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HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il)
Reynolds number: 50,000
Max Cl/Cd: 32.43 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1010-il-50000.txt
Download as CSV file: xf-hq1010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/10 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4861   0.09645   0.08966   0.0036   1.0000   0.3518
  -8.250  -0.4849   0.09306   0.08632   0.0036   1.0000   0.3643
  -8.000  -0.5806   0.07561   0.06927  -0.0259   1.0000   0.1714
  -7.500  -0.6170   0.06048   0.05366  -0.0348   1.0000   0.1316
  -7.250  -0.6255   0.05445   0.04692  -0.0357   1.0000   0.1209
  -7.000  -0.6145   0.04993   0.04221  -0.0350   1.0000   0.1184
  -6.750  -0.6051   0.04556   0.03743  -0.0340   1.0000   0.1165
  -6.500  -0.5935   0.04180   0.03317  -0.0327   1.0000   0.1172
  -6.250  -0.5810   0.03869   0.02928  -0.0309   1.0000   0.1215
  -6.000  -0.5627   0.03561   0.02625  -0.0297   1.0000   0.1283
  -5.750  -0.5444   0.03290   0.02285  -0.0279   1.0000   0.1352
  -5.500  -0.5252   0.03065   0.02039  -0.0265   1.0000   0.1494
  -5.250  -0.5041   0.02840   0.01808  -0.0250   1.0000   0.1644
  -5.000  -0.4836   0.02658   0.01634  -0.0235   1.0000   0.1855
  -4.750  -0.4625   0.02482   0.01463  -0.0219   1.0000   0.2110
  -4.500  -0.4424   0.02313   0.01308  -0.0201   1.0000   0.2489
  -4.250  -0.4250   0.02108   0.01163  -0.0181   1.0000   0.3195
  -4.000  -0.4220   0.01903   0.01137  -0.0123   1.0000   0.5242
  -3.750  -0.4224   0.01954   0.01234  -0.0032   1.0000   0.6785
  -3.500  -0.4189   0.02012   0.01290   0.0048   1.0000   0.7501
  -3.250  -0.4125   0.02049   0.01320   0.0120   1.0000   0.8062
  -3.000  -0.3917   0.02098   0.01349   0.0175   1.0000   0.8661
  -2.750  -0.1589   0.02213   0.01318  -0.0125   1.0000   0.9584
  -2.500  -0.0513   0.02087   0.01134  -0.0287   1.0000   0.9930
  -2.250  -0.0252   0.02027   0.01062  -0.0309   1.0000   1.0000
  -2.000  -0.0262   0.02004   0.01039  -0.0280   1.0000   1.0000
  -1.750  -0.0325   0.01989   0.01024  -0.0244   1.0000   1.0000
  -1.500  -0.0429   0.01977   0.01011  -0.0201   1.0000   1.0000
  -1.250  -0.0555   0.01963   0.00993  -0.0156   1.0000   1.0000
  -1.000  -0.0692   0.01943   0.00972  -0.0109   1.0000   1.0000
  -0.750  -0.0824   0.01918   0.00944  -0.0062   1.0000   1.0000
  -0.500  -0.0895   0.01899   0.00917  -0.0024   1.0000   1.0000
  -0.250  -0.0855   0.01896   0.00902  -0.0002   1.0000   1.0000
   0.000  -0.0737   0.01909   0.00902   0.0009   1.0000   1.0000
   0.250  -0.0584   0.01931   0.00912   0.0015   1.0000   1.0000
   0.500  -0.0415   0.01961   0.00931   0.0019   1.0000   1.0000
   0.750  -0.0239   0.01997   0.00955   0.0021   1.0000   1.0000
   1.000  -0.0059   0.02038   0.00988   0.0023   1.0000   1.0000
   1.250   0.0123   0.02084   0.01028   0.0025   1.0000   1.0000
   1.500   0.0305   0.02135   0.01075   0.0025   1.0000   1.0000
   1.750   0.0485   0.02192   0.01129   0.0025   1.0000   1.0000
   2.000   0.0664   0.02255   0.01189   0.0025   1.0000   1.0000
   2.250   0.1235   0.02403   0.01342  -0.0049   0.9827   1.0000
   2.500   0.1858   0.02533   0.01485  -0.0128   0.9579   1.0000
   2.750   0.2429   0.02637   0.01602  -0.0192   0.9329   1.0000
   3.000   0.2988   0.02725   0.01706  -0.0249   0.9095   1.0000
   3.250   0.3548   0.02791   0.01797  -0.0301   0.8855   1.0000
   3.500   0.4030   0.02837   0.01866  -0.0336   0.8597   1.0000
   3.750   0.4590   0.02852   0.01914  -0.0377   0.8321   1.0000
   4.000   0.5175   0.02822   0.01919  -0.0410   0.8029   1.0000
   4.250   0.5728   0.02735   0.01867  -0.0423   0.7719   1.0000
   4.500   0.6089   0.02660   0.01821  -0.0405   0.7353   1.0000
   4.750   0.6463   0.02536   0.01719  -0.0379   0.6959   1.0000
   5.000   0.6767   0.02420   0.01615  -0.0342   0.6503   1.0000
   5.250   0.7032   0.02328   0.01521  -0.0303   0.5983   1.0000
   5.500   0.7252   0.02287   0.01470  -0.0265   0.5384   1.0000
   5.750   0.7456   0.02299   0.01447  -0.0231   0.4752   1.0000
   6.000   0.7630   0.02386   0.01499  -0.0202   0.4093   1.0000
   6.250   0.7810   0.02524   0.01591  -0.0178   0.3473   1.0000
   6.500   0.7996   0.02700   0.01726  -0.0158   0.2919   1.0000
   6.750   0.8192   0.02903   0.01916  -0.0142   0.2469   1.0000
   7.000   0.8403   0.03117   0.02127  -0.0128   0.2125   1.0000
   7.250   0.8600   0.03312   0.02319  -0.0114   0.1839   1.0000
   7.500   0.8807   0.03557   0.02575  -0.0102   0.1627   1.0000
   7.750   0.8995   0.03800   0.02837  -0.0087   0.1443   1.0000
   8.000   0.9170   0.04119   0.03188  -0.0073   0.1322   1.0000
   8.250   0.9313   0.04425   0.03522  -0.0058   0.1207   1.0000
   8.500   0.9486   0.04784   0.03898  -0.0046   0.1138   1.0000
   8.750   0.9522   0.05215   0.04386  -0.0026   0.1100   1.0000
   9.000   0.9499   0.05639   0.04866  -0.0007   0.1067   1.0000
   9.250   0.9467   0.06068   0.05334   0.0009   0.1043   1.0000
   9.500   0.9348   0.06569   0.05872   0.0022   0.1049   1.0000
   9.750   0.9156   0.07084   0.06414   0.0030   0.1068   1.0000
  10.000   0.8934   0.07577   0.06923   0.0035   0.1084   1.0000
  10.250   0.8741   0.08088   0.07440   0.0030   0.1100   1.0000
  10.500   0.8594   0.08663   0.08018   0.0015   0.1114   1.0000
  10.750   0.8077   0.09914   0.09271  -0.0086   0.1271   1.0000
  11.000   0.8097   0.10460   0.09813  -0.0096   0.1257   1.0000
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