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HQ 2.0/8 AIRFOIL (hq208-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.0/8 AIRFOIL (hq208-il)
Reynolds number: 50,000
Max Cl/Cd: 38.51 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq208-il-50000-n5.txt
Download as CSV file: xf-hq208-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4709   0.11712   0.11007  -0.0139   1.0000   0.0938
  -9.500  -0.4698   0.11399   0.10700  -0.0158   1.0000   0.0966
  -9.250  -0.4851   0.11226   0.10544  -0.0211   1.0000   0.0995
  -9.000  -0.4800   0.10788   0.10113  -0.0219   1.0000   0.1003
  -8.750  -0.4703   0.10346   0.09673  -0.0216   1.0000   0.1013
  -8.500  -0.4568   0.09633   0.08952  -0.0239   1.0000   0.0549
  -8.000  -0.4597   0.08745   0.08080  -0.0296   1.0000   0.0442
  -7.750  -0.4590   0.08369   0.07712  -0.0307   1.0000   0.0427
  -7.500  -0.4581   0.07942   0.07292  -0.0336   1.0000   0.0414
  -7.250  -0.4574   0.07482   0.06836  -0.0368   1.0000   0.0401
  -7.000  -0.4561   0.07030   0.06383  -0.0395   1.0000   0.0388
  -6.750  -0.4541   0.06561   0.05908  -0.0420   1.0000   0.0375
  -6.500  -0.4507   0.06082   0.05415  -0.0441   1.0000   0.0362
  -6.250  -0.4445   0.05576   0.04870  -0.0460   1.0000   0.0344
  -5.750  -0.4211   0.04856   0.04081  -0.0461   1.0000   0.0334
  -5.500  -0.4077   0.04492   0.03683  -0.0458   1.0000   0.0334
  -5.250  -0.3932   0.04134   0.03293  -0.0455   1.0000   0.0340
  -5.000  -0.3788   0.03872   0.03024  -0.0450   1.0000   0.0370
  -4.750  -0.3605   0.03625   0.02741  -0.0444   1.0000   0.0400
  -4.500  -0.3398   0.03357   0.02415  -0.0437   1.0000   0.0417
  -4.250  -0.3177   0.03099   0.02108  -0.0427   1.0000   0.0427
  -4.000  -0.2946   0.02874   0.01835  -0.0416   1.0000   0.0442
  -3.750  -0.2713   0.02687   0.01603  -0.0402   1.0000   0.0462
  -3.500  -0.2502   0.02509   0.01424  -0.0389   1.0000   0.0502
  -3.250  -0.2286   0.02398   0.01290  -0.0376   1.0000   0.0604
  -3.000  -0.2069   0.02274   0.01152  -0.0363   1.0000   0.0717
  -2.750  -0.1849   0.02156   0.01036  -0.0357   1.0000   0.0953
  -2.500  -0.1607   0.02014   0.00914  -0.0355   1.0000   0.1456
  -2.250  -0.1412   0.01751   0.00855  -0.0347   1.0000   0.5412
  -2.000  -0.1334   0.01715   0.00866  -0.0292   1.0000   0.7380
  -1.750  -0.1229   0.01676   0.00848  -0.0240   1.0000   0.8709
  -1.500  -0.0850   0.01650   0.00791  -0.0263   1.0000   1.0000
  -1.250  -0.0627   0.01661   0.00765  -0.0263   1.0000   1.0000
  -1.000  -0.0378   0.01679   0.00749  -0.0268   0.9987   1.0000
  -0.750   0.0010   0.01710   0.00740  -0.0298   0.9905   1.0000
  -0.500   0.0394   0.01743   0.00743  -0.0327   0.9825   1.0000
  -0.250   0.0780   0.01776   0.00753  -0.0356   0.9742   1.0000
   0.000   0.1143   0.01806   0.00763  -0.0380   0.9649   1.0000
   0.250   0.1518   0.01838   0.00777  -0.0405   0.9560   1.0000
