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HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il)
Reynolds number: 200,000
Max Cl/Cd: 67.32 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq159b-il-200000.txt
Download as CSV file: xf-hq159b-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/9 B AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4229   0.10563   0.10227  -0.0147   1.0000   0.0442
 -10.250  -0.4209   0.10170   0.09834  -0.0158   1.0000   0.0455
 -10.000  -0.5198   0.10282   0.09926  -0.0164   1.0000   0.0425
  -9.750  -0.5147   0.09946   0.09591  -0.0171   1.0000   0.0434
  -9.500  -0.5114   0.09586   0.09234  -0.0183   1.0000   0.0446
  -9.250  -0.5102   0.09200   0.08850  -0.0201   1.0000   0.0458
  -9.000  -0.5110   0.08778   0.08432  -0.0224   1.0000   0.0473
  -8.750  -0.5139   0.08324   0.07983  -0.0254   1.0000   0.0485
  -8.500  -0.5209   0.07799   0.07464  -0.0299   1.0000   0.0497
  -8.250  -0.5331   0.07233   0.06904  -0.0355   1.0000   0.0499
  -8.000  -0.5459   0.06665   0.06323  -0.0411   1.0000   0.0512
  -7.750  -0.5576   0.06346   0.05962  -0.0437   1.0000   0.0524
  -7.500  -0.5566   0.06027   0.05614  -0.0436   1.0000   0.0526
  -7.250  -0.5555   0.05212   0.04816  -0.0442   1.0000   0.0543
  -7.000  -0.5457   0.04901   0.04505  -0.0434   1.0000   0.0555
  -6.750  -0.5368   0.04611   0.04206  -0.0424   1.0000   0.0572
  -6.500  -0.5283   0.04315   0.03891  -0.0412   1.0000   0.0594
  -6.250  -0.5201   0.04014   0.03562  -0.0396   1.0000   0.0619
  -6.000  -0.5150   0.03475   0.02961  -0.0372   1.0000   0.0555
  -5.500  -0.4861   0.02268   0.01588  -0.0312   1.0000   0.0319
  -5.250  -0.4690   0.02174   0.01485  -0.0298   1.0000   0.0347
  -5.000  -0.4497   0.02009   0.01292  -0.0282   1.0000   0.0368
  -4.750  -0.4296   0.01892   0.01148  -0.0266   1.0000   0.0396
  -4.500  -0.4101   0.01737   0.00989  -0.0255   1.0000   0.0454
  -4.250  -0.3747   0.01657   0.00891  -0.0271   0.9965   0.0532
  -4.000  -0.3383   0.01553   0.00787  -0.0293   0.9925   0.0635
  -3.750  -0.3026   0.01457   0.00693  -0.0312   0.9876   0.0741
  -3.500  -0.2649   0.01373   0.00610  -0.0336   0.9835   0.0867
  -3.250  -0.2303   0.01294   0.00540  -0.0354   0.9778   0.1118
  -3.000  -0.1985   0.01110   0.00477  -0.0376   0.9728   0.3569
  -2.750  -0.1660   0.01032   0.00496  -0.0387   0.9686   0.6157
  -2.500  -0.1348   0.01028   0.00503  -0.0391   0.9610   0.6819
  -2.250  -0.0953   0.01027   0.00503  -0.0411   0.9566   0.7253
  -2.000  -0.0621   0.01026   0.00503  -0.0417   0.9492   0.7589
  -1.750  -0.0248   0.01021   0.00500  -0.0431   0.9438   0.7885
  -1.500   0.0063   0.01017   0.00498  -0.0431   0.9364   0.8153
  -1.250   0.0379   0.01009   0.00491  -0.0433   0.9300   0.8369
  -1.000   0.0666   0.01004   0.00485  -0.0430   0.9230   0.8565
  -0.750   0.0945   0.00995   0.00479  -0.0424   0.9161   0.8753
