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HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 1.5/8 AIRFOIL (hq158-il)
Reynolds number: 50,000
Max Cl/Cd: 35.92 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq158-il-50000-n5.txt
Download as CSV file: xf-hq158-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5052   0.09486   0.08805  -0.0186   1.0000   0.0395
  -8.500  -0.5044   0.09077   0.08404  -0.0207   1.0000   0.0385
  -8.250  -0.5058   0.08632   0.07967  -0.0232   1.0000   0.0376
  -8.000  -0.5087   0.08173   0.07517  -0.0263   1.0000   0.0367
  -7.750  -0.5116   0.07660   0.07009  -0.0306   1.0000   0.0358
  -7.500  -0.5131   0.07156   0.06504  -0.0340   1.0000   0.0348
  -7.250  -0.5139   0.06640   0.05979  -0.0369   1.0000   0.0339
  -7.000  -0.5125   0.06132   0.05453  -0.0392   1.0000   0.0331
  -6.750  -0.5074   0.05691   0.04992  -0.0402   1.0000   0.0331
  -6.500  -0.4989   0.05327   0.04608  -0.0406   1.0000   0.0339
  -6.250  -0.4885   0.04975   0.04232  -0.0408   1.0000   0.0353
  -6.000  -0.4762   0.04628   0.03852  -0.0407   1.0000   0.0370
  -5.750  -0.4620   0.04277   0.03458  -0.0403   1.0000   0.0387
  -5.500  -0.4457   0.03932   0.03055  -0.0396   1.0000   0.0394
  -5.250  -0.4273   0.03610   0.02684  -0.0387   1.0000   0.0399
  -5.000  -0.4071   0.03315   0.02338  -0.0375   1.0000   0.0408
  -4.750  -0.3852   0.03055   0.02025  -0.0362   1.0000   0.0422
  -4.500  -0.3630   0.02821   0.01755  -0.0348   1.0000   0.0442
  -4.250  -0.3427   0.02658   0.01585  -0.0337   1.0000   0.0507
  -4.000  -0.3212   0.02503   0.01409  -0.0322   1.0000   0.0593
  -3.750  -0.3002   0.02355   0.01244  -0.0304   1.0000   0.0677
  -3.500  -0.2800   0.02230   0.01117  -0.0292   1.0000   0.0866
  -3.250  -0.2590   0.02093   0.00981  -0.0281   1.0000   0.1140
  -3.000  -0.2384   0.01939   0.00855  -0.0275   1.0000   0.1785
  -2.750  -0.2248   0.01709   0.00810  -0.0255   1.0000   0.5066
  -2.500  -0.2151   0.01672   0.00819  -0.0205   1.0000   0.6898
  -2.250  -0.2045   0.01653   0.00809  -0.0158   1.0000   0.7855
  -2.000  -0.1690   0.01623   0.00786  -0.0144   1.0000   0.9237
  -1.750  -0.1115   0.01609   0.00727  -0.0206   1.0000   1.0000
  -1.500  -0.1000   0.01604   0.00688  -0.0190   1.0000   1.0000
  -1.250  -0.0833   0.01607   0.00664  -0.0181   1.0000   1.0000
  -1.000  -0.0644   0.01616   0.00648  -0.0175   1.0000   1.0000
  -0.750  -0.0446   0.01629   0.00639  -0.0171   1.0000   1.0000
  -0.500  -0.0210   0.01649   0.00637  -0.0173   0.9984   1.0000
  -0.250   0.0191   0.01680   0.00647  -0.0206   0.9885   1.0000
   0.000   0.0585   0.01711   0.00658  -0.0237   0.9786   1.0000
   0.250   0.0980   0.01742   0.00676  -0.0267   0.9685   1.0000
   0.500   0.1389   0.01773   0.00697  -0.0299   0.9585   1.0000
   0.750   0.1776   0.01799   0.00717  -0.0326   0.9470   1.0000
   1.000   0.2158   0.01823   0.00738  -0.0350   0.9351   1.0000
   1.250   0.2539   0.01846   0.00761  -0.0374   0.9231   1.0000
   1.500   0.2931   0.01866   0.00788  -0.0398   0.9110   1.0000
   1.750   0.3319   0.01883   0.00811  -0.0420   0.8984   1.0000
   2.000   0.3684   0.01898   0.00835  -0.0437   0.8849   1.0000
   2.250   0.4039   0.01909   0.00856  -0.0450   0.8701   1.0000
   2.500   0.4400   0.01911   0.00877  -0.0460   0.8532   1.0000
   2.750   0.4729   0.01910   0.00890  -0.0463   0.8340   1.0000
   3.000   0.5033   0.01909   0.00904  -0.0460   0.8130   1.0000
   3.250   0.5322   0.01909   0.00920  -0.0454   0.7913   1.0000
   3.500   0.5614   0.01906   0.00934  -0.0447   0.7691   1.0000
   3.750   0.5877   0.01907   0.00962  -0.0435   0.7427   1.0000
   4.000   0.6134   0.01907   0.00979  -0.0421   0.7121   1.0000
   4.250   0.6384   0.01906   0.00995  -0.0404   0.6750   1.0000
   4.500   0.6622   0.01910   0.01010  -0.0385   0.6275   1.0000
   4.750   0.6849   0.01923   0.01019  -0.0363   0.5651   1.0000
   5.000   0.7054   0.01964   0.01035  -0.0339   0.4855   1.0000
   5.250   0.7235   0.02048   0.01091  -0.0317   0.4031   1.0000
   5.500   0.7407   0.02156   0.01168  -0.0299   0.3250   1.0000
   5.750   0.7578   0.02286   0.01270  -0.0285   0.2586   1.0000
   6.000   0.7750   0.02429   0.01387  -0.0272   0.1949   1.0000
   6.250   0.7923   0.02592   0.01520  -0.0260   0.1404   1.0000
   6.500   0.8110   0.02773   0.01688  -0.0248   0.1123   1.0000
   6.750   0.8305   0.02944   0.01857  -0.0238   0.0879   1.0000
   7.000   0.8516   0.03156   0.02080  -0.0226   0.0746   1.0000
   7.250   0.8724   0.03384   0.02319  -0.0215   0.0611   1.0000
   7.500   0.8926   0.03632   0.02603  -0.0204   0.0492   1.0000
   7.750   0.9112   0.03887   0.02896  -0.0192   0.0399   1.0000
   8.000   0.9281   0.04198   0.03219  -0.0183   0.0359   1.0000
   8.250   0.9437   0.04576   0.03655  -0.0168   0.0337   1.0000
   8.500   0.9539   0.04952   0.04093  -0.0152   0.0314   1.0000
   8.750   0.9596   0.05311   0.04499  -0.0138   0.0291   1.0000
   9.000   0.9624   0.05652   0.04873  -0.0125   0.0274   1.0000
   9.250   0.9612   0.06000   0.05251  -0.0112   0.0263   1.0000
   9.500   0.9560   0.06363   0.05638  -0.0101   0.0255   1.0000
   9.750   0.9445   0.06747   0.06044  -0.0089   0.0252   1.0000
  10.000   0.9275   0.07182   0.06503  -0.0085   0.0255   1.0000
  10.250   0.9058   0.07742   0.07087  -0.0102   0.0262   1.0000
  10.500   0.8749   0.08605   0.07970  -0.0155   0.0282   1.0000
  10.750   0.8540   0.09490   0.08860  -0.0217   0.0304   1.0000
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