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HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il)
Reynolds number: 100,000
Max Cl/Cd: 47.97 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq108-il-100000.txt
Download as CSV file: xf-hq108-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5546   0.09532   0.09058  -0.0102   1.0000   0.0945
  -8.500  -0.5748   0.09111   0.08650  -0.0181   1.0000   0.0963
  -8.250  -0.5978   0.08667   0.08203  -0.0262   1.0000   0.0968
  -8.000  -0.5571   0.08317   0.07861  -0.0156   1.0000   0.1029
  -7.750  -0.5640   0.07880   0.07430  -0.0198   1.0000   0.1063
  -7.250  -0.5678   0.06953   0.06501  -0.0261   1.0000   0.1148
  -7.000  -0.5795   0.06533   0.06048  -0.0319   1.0000   0.1244
  -6.750  -0.5595   0.06156   0.05693  -0.0290   1.0000   0.1297
  -6.500  -0.5533   0.05792   0.05323  -0.0297   1.0000   0.1424
  -5.750  -0.5108   0.03828   0.03164  -0.0334   1.0000   0.0727
  -5.500  -0.4886   0.03416   0.02655  -0.0313   1.0000   0.0578
  -5.250  -0.4689   0.03018   0.02235  -0.0302   1.0000   0.0551
  -5.000  -0.4477   0.02707   0.01883  -0.0288   1.0000   0.0537
  -4.750  -0.4251   0.02452   0.01587  -0.0273   1.0000   0.0538
  -4.500  -0.4017   0.02244   0.01343  -0.0258   1.0000   0.0557
  -4.250  -0.3802   0.02059   0.01154  -0.0246   1.0000   0.0637
  -4.000  -0.3581   0.01890   0.00969  -0.0230   1.0000   0.0725
  -3.750  -0.3382   0.01731   0.00826  -0.0214   1.0000   0.0901
  -3.500  -0.3197   0.01583   0.00702  -0.0198   1.0000   0.1312
  -3.250  -0.3043   0.01348   0.00582  -0.0183   1.0000   0.2921
  -3.000  -0.2998   0.01209   0.00614  -0.0127   1.0000   0.6787
  -2.750  -0.2878   0.01209   0.00621  -0.0084   1.0000   0.7558
  -2.500  -0.2768   0.01209   0.00622  -0.0039   1.0000   0.8118
  -2.250  -0.2654   0.01206   0.00618   0.0005   1.0000   0.8618
  -2.000  -0.2447   0.01206   0.00609   0.0032   1.0000   0.9141
  -1.750  -0.1542   0.01242   0.00609  -0.0069   1.0000   0.9738
  -1.500  -0.0650   0.01237   0.00566  -0.0188   1.0000   0.9978
  -1.250  -0.0801   0.01213   0.00541  -0.0137   1.0000   1.0000
  -1.000  -0.0911   0.01196   0.00517  -0.0089   1.0000   1.0000
  -0.750  -0.0794   0.01201   0.00507  -0.0075   1.0000   1.0000
  -0.500  -0.0621   0.01215   0.00507  -0.0069   1.0000   1.0000
  -0.250  -0.0431   0.01234   0.00515  -0.0064   1.0000   1.0000
   0.000  -0.0235   0.01258   0.00529  -0.0061   1.0000   1.0000
   0.250   0.0113   0.01296   0.00556  -0.0086   0.9950   1.0000
   0.500   0.0579   0.01338   0.00593  -0.0132   0.9852   1.0000
   0.750   0.1037   0.01377   0.00629  -0.0176   0.9750   1.0000
   1.000   0.1505   0.01413   0.00665  -0.0220   0.9650   1.0000
   1.250   0.2016   0.01437   0.00693  -0.0270   0.9529   1.0000
   1.500   0.2568   0.01439   0.00702  -0.0323   0.9366   1.0000
   1.750   0.3140   0.01427   0.00705  -0.0376   0.9217   1.0000
   2.000   0.3610   0.01420   0.00710  -0.0410   0.9083   1.0000
   2.250   0.4027   0.01412   0.00714  -0.0432   0.8937   1.0000
   2.500   0.4398   0.01401   0.00716  -0.0442   0.8776   1.0000
   2.750   0.4736   0.01387   0.00719  -0.0443   0.8604   1.0000
   3.000   0.4992   0.01380   0.00723  -0.0428   0.8381   1.0000
   3.250   0.5241   0.01364   0.00717  -0.0408   0.8139   1.0000
   3.500   0.5470   0.01346   0.00707  -0.0383   0.7852   1.0000
   3.750   0.5692   0.01325   0.00691  -0.0356   0.7519   1.0000
   4.000   0.5907   0.01309   0.00684  -0.0329   0.7102   1.0000
   4.250   0.6119   0.01305   0.00679  -0.0304   0.6565   1.0000
   4.500   0.6322   0.01318   0.00680  -0.0278   0.5800   1.0000
   4.750   0.6499   0.01376   0.00690  -0.0251   0.4662   1.0000
   5.000   0.6656   0.01496   0.00742  -0.0228   0.3387   1.0000
   5.250   0.6803   0.01668   0.00840  -0.0209   0.2223   1.0000
   5.500   0.6966   0.01859   0.00968  -0.0192   0.1364   1.0000
   5.750   0.7150   0.02053   0.01140  -0.0176   0.0933   1.0000
   6.000   0.7355   0.02253   0.01329  -0.0163   0.0694   1.0000
   6.250   0.7590   0.02494   0.01571  -0.0152   0.0577   1.0000
   6.500   0.7818   0.02734   0.01814  -0.0144   0.0483   1.0000
   6.750   0.8068   0.03000   0.02123  -0.0131   0.0455   1.0000
   7.000   0.8292   0.03325   0.02498  -0.0117   0.0442   1.0000
   7.250   0.8480   0.03705   0.02934  -0.0101   0.0445   1.0000
   7.500   0.8626   0.04125   0.03411  -0.0084   0.0453   1.0000
   7.750   0.8735   0.04549   0.03887  -0.0068   0.0455   1.0000
   8.000   0.8804   0.04992   0.04377  -0.0053   0.0457   1.0000
   8.250   0.8837   0.05457   0.04881  -0.0040   0.0463   1.0000
   8.500   0.8850   0.05954   0.05403  -0.0031   0.0477   1.0000
   8.750   0.8727   0.06762   0.06263  -0.0024   0.0602   1.0000
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