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

HQ 2.0/12 AIRFOIL (hq2012-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.0/12 AIRFOIL (hq2012-il)
Reynolds number: 100,000
Max Cl/Cd: 52.4 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq2012-il-100000.txt
Download as CSV file: xf-hq2012-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/12 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4526   0.09468   0.08988  -0.0389   1.0000   0.1368
  -9.250  -0.4399   0.09081   0.08603  -0.0372   1.0000   0.1407
  -9.000  -0.4368   0.08804   0.08330  -0.0371   1.0000   0.1489
  -8.750  -0.4736   0.08318   0.07863  -0.0429   1.0000   0.1512
  -8.500  -0.4438   0.08087   0.07629  -0.0379   1.0000   0.1592
  -8.250  -0.4736   0.07707   0.07265  -0.0403   1.0000   0.1643
  -8.000  -0.5902   0.05746   0.05262  -0.0485   1.0000   0.0767
  -7.750  -0.6251   0.05116   0.04567  -0.0447   1.0000   0.0663
  -7.500  -0.6287   0.04772   0.04206  -0.0421   1.0000   0.0649
  -7.250  -0.6300   0.04422   0.03827  -0.0397   1.0000   0.0637
  -7.000  -0.6277   0.04096   0.03461  -0.0375   1.0000   0.0640
  -6.750  -0.6212   0.03807   0.03124  -0.0354   1.0000   0.0656
  -6.500  -0.6106   0.03542   0.02804  -0.0335   1.0000   0.0669
  -6.250  -0.5966   0.03313   0.02520  -0.0317   1.0000   0.0680
  -6.000  -0.5807   0.03006   0.02199  -0.0308   1.0000   0.0711
  -5.750  -0.5630   0.02888   0.02068  -0.0296   1.0000   0.0771
  -5.500  -0.5432   0.02713   0.01847  -0.0284   1.0000   0.0822
  -5.250  -0.5243   0.02586   0.01724  -0.0275   1.0000   0.0899
  -5.000  -0.5042   0.02455   0.01577  -0.0264   1.0000   0.0981
  -4.750  -0.4814   0.02383   0.01486  -0.0259   0.9991   0.1087
  -4.500  -0.4442   0.02281   0.01381  -0.0282   0.9942   0.1234
  -4.250  -0.4104   0.02170   0.01286  -0.0298   0.9889   0.1396
  -4.000  -0.3742   0.02093   0.01224  -0.0319   0.9837   0.1653
  -3.750  -0.3429   0.01996   0.01160  -0.0331   0.9778   0.2076
  -3.500  -0.3152   0.01816   0.01101  -0.0343   0.9722   0.3720
  -3.250  -0.2908   0.01803   0.01178  -0.0331   0.9656   0.5982
  -3.000  -0.2625   0.01856   0.01233  -0.0324   0.9583   0.6623
  -2.750  -0.2342   0.01907   0.01277  -0.0318   0.9515   0.7038
  -2.500  -0.2080   0.01948   0.01314  -0.0307   0.9440   0.7357
  -2.250  -0.1810   0.01986   0.01347  -0.0298   0.9373   0.7650
  -2.000  -0.1572   0.02010   0.01367  -0.0284   0.9295   0.7921
  -1.750  -0.1316   0.02031   0.01384  -0.0271   0.9228   0.8159
  -1.500  -0.1090   0.02039   0.01387  -0.0256   0.9148   0.8386
  -1.250  -0.0814   0.02047   0.01391  -0.0250   0.9081   0.8595
  -1.000  -0.0576   0.02049   0.01388  -0.0239   0.9000   0.8815
  -0.750  -0.0204   0.02060   0.01395  -0.0247   0.8946   0.9065
  -0.500   0.0154   0.02076   0.01408  -0.0260   0.8867   0.9285
  -0.250   0.0744   0.02087   0.01408  -0.0319   0.8825   0.9412
   0.000   0.1216   0.02104   0.01419  -0.0363   0.8752   0.9526
   0.250   0.1844   0.02108   0.01417  -0.0433   0.8703   0.9598
   0.500   0.2548   0.02092   0.01394  -0.0513   0.8670   0.9650
   0.750   0.3094   0.02080   0.01380  -0.0567   0.8573   0.9730
   1.000   0.3812   0.02000   0.01296  -0.0638   0.8514   0.9761
   1.250   0.4317   0.01954   0.01250  -0.0678   0.8391   0.9846
   1.500   0.4823   0.01912   0.01211  -0.0720   0.8279   0.9937
   1.750   0.5268   0.01871   0.01170  -0.0752   0.8188   1.0000
   2.000   0.5436   0.01853   0.01153  -0.0733   0.8079   1.0000
