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HQ 1.5/10 AIRFOIL (hq1510-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/10 AIRFOIL (hq1510-il)
Reynolds number: 50,000
Max Cl/Cd: 34.29 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1510-il-50000.txt
Download as CSV file: xf-hq1510-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/10 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4554   0.10107   0.09425  -0.0060   1.0000   0.3082
  -8.500  -0.4488   0.09775   0.09097  -0.0054   1.0000   0.3244
  -8.250  -0.4560   0.09571   0.08903  -0.0049   1.0000   0.3448
  -8.000  -0.4374   0.09149   0.08480  -0.0036   1.0000   0.3649
  -7.750  -0.4314   0.08826   0.08162  -0.0028   1.0000   0.3826
  -7.500  -0.4278   0.08493   0.07835  -0.0024   1.0000   0.3947
  -7.250  -0.5311   0.06929   0.06302  -0.0327   1.0000   0.1841
  -7.000  -0.5507   0.06035   0.05374  -0.0393   1.0000   0.1447
  -6.750  -0.5519   0.05451   0.04751  -0.0403   1.0000   0.1304
  -6.500  -0.5513   0.04930   0.04152  -0.0402   1.0000   0.1197
  -6.250  -0.5393   0.04530   0.03732  -0.0390   1.0000   0.1175
  -6.000  -0.5276   0.04168   0.03325  -0.0377   1.0000   0.1172
  -5.750  -0.5147   0.03865   0.02946  -0.0362   1.0000   0.1200
  -5.500  -0.4982   0.03570   0.02651  -0.0348   1.0000   0.1259
  -5.250  -0.4796   0.03296   0.02330  -0.0333   1.0000   0.1303
  -5.000  -0.4613   0.03074   0.02086  -0.0318   1.0000   0.1427
  -4.750  -0.4406   0.02853   0.01841  -0.0303   1.0000   0.1548
  -4.500  -0.4203   0.02685   0.01652  -0.0288   1.0000   0.1747
  -4.250  -0.3986   0.02506   0.01479  -0.0273   1.0000   0.1947
  -4.000  -0.3783   0.02358   0.01343  -0.0257   1.0000   0.2260
  -3.750  -0.3571   0.02188   0.01194  -0.0242   1.0000   0.2729
  -3.500  -0.3448   0.01929   0.01101  -0.0213   1.0000   0.4482
  -3.250  -0.3520   0.01935   0.01217  -0.0104   1.0000   0.7015
  -3.000  -0.3532   0.01975   0.01258  -0.0016   1.0000   0.7838
  -2.750  -0.3494   0.01988   0.01265   0.0060   1.0000   0.8470
  -2.500  -0.0523   0.02079   0.01176  -0.0336   1.0000   1.0000
  -2.250  -0.0607   0.02056   0.01151  -0.0298   1.0000   1.0000
  -2.000  -0.0715   0.02034   0.01127  -0.0257   1.0000   1.0000
  -1.750  -0.0839   0.02009   0.01099  -0.0214   1.0000   1.0000
  -1.500  -0.0974   0.01979   0.01066  -0.0169   1.0000   1.0000
  -1.250  -0.1101   0.01944   0.01027  -0.0125   1.0000   1.0000
  -1.000  -0.1156   0.01917   0.00989  -0.0090   1.0000   1.0000
  -0.750  -0.1081   0.01911   0.00966  -0.0075   1.0000   1.0000
  -0.500  -0.0933   0.01921   0.00958  -0.0069   1.0000   1.0000
  -0.250  -0.0756   0.01942   0.00959  -0.0068   1.0000   1.0000
   0.000  -0.0568   0.01970   0.00971  -0.0067   1.0000   1.0000
   0.250  -0.0374   0.02004   0.00991  -0.0067   1.0000   1.0000
   0.500  -0.0179   0.02044   0.01018  -0.0067   1.0000   1.0000
   0.750   0.0016   0.02088   0.01051  -0.0068   1.0000   1.0000
   1.000   0.0210   0.02137   0.01089  -0.0068   1.0000   1.0000
   1.250   0.0402   0.02191   0.01136  -0.0069   1.0000   1.0000
   1.500   0.0591   0.02250   0.01189  -0.0069   1.0000   1.0000
   1.750   0.0778   0.02315   0.01250  -0.0071   1.0000   1.0000
   2.000   0.1120   0.02422   0.01354  -0.0102   0.9930   1.0000
   2.250   0.1662   0.02563   0.01498  -0.0169   0.9738   1.0000
   2.500   0.2255   0.02695   0.01639  -0.0240   0.9499   1.0000
   2.750   0.2861   0.02804   0.01758  -0.0308   0.9253   1.0000
   3.000   0.3344   0.02883   0.01849  -0.0351   0.9019   1.0000
   3.250   0.3809   0.02951   0.01935  -0.0387   0.8787   1.0000
   3.500   0.4390   0.02996   0.02002  -0.0437   0.8559   1.0000
   3.750   0.4832   0.03025   0.02052  -0.0460   0.8304   1.0000
   4.000   0.5331   0.03024   0.02081  -0.0487   0.8040   1.0000
   4.250   0.5840   0.02983   0.02069  -0.0505   0.7764   1.0000
   4.500   0.6320   0.02903   0.02022  -0.0510   0.7473   1.0000
   4.750   0.6788   0.02779   0.01926  -0.0503   0.7164   1.0000
   5.000   0.7130   0.02682   0.01849  -0.0480   0.6794   1.0000
   5.250   0.7475   0.02559   0.01739  -0.0451   0.6389   1.0000
   5.500   0.7768   0.02459   0.01645  -0.0419   0.5927   1.0000
   5.750   0.8014   0.02403   0.01576  -0.0385   0.5413   1.0000
   6.000   0.8225   0.02399   0.01549  -0.0353   0.4859   1.0000
   6.250   0.8405   0.02453   0.01573  -0.0323   0.4267   1.0000
   6.500   0.8567   0.02560   0.01637  -0.0295   0.3640   1.0000
   6.750   0.8720   0.02733   0.01775  -0.0270   0.3004   1.0000
   7.000   0.8899   0.02940   0.01938  -0.0251   0.2484   1.0000
   7.250   0.9121   0.03170   0.02163  -0.0239   0.2144   1.0000
   7.500   0.9345   0.03384   0.02374  -0.0228   0.1909   1.0000
   7.750   0.9553   0.03591   0.02596  -0.0217   0.1726   1.0000
   8.000   0.9776   0.03871   0.02902  -0.0207   0.1615   1.0000
   8.250   0.9973   0.04137   0.03186  -0.0196   0.1506   1.0000
   8.500   1.0176   0.04401   0.03459  -0.0187   0.1397   1.0000
   8.750   1.0254   0.04794   0.03930  -0.0166   0.1359   1.0000
   9.000   1.0317   0.05204   0.04394  -0.0147   0.1324   1.0000
   9.250   1.0441   0.05547   0.04746  -0.0135   0.1251   1.0000
   9.500   1.0356   0.05991   0.05248  -0.0112   0.1226   1.0000
   9.750   1.0243   0.06478   0.05777  -0.0094   0.1221   1.0000
  10.000   1.0084   0.06983   0.06312  -0.0080   0.1229   1.0000
  10.250   0.9895   0.07486   0.06836  -0.0069   0.1240   1.0000
  10.500   0.9699   0.07982   0.07342  -0.0062   0.1252   1.0000
  10.750   0.9549   0.08537   0.07903  -0.0066   0.1262   1.0000
  11.000   0.7550   0.13147   0.12448  -0.0516   0.3207   1.0000
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