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HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/9 AIRFOIL (hq159-il)
Reynolds number: 200,000
Max Cl/Cd: 68.64 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq159-il-200000.txt
Download as CSV file: xf-hq159-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4987   0.10622   0.10261  -0.0150   1.0000   0.0396
  -9.750  -0.5026   0.10165   0.09810  -0.0198   1.0000   0.0414
  -8.000  -0.5446   0.06609   0.06265  -0.0434   1.0000   0.0436
  -7.750  -0.5373   0.06318   0.05979  -0.0426   1.0000   0.0448
  -7.500  -0.5344   0.05984   0.05641  -0.0428   1.0000   0.0458
  -7.250  -0.5313   0.05642   0.05292  -0.0430   1.0000   0.0475
  -7.000  -0.5281   0.05277   0.04914  -0.0430   1.0000   0.0493
  -5.750  -0.4946   0.03514   0.02971  -0.0343   1.0000   0.0421
  -5.500  -0.4840   0.02773   0.02166  -0.0317   1.0000   0.0299
  -5.250  -0.4694   0.02441   0.01797  -0.0299   1.0000   0.0290
  -5.000  -0.4515   0.02205   0.01523  -0.0282   1.0000   0.0289
  -4.750  -0.4318   0.02024   0.01312  -0.0267   1.0000   0.0293
  -4.500  -0.4114   0.01905   0.01165  -0.0253   1.0000   0.0304
  -4.250  -0.3781   0.01690   0.00943  -0.0268   0.9970   0.0352
  -4.000  -0.3402   0.01598   0.00838  -0.0287   0.9929   0.0419
  -3.750  -0.3061   0.01446   0.00690  -0.0303   0.9883   0.0574
  -3.500  -0.2684   0.01355   0.00608  -0.0327   0.9838   0.0866
  -3.250  -0.2333   0.01272   0.00543  -0.0347   0.9785   0.1352
  -3.000  -0.2048   0.01072   0.00502  -0.0364   0.9729   0.4688
  -2.750  -0.1687   0.01038   0.00514  -0.0379   0.9687   0.6249
  -2.500  -0.1378   0.01033   0.00514  -0.0382   0.9610   0.6784
  -2.250  -0.1000   0.01032   0.00514  -0.0399   0.9565   0.7230
  -2.000  -0.0676   0.01032   0.00515  -0.0404   0.9497   0.7566
  -1.750  -0.0312   0.01030   0.00513  -0.0417   0.9444   0.7860
  -1.500   0.0030   0.01026   0.00511  -0.0424   0.9393   0.8167
  -1.250   0.0327   0.01017   0.00502  -0.0422   0.9315   0.8390
  -1.000   0.0666   0.01004   0.00489  -0.0429   0.9260   0.8592
  -0.750   0.0932   0.00993   0.00480  -0.0420   0.9174   0.8803
  -0.500   0.1230   0.00981   0.00468  -0.0419   0.9106   0.8999
  -0.250   0.1542   0.00971   0.00457  -0.0421   0.9028   0.9156
   0.000   0.1883   0.00960   0.00445  -0.0429   0.8940   0.9308
   0.250   0.2268   0.00945   0.00425  -0.0445   0.8852   0.9449
   0.500   0.2660   0.00932   0.00410  -0.0465   0.8737   0.9590
   0.750   0.3071   0.00921   0.00397  -0.0490   0.8618   0.9724
   1.000   0.3499   0.00911   0.00384  -0.0519   0.8499   0.9852
   1.250   0.3936   0.00903   0.00375  -0.0553   0.8389   0.9993
   1.500   0.4107   0.00902   0.00370  -0.0535   0.8259   1.0000
   1.750   0.4304   0.00905   0.00368  -0.0520   0.8131   1.0000
   2.000   0.4531   0.00908   0.00368  -0.0509   0.7999   1.0000
   2.250   0.4771   0.00912   0.00372  -0.0501   0.7861   1.0000
   2.500   0.5017   0.00916   0.00374  -0.0492   0.7711   1.0000
   2.750   0.5266   0.00921   0.00378  -0.0484   0.7550   1.0000
   3.000   0.5517   0.00925   0.00381  -0.0476   0.7372   1.0000
   3.250   0.5766   0.00930   0.00383  -0.0467   0.7163   1.0000
   3.500   0.6014   0.00935   0.00387  -0.0457   0.6915   1.0000
   3.750   0.6258   0.00944   0.00392  -0.0447   0.6606   1.0000
   4.000   0.6496   0.00958   0.00397  -0.0436   0.6223   1.0000
   4.250   0.6727   0.00980   0.00406  -0.0424   0.5718   1.0000
   4.500   0.6945   0.01020   0.00424  -0.0411   0.5063   1.0000
   4.750   0.7149   0.01080   0.00450  -0.0397   0.4323   1.0000
   5.000   0.7355   0.01147   0.00488  -0.0385   0.3648   1.0000
   5.250   0.7568   0.01215   0.00532  -0.0376   0.3110   1.0000
   5.500   0.7780   0.01286   0.00581  -0.0367   0.2648   1.0000
   5.750   0.7998   0.01354   0.00634  -0.0359   0.2162   1.0000
   6.000   0.8197   0.01448   0.00698  -0.0349   0.1477   1.0000
   6.250   0.8376   0.01576   0.00790  -0.0336   0.1020   1.0000
   6.500   0.8578   0.01679   0.00889  -0.0324   0.0829   1.0000
   6.750   0.8774   0.01792   0.01001  -0.0312   0.0722   1.0000
   7.000   0.8962   0.01923   0.01126  -0.0300   0.0627   1.0000
   7.250   0.9189   0.01976   0.01188  -0.0294   0.0524   1.0000
   7.500   0.9401   0.02066   0.01288  -0.0284   0.0433   1.0000
   7.750   0.9586   0.02208   0.01443  -0.0271   0.0331   1.0000
   8.000   0.9760   0.02377   0.01617  -0.0257   0.0254   1.0000
   8.250   0.9954   0.02528   0.01780  -0.0245   0.0219   1.0000
   8.500   1.0130   0.02744   0.02005  -0.0232   0.0198   1.0000
   8.750   1.0284   0.03154   0.02450  -0.0217   0.0186   1.0000
   9.000   1.0442   0.03409   0.02741  -0.0201   0.0182   1.0000
   9.250   1.0558   0.03721   0.03093  -0.0182   0.0180   1.0000
   9.500   1.0620   0.04081   0.03496  -0.0161   0.0180   1.0000
   9.750   1.0623   0.04476   0.03931  -0.0136   0.0181   1.0000
  10.000   1.0565   0.04890   0.04382  -0.0111   0.0183   1.0000
  10.250   1.0442   0.05311   0.04833  -0.0083   0.0185   1.0000
  10.500   1.0285   0.05779   0.05323  -0.0057   0.0189   1.0000
  10.750   1.0191   0.06031   0.05594  -0.0038   0.0190   1.0000
  11.000   1.0100   0.06293   0.05874  -0.0028   0.0193   1.0000
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