  -0.500   0.1218   0.00987   0.00471  -0.0418   0.9082   0.8920
  -0.250   0.1513   0.00974   0.00457  -0.0415   0.9004   0.9087
   0.000   0.1792   0.00963   0.00445  -0.0409   0.8901   0.9260
   0.250   0.2124   0.00950   0.00430  -0.0415   0.8805   0.9407
   0.500   0.2502   0.00935   0.00412  -0.0430   0.8713   0.9533
   0.750   0.2896   0.00924   0.00399  -0.0451   0.8593   0.9656
   1.000   0.3308   0.00915   0.00388  -0.0477   0.8473   0.9776
   1.250   0.3737   0.00908   0.00381  -0.0509   0.8370   0.9896
   1.500   0.4088   0.00902   0.00374  -0.0526   0.8261   1.0000
   1.750   0.4261   0.00903   0.00371  -0.0508   0.8135   1.0000
   2.000   0.4474   0.00907   0.00374  -0.0496   0.8008   1.0000
   2.250   0.4712   0.00912   0.00377  -0.0487   0.7876   1.0000
   2.500   0.4960   0.00917   0.00382  -0.0480   0.7741   1.0000
   2.750   0.5213   0.00922   0.00386  -0.0473   0.7593   1.0000
   3.000   0.5462   0.00925   0.00390  -0.0464   0.7406   1.0000
   3.250   0.5709   0.00927   0.00389  -0.0454   0.7169   1.0000
   3.500   0.5953   0.00932   0.00388  -0.0443   0.6874   1.0000
   3.750   0.6192   0.00940   0.00389  -0.0431   0.6486   1.0000
   4.000   0.6426   0.00957   0.00395  -0.0419   0.5993   1.0000
   4.250   0.6651   0.00988   0.00404  -0.0407   0.5343   1.0000
   4.500   0.6858   0.01045   0.00425  -0.0393   0.4592   1.0000
   4.750   0.7065   0.01113   0.00460  -0.0381   0.3842   1.0000
   5.000   0.7271   0.01187   0.00500  -0.0371   0.3144   1.0000
   5.250   0.7487   0.01258   0.00550  -0.0363   0.2650   1.0000
   5.500   0.7706   0.01329   0.00603  -0.0355   0.2104   1.0000
   5.750   0.7850   0.01512   0.00704  -0.0339   0.0871   1.0000
   6.000   0.8053   0.01624   0.00808  -0.0327   0.0651   1.0000
   6.250   0.8262   0.01724   0.00911  -0.0316   0.0569   1.0000
   6.500   0.8464   0.01834   0.01019  -0.0305   0.0507   1.0000
   6.750   0.8674   0.01944   0.01140  -0.0294   0.0461   1.0000
   7.000   0.8889   0.02056   0.01257  -0.0284   0.0419   1.0000
   7.250   0.9085   0.02257   0.01457  -0.0273   0.0375   1.0000
   7.500   0.9318   0.02398   0.01617  -0.0264   0.0350   1.0000
   7.750   0.9548   0.02580   0.01815  -0.0255   0.0326   1.0000
   8.000   0.9757   0.02753   0.01994  -0.0248   0.0295   1.0000
   8.250   0.9934   0.03207   0.02486  -0.0237   0.0277   1.0000
   8.500   1.0111   0.03423   0.02741  -0.0222   0.0271   1.0000
   8.750   1.0249   0.03748   0.03110  -0.0204   0.0268   1.0000
   9.000   1.0334   0.04149   0.03555  -0.0184   0.0270   1.0000
   9.250   1.0401   0.04715   0.04154  -0.0168   0.0281   1.0000
   9.500   1.0509   0.04906   0.04376  -0.0148   0.0287   1.0000
  13.250   0.6647   0.15000   0.14698  -0.0497   0.0557   1.0000
  13.500   0.6658   0.15324   0.15024  -0.0513   0.0535   1.0000
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