Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 3.5/8 AIRFOIL (hq358-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 3.5/8 AIRFOIL (hq358-il)
Reynolds number: 200,000
Max Cl/Cd: 84.17 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq358-il-200000-n5.txt
Download as CSV file: xf-hq358-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.5/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3387   0.09357   0.09020  -0.0328   1.0000   0.0167
  -7.750  -0.3438   0.09153   0.08823  -0.0316   1.0000   0.0167
  -7.500  -0.3540   0.08998   0.08676  -0.0296   1.0000   0.0170
  -7.250  -0.3405   0.08655   0.08336  -0.0369   0.9934   0.0180
  -5.750  -0.1852   0.05527   0.05167  -0.0838   0.9614   0.0185
  -4.750  -0.0793   0.03245   0.02783  -0.0999   0.9395   0.0106
  -4.500  -0.0508   0.03093   0.02616  -0.1026   0.9360   0.0145
  -4.250  -0.0203   0.02719   0.02200  -0.1040   0.9309   0.0139
  -4.000   0.0110   0.02350   0.01781  -0.1049   0.9259   0.0128
  -3.750   0.0448   0.02025   0.01400  -0.1059   0.9226   0.0122
  -3.500   0.0748   0.01806   0.01135  -0.1060   0.9179   0.0121
  -3.250   0.1040   0.01639   0.00937  -0.1060   0.9128   0.0123
  -3.000   0.1351   0.01504   0.00780  -0.1064   0.9091   0.0130
  -2.750   0.1631   0.01423   0.00688  -0.1063   0.9038   0.0154
  -2.500   0.1917   0.01350   0.00604  -0.1063   0.8984   0.0176
  -2.250   0.2223   0.01270   0.00509  -0.1067   0.8944   0.0183
  -2.000   0.2492   0.01212   0.00439  -0.1065   0.8877   0.0203
  -1.750   0.2789   0.01173   0.00387  -0.1067   0.8826   0.0281
  -1.500   0.3070   0.01140   0.00358  -0.1067   0.8765   0.0627
  -1.250   0.3340   0.01024   0.00339  -0.1074   0.8703   0.3634
  -0.750   0.3798   0.00930   0.00358  -0.1046   0.8554   0.7567
  -0.500   0.4024   0.00912   0.00344  -0.1030   0.8451   0.8074
  -0.250   0.4266   0.00895   0.00327  -0.1017   0.8346   0.8424
   0.000   0.4532   0.00881   0.00309  -0.1011   0.8240   0.8673
   0.250   0.4827   0.00864   0.00291  -0.1010   0.8134   0.9114
   0.500   0.5164   0.00858   0.00278  -0.1022   0.8014   1.0000
   0.750   0.5441   0.00862   0.00271  -0.1021   0.7886   1.0000
   1.000   0.5718   0.00866   0.00266  -0.1020   0.7759   1.0000
   1.250   0.5994   0.00873   0.00265  -0.1019   0.7639   1.0000
   1.500   0.6269   0.00880   0.00268  -0.1018   0.7517   1.0000
   1.750   0.6542   0.00888   0.00269  -0.1016   0.7384   1.0000
   2.000   0.6813   0.00897   0.00273  -0.1014   0.7239   1.0000
   2.250   0.7082   0.00908   0.00277  -0.1012   0.7082   1.0000
   2.500   0.7349   0.00920   0.00283  -0.1009   0.6906   1.0000
   2.750   0.7612   0.00934   0.00296  -0.1005   0.6704   1.0000
   3.000   0.7872   0.00951   0.00306  -0.1001   0.6482   1.0000
   3.250   0.8127   0.00971   0.00318  -0.0995   0.6215   1.0000
   3.500   0.8375   0.00995   0.00333  -0.0989   0.5891   1.0000
   3.750   0.8613   0.01027   0.00350  -0.0981   0.5503   1.0000
   4.000   0.8839   0.01069   0.00378  -0.0971   0.5043   1.0000
   4.250   0.9056   0.01121   0.00408  -0.0960   0.4555   1.0000
   4.500   0.9271   0.01175   0.00444  -0.0950   0.4108   1.0000
   4.750   0.9492   0.01229   0.00483  -0.0941   0.3725   1.0000
   5.000   0.9715   0.01281   0.00525  -0.0932   0.3391   1.0000
   5.250   0.9935   0.01337   0.00576  -0.0924   0.3067   1.0000
   5.500   1.0130   0.01416   0.00628  -0.0913   0.2599   1.0000
   5.750   1.0345   0.01477   0.00678  -0.0904   0.2287   1.0000
   6.000   1.0579   0.01522   0.00727  -0.0898   0.2085   1.0000
   6.250   1.0787   0.01592   0.00782  -0.0889   0.1646   1.0000
   6.500   1.0938   0.01730   0.00867  -0.0873   0.0860   1.0000
   6.750   1.1048   0.01931   0.01009  -0.0852   0.0130   1.0000
   7.000   1.1240   0.02032   0.01117  -0.0838   0.0076   1.0000
   7.250   1.1438   0.02120   0.01230  -0.0825   0.0066   1.0000
   7.500   1.1626   0.02218   0.01348  -0.0811   0.0060   1.0000
   7.750   1.1798   0.02329   0.01478  -0.0795   0.0056   1.0000
   8.000   1.1950   0.02456   0.01624  -0.0776   0.0053   1.0000
   8.250   1.2081   0.02597   0.01783  -0.0756   0.0051   1.0000
   8.500   1.2191   0.02754   0.01957  -0.0732   0.0050   1.0000
   8.750   1.2286   0.02924   0.02143  -0.0707   0.0049   1.0000
   9.000   1.2364   0.03111   0.02346  -0.0680   0.0048   1.0000
   9.250   1.2447   0.03313   0.02566  -0.0655   0.0048   1.0000
   9.500   1.2533   0.03533   0.02805  -0.0632   0.0047   1.0000
   9.750   1.2615   0.03776   0.03070  -0.0609   0.0047   1.0000
  10.000   1.2681   0.04039   0.03358  -0.0586   0.0047   1.0000
  10.250   1.2703   0.04347   0.03694  -0.0562   0.0046   1.0000
  10.750   1.2619   0.04974   0.04392  -0.0511   0.0042   1.0000
  11.000   1.2570   0.05269   0.04715  -0.0491   0.0041   1.0000
  11.250   1.2520   0.05556   0.05031  -0.0477   0.0039   1.0000
  11.500   1.2431   0.05924   0.05425  -0.0467   0.0037   1.0000
  11.750   1.2320   0.06339   0.05867  -0.0463   0.0036   1.0000
  12.000   1.2189   0.06812   0.06365  -0.0467   0.0036   1.0000
  12.250   1.2043   0.07344   0.06920  -0.0480   0.0035   1.0000
  12.500   1.1887   0.07940   0.07538  -0.0502   0.0035   1.0000
  12.750   1.1721   0.08611   0.08229  -0.0535   0.0036   1.0000
  13.000   1.1551   0.09360   0.08998  -0.0577   0.0036   1.0000
  13.250   1.1380   0.10195   0.09851  -0.0629   0.0037   1.0000
<< Back to HQ 3.5/8 AIRFOIL (hq358-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.5/8 AIRFOIL (hq358-il)