   0.500   0.1906   0.01871   0.00797  -0.0433   0.9471   1.0000
   0.750   0.2255   0.01899   0.00817  -0.0452   0.9366   1.0000
   1.000   0.2615   0.01927   0.00840  -0.0473   0.9262   1.0000
   1.250   0.2995   0.01954   0.00864  -0.0496   0.9162   1.0000
   1.500   0.3397   0.01978   0.00889  -0.0522   0.9063   1.0000
   1.750   0.3754   0.02000   0.00915  -0.0540   0.8944   1.0000
   2.000   0.4134   0.02012   0.00937  -0.0557   0.8807   1.0000
   2.250   0.4530   0.02006   0.00939  -0.0573   0.8640   1.0000
   2.500   0.4862   0.02001   0.00943  -0.0577   0.8441   1.0000
   2.750   0.5207   0.01993   0.00946  -0.0581   0.8263   1.0000
   3.000   0.5512   0.01994   0.00968  -0.0580   0.8087   1.0000
   3.250   0.5792   0.01998   0.00987  -0.0574   0.7893   1.0000
   3.500   0.6100   0.01991   0.00995  -0.0571   0.7704   1.0000
   3.750   0.6361   0.01994   0.01016  -0.0561   0.7471   1.0000
   4.000   0.6630   0.01992   0.01032  -0.0550   0.7220   1.0000
   4.250   0.6897   0.01988   0.01056  -0.0538   0.6932   1.0000
   4.500   0.7160   0.01986   0.01069  -0.0524   0.6593   1.0000
   4.750   0.7410   0.01990   0.01085  -0.0508   0.6173   1.0000
   5.000   0.7648   0.02004   0.01100  -0.0489   0.5643   1.0000
   5.250   0.7859   0.02041   0.01119  -0.0466   0.4954   1.0000
   5.500   0.8037   0.02116   0.01156  -0.0443   0.4174   1.0000
   5.750   0.8201   0.02222   0.01239  -0.0423   0.3437   1.0000
   6.000   0.8361   0.02344   0.01331  -0.0407   0.2809   1.0000
   6.250   0.8527   0.02477   0.01447  -0.0393   0.2210   1.0000
   6.500   0.8696   0.02629   0.01584  -0.0380   0.1674   1.0000
   6.750   0.8864   0.02797   0.01729  -0.0368   0.1276   1.0000
   7.000   0.9044   0.02973   0.01895  -0.0356   0.1056   1.0000
   7.250   0.9248   0.03164   0.02097  -0.0344   0.0919   1.0000
   7.500   0.9442   0.03353   0.02303  -0.0333   0.0750   1.0000
   7.750   0.9661   0.03613   0.02579  -0.0321   0.0642   1.0000
   8.000   0.9817   0.03841   0.02829  -0.0309   0.0497   1.0000
   8.250   0.9956   0.04082   0.03105  -0.0296   0.0392   1.0000
   8.500   1.0127   0.04406   0.03466  -0.0283   0.0346   1.0000
   8.750   1.0251   0.04762   0.03838  -0.0272   0.0318   1.0000
   9.000   1.0339   0.05145   0.04278  -0.0254   0.0301   1.0000
   9.250   1.0367   0.05536   0.04727  -0.0236   0.0284   1.0000
   9.500   1.0343   0.05928   0.05164  -0.0219   0.0272   1.0000
   9.750   1.0265   0.06323   0.05595  -0.0202   0.0264   1.0000
  10.000   1.0134   0.06707   0.06008  -0.0186   0.0261   1.0000
  10.250   0.9971   0.07135   0.06459  -0.0180   0.0261   1.0000
  10.500   0.9783   0.07638   0.06983  -0.0189   0.0263   1.0000
  10.750   0.9586   0.08232   0.07595  -0.0214   0.0271   1.0000
  11.000   0.9400   0.08897   0.08271  -0.0252   0.0279   1.0000
  11.250   0.9225   0.09650   0.09031  -0.0301   0.0288   1.0000
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