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HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/9 AIRFOIL (hq159-il)
Reynolds number: 100,000
Max Cl/Cd: 51.89 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq159-il-100000.txt
Download as CSV file: xf-hq159-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4982   0.09664   0.09183  -0.0193   1.0000   0.0986
  -8.750  -0.4941   0.09322   0.08845  -0.0200   1.0000   0.1030
  -8.500  -0.5072   0.08946   0.08480  -0.0249   1.0000   0.1070
  -8.250  -0.5373   0.08487   0.08030  -0.0341   1.0000   0.1081
  -8.000  -0.5070   0.08143   0.07689  -0.0266   1.0000   0.1126
  -7.750  -0.4282   0.06824   0.06401  -0.0274   1.0000   0.1281
  -7.500  -0.4431   0.06409   0.05995  -0.0298   1.0000   0.1323
  -7.250  -0.4939   0.05866   0.05437  -0.0388   1.0000   0.1357
  -7.000  -0.4615   0.05481   0.05074  -0.0335   1.0000   0.1411
  -6.750  -0.4879   0.04984   0.04560  -0.0374   1.0000   0.1504
  -6.500  -0.4701   0.04668   0.04256  -0.0342   1.0000   0.1557
  -6.250  -0.4760   0.04276   0.03859  -0.0340   1.0000   0.1667
  -6.000  -0.5038   0.04088   0.03460  -0.0398   1.0000   0.0699
  -5.750  -0.4886   0.03621   0.02907  -0.0374   1.0000   0.0556
  -5.500  -0.4725   0.03336   0.02591  -0.0356   1.0000   0.0544
  -5.250  -0.4571   0.03050   0.02260  -0.0340   1.0000   0.0556
  -5.000  -0.4406   0.02766   0.01966  -0.0328   1.0000   0.0587
  -4.750  -0.4206   0.02532   0.01687  -0.0311   1.0000   0.0590
  -4.500  -0.3998   0.02339   0.01469  -0.0296   1.0000   0.0613
  -4.250  -0.3782   0.02186   0.01287  -0.0281   1.0000   0.0662
  -4.000  -0.3578   0.02038   0.01138  -0.0269   1.0000   0.0771
  -3.750  -0.3367   0.01885   0.00986  -0.0255   1.0000   0.0906
  -3.500  -0.3167   0.01762   0.00877  -0.0243   1.0000   0.1154
  -3.250  -0.2967   0.01658   0.00789  -0.0231   1.0000   0.1504
  -3.000  -0.2777   0.01469   0.00692  -0.0223   1.0000   0.2789
  -2.750  -0.2700   0.01340   0.00732  -0.0178   1.0000   0.6621
  -2.500  -0.2585   0.01351   0.00753  -0.0138   1.0000   0.7385
  -2.250  -0.2472   0.01360   0.00764  -0.0099   1.0000   0.7917
  -2.000  -0.2362   0.01361   0.00766  -0.0059   1.0000   0.8385
  -1.750  -0.2233   0.01356   0.00762  -0.0024   1.0000   0.8867
  -1.500  -0.1721   0.01368   0.00763  -0.0059   1.0000   0.9625
  -1.250  -0.1266   0.01372   0.00745  -0.0110   1.0000   1.0000
  -1.000  -0.1079   0.01374   0.00729  -0.0115   0.9984   1.0000
  -0.750  -0.0622   0.01409   0.00741  -0.0163   0.9900   1.0000
  -0.500  -0.0147   0.01451   0.00764  -0.0214   0.9822   1.0000
  -0.250   0.0267   0.01479   0.00779  -0.0251   0.9727   1.0000
   0.000   0.0695   0.01512   0.00798  -0.0289   0.9638   1.0000
   0.250   0.1156   0.01544   0.00821  -0.0332   0.9557   1.0000
   0.500   0.1532   0.01569   0.00839  -0.0358   0.9453   1.0000
