0000
  11.750   0.7961   0.07653   0.07442  -0.0085   0.0102   1.0000
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