   2.250   0.5503   0.01856   0.01157  -0.0700   0.7947   1.0000
   2.500   0.5570   0.01858   0.01159  -0.0666   0.7819   1.0000
   2.750   0.5699   0.01854   0.01155  -0.0640   0.7697   1.0000
   3.000   0.5932   0.01834   0.01134  -0.0628   0.7586   1.0000
   3.250   0.6229   0.01797   0.01095  -0.0623   0.7481   1.0000
   3.500   0.6470   0.01784   0.01084  -0.0612   0.7342   1.0000
   3.750   0.6731   0.01763   0.01065  -0.0603   0.7199   1.0000
   4.000   0.7004   0.01738   0.01040  -0.0594   0.7048   1.0000
   4.250   0.7285   0.01709   0.01012  -0.0586   0.6887   1.0000
   4.500   0.7518   0.01699   0.01005  -0.0572   0.6677   1.0000
   4.750   0.7791   0.01671   0.00976  -0.0561   0.6465   1.0000
   5.000   0.8033   0.01656   0.00960  -0.0547   0.6199   1.0000
   5.250   0.8275   0.01644   0.00944  -0.0533   0.5890   1.0000
   5.500   0.8502   0.01645   0.00934  -0.0517   0.5524   1.0000
   5.750   0.8720   0.01664   0.00934  -0.0500   0.5127   1.0000
   6.000   0.8922   0.01707   0.00958  -0.0484   0.4725   1.0000
   6.250   0.9125   0.01763   0.00992  -0.0469   0.4375   1.0000
   6.500   0.9323   0.01826   0.01039  -0.0455   0.4058   1.0000
   6.750   0.9522   0.01893   0.01090  -0.0442   0.3780   1.0000
   7.000   0.9716   0.01960   0.01148  -0.0429   0.3521   1.0000
   7.250   0.9906   0.02029   0.01214  -0.0416   0.3276   1.0000
   7.500   1.0092   0.02105   0.01279  -0.0402   0.3042   1.0000
   7.750   1.0266   0.02186   0.01354  -0.0388   0.2801   1.0000
   8.000   1.0425   0.02275   0.01439  -0.0371   0.2548   1.0000
   8.250   1.0568   0.02374   0.01530  -0.0352   0.2284   1.0000
   8.500   1.0688   0.02477   0.01626  -0.0331   0.2014   1.0000
   8.750   1.0796   0.02583   0.01721  -0.0308   0.1764   1.0000
   9.000   1.0917   0.02700   0.01835  -0.0287   0.1553   1.0000
   9.250   1.1065   0.02833   0.01960  -0.0270   0.1401   1.0000
   9.500   1.1236   0.02977   0.02100  -0.0257   0.1283   1.0000
   9.750   1.1427   0.03130   0.02244  -0.0248   0.1186   1.0000
  10.000   1.1619   0.03277   0.02403  -0.0238   0.1107   1.0000
  10.250   1.1864   0.03479   0.02609  -0.0237   0.1044   1.0000
  10.500   1.2041   0.03640   0.02788  -0.0226   0.0986   1.0000
  10.750   1.2303   0.03882   0.03024  -0.0229   0.0932   1.0000
  11.000   1.2436   0.04091   0.03273  -0.0212   0.0899   1.0000
  11.250   1.2574   0.04313   0.03522  -0.0198   0.0865   1.0000
  11.500   1.2805   0.04588   0.03790  -0.0201   0.0818   1.0000
  11.750   1.2786   0.04836   0.04084  -0.0170   0.0795   1.0000
  12.000   1.2720   0.05066   0.04354  -0.0135   0.0772   1.0000
  12.250   1.2716   0.05228   0.04531  -0.0109   0.0733   1.0000
  12.500   1.2784   0.05462   0.04754  -0.0099   0.0681   1.0000
  12.750   1.2582   0.05672   0.05001  -0.0062   0.0664   1.0000
  13.000   1.2417   0.05936   0.05294  -0.0038   0.0642   1.0000
  13.250   1.2399   0.06123   0.05486  -0.0027   0.0603   1.0000
  13.500   1.2417   0.06492   0.05850  -0.0023   0.0567   1.0000
  13.750   1.2186   0.06891   0.06281  -0.0013   0.0563   1.0000
  14.000   1.1940   0.07358   0.06777  -0.0014   0.0559   1.0000
  14.250   1.1685   0.07905   0.07352  -0.0025   0.0558   1.0000
  14.500   1.1417   0.08541   0.08012  -0.0046   0.0560   1.0000
  14.750   1.1147   0.09247   0.08738  -0.0078   0.0562   1.0000
  15.000   1.0878   0.10039   0.09546  -0.0117   0.0567   1.0000
<< Back to HQ 2.0/12 AIRFOIL (hq2012-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.0/12 AIRFOIL (hq2012-il)