   0.750   0.1941   0.01593   0.00860  -0.0389   0.9357   1.0000
   1.000   0.2460   0.01606   0.00871  -0.0437   0.9267   1.0000
   1.250   0.2995   0.01594   0.00861  -0.0484   0.9140   1.0000
   1.500   0.3515   0.01568   0.00842  -0.0524   0.9006   1.0000
   1.750   0.3915   0.01557   0.00835  -0.0544   0.8877   1.0000
   2.000   0.4270   0.01552   0.00835  -0.0554   0.8749   1.0000
   2.250   0.4608   0.01544   0.00833  -0.0560   0.8616   1.0000
   2.500   0.4926   0.01535   0.00833  -0.0560   0.8477   1.0000
   2.750   0.5228   0.01523   0.00828  -0.0556   0.8328   1.0000
   3.000   0.5517   0.01507   0.00819  -0.0548   0.8172   1.0000
   3.250   0.5763   0.01500   0.00820  -0.0533   0.7981   1.0000
   3.500   0.6022   0.01483   0.00813  -0.0518   0.7784   1.0000
   3.750   0.6270   0.01465   0.00802  -0.0500   0.7556   1.0000
   4.000   0.6515   0.01442   0.00784  -0.0481   0.7289   1.0000
   4.250   0.6751   0.01420   0.00765  -0.0460   0.6961   1.0000
   4.500   0.6979   0.01406   0.00754  -0.0438   0.6550   1.0000
   4.750   0.7196   0.01404   0.00744  -0.0416   0.6013   1.0000
   5.000   0.7400   0.01426   0.00745  -0.0393   0.5329   1.0000
   5.250   0.7589   0.01484   0.00766  -0.0372   0.4569   1.0000
   5.500   0.7771   0.01567   0.00813  -0.0354   0.3854   1.0000
   5.750   0.7948   0.01664   0.00875  -0.0337   0.3196   1.0000
   6.000   0.8115   0.01787   0.00968  -0.0320   0.2535   1.0000
   6.250   0.8266   0.01940   0.01077  -0.0302   0.1834   1.0000
   6.500   0.8439   0.02092   0.01194  -0.0287   0.1420   1.0000
   6.750   0.8641   0.02245   0.01338  -0.0274   0.1196   1.0000
   7.000   0.8848   0.02399   0.01479  -0.0264   0.1029   1.0000
   7.250   0.9085   0.02581   0.01667  -0.0255   0.0916   1.0000
   7.500   0.9324   0.02794   0.01894  -0.0246   0.0812   1.0000
   7.750   0.9530   0.02993   0.02105  -0.0236   0.0688   1.0000
   8.000   0.9708   0.03187   0.02320  -0.0224   0.0561   1.0000
   8.250   0.9894   0.03515   0.02665  -0.0213   0.0489   1.0000
   8.500   1.0057   0.03849   0.03059  -0.0194   0.0456   1.0000
   8.750   1.0185   0.04199   0.03455  -0.0175   0.0434   1.0000
   9.000   1.0316   0.04492   0.03758  -0.0164   0.0402   1.0000
   9.250   1.0328   0.05059   0.04359  -0.0149   0.0387   1.0000
   9.500   1.0320   0.05360   0.04714  -0.0123   0.0380   1.0000
   9.750   1.0273   0.05760   0.05155  -0.0101   0.0379   1.0000
  10.000   1.0180   0.06174   0.05602  -0.0080   0.0379   1.0000
  10.250   1.0056   0.06609   0.06060  -0.0062   0.0382   1.0000
  10.500   0.8813   0.06456   0.05998  -0.0005   0.0441   1.0000
  10.750   0.8496   0.07150   0.06709  -0.0025   0.0462   1.0000
  11.000   0.8238   0.07865   0.07434  -0.0055   0.0476   1.0000
  11.250   0.8011   0.08619   0.08195  -0.0091   0.0488   1.0000
  11.500   0.7834   0.09375   0.08954  -0.0126   0.0497   1.0000
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