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HQ 2.0/8 AIRFOIL (hq208-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.0/8 AIRFOIL (hq208-il)
Reynolds number: 200,000
Max Cl/Cd: 68.64 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq208-il-200000-n5.txt
Download as CSV file: xf-hq208-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4532   0.07872   0.07533  -0.0293   1.0000   0.0116
  -7.750  -0.4585   0.07483   0.07152  -0.0311   1.0000   0.0112
  -7.500  -0.4612   0.07028   0.06702  -0.0350   1.0000   0.0110
  -7.250  -0.4616   0.06507   0.06180  -0.0396   1.0000   0.0107
  -7.000  -0.4607   0.06043   0.05711  -0.0423   1.0000   0.0105
  -6.750  -0.4590   0.05589   0.05250  -0.0439   1.0000   0.0102
  -6.500  -0.4562   0.05153   0.04802  -0.0446   1.0000   0.0099
  -6.250  -0.4527   0.04742   0.04376  -0.0444   1.0000   0.0096
  -6.000  -0.4340   0.04184   0.03787  -0.0470   0.9965   0.0092
  -5.750  -0.4050   0.03571   0.03128  -0.0507   0.9905   0.0089
  -5.500  -0.3736   0.02989   0.02487  -0.0535   0.9857   0.0086
  -5.250  -0.3437   0.02542   0.01979  -0.0549   0.9801   0.0087
  -5.000  -0.3113   0.02206   0.01586  -0.0562   0.9758   0.0092
  -4.750  -0.2773   0.01997   0.01334  -0.0576   0.9726   0.0106
  -4.500  -0.2459   0.01915   0.01220  -0.0582   0.9668   0.0124
  -4.250  -0.2148   0.01638   0.00908  -0.0589   0.9627   0.0133
  -4.000  -0.1827   0.01491   0.00747  -0.0599   0.9586   0.0146
  -3.750  -0.1533   0.01398   0.00644  -0.0603   0.9521   0.0166
  -3.500  -0.1202   0.01339   0.00574  -0.0614   0.9475   0.0209
  -3.250  -0.0907   0.01255   0.00480  -0.0618   0.9408   0.0262
  -3.000  -0.0590   0.01193   0.00414  -0.0626   0.9351   0.0434
  -2.750  -0.0291   0.01141   0.00372  -0.0633   0.9287   0.0851
  -2.500  -0.0009   0.01041   0.00337  -0.0641   0.9221   0.2445
  -2.250   0.0251   0.00946   0.00318  -0.0643   0.9151   0.4683
  -2.000   0.0525   0.00913   0.00312  -0.0640   0.9082   0.5745
  -1.750   0.0793   0.00896   0.00309  -0.0635   0.9008   0.6452
  -1.500   0.1062   0.00884   0.00303  -0.0629   0.8937   0.6996
  -1.250   0.1307   0.00874   0.00304  -0.0618   0.8851   0.7501
  -1.000   0.1566   0.00866   0.00298  -0.0609   0.8777   0.7854
  -0.750   0.1821   0.00859   0.00291  -0.0601   0.8685   0.8077
  -0.500   0.2087   0.00854   0.00281  -0.0595   0.8596   0.8234
  -0.250   0.2354   0.00847   0.00272  -0.0590   0.8505   0.8389
   0.000   0.2616   0.00841   0.00263  -0.0583   0.8386   0.8562
   0.250   0.2878   0.00834   0.00254  -0.0576   0.8248   0.8770
   0.500   0.3160   0.00826   0.00246  -0.0573   0.8108   0.9028
   0.750   0.3503   0.00819   0.00239  -0.0584   0.7961   0.9380
   1.000   0.3876   0.00815   0.00230  -0.0603   0.7803   1.0000
   1.250   0.4143   0.00821   0.00230  -0.0601   0.7659   1.0000
   1.500   0.4411   0.00828   0.00233  -0.0598   0.7519   1.0000
   1.750   0.4678   0.00836   0.00236  -0.0595   0.7368   1.0000
   2.000   0.4945   0.00845   0.00241  -0.0592   0.7207   1.0000
   2.250   0.5210   0.00855   0.00246  -0.0589   0.7028   1.0000
   2.500   0.5474   0.00867   0.00254  -0.0585   0.6823   1.0000
   2.750   0.5735   0.00881   0.00266  -0.0581   0.6596   1.0000
   3.000   0.5993   0.00898   0.00277  -0.0576   0.6325   1.0000
   3.250   0.6247   0.00919   0.00290  -0.0570   0.5994   1.0000
   3.500   0.6493   0.00946   0.00304  -0.0563   0.5577   1.0000
   3.750   0.6723   0.00988   0.00322  -0.0553   0.4979   1.0000
   4.000   0.6926   0.01058   0.00354  -0.0540   0.4085   1.0000
   4.250   0.7134   0.01134   0.00390  -0.0530   0.3272   1.0000
   4.750   0.7605   0.01240   0.00469  -0.0519   0.2526   1.0000
   5.000   0.7845   0.01291   0.00510  -0.0515   0.2186   1.0000
   5.250   0.8080   0.01346   0.00553  -0.0510   0.1782   1.0000
   5.500   0.8308   0.01413   0.00608  -0.0504   0.1311   1.0000
   5.750   0.8519   0.01506   0.00673  -0.0496   0.0826   1.0000
   6.000   0.8729   0.01601   0.00741  -0.0489   0.0470   1.0000
   6.250   0.8960   0.01665   0.00802  -0.0484   0.0339   1.0000
   6.500   0.9198   0.01720   0.00869  -0.0478   0.0274   1.0000
   6.750   0.9423   0.01795   0.00948  -0.0471   0.0169   1.0000
   7.000   0.9636   0.01891   0.01050  -0.0461   0.0110   1.0000
   7.250   0.9849   0.01987   0.01162  -0.0450   0.0090   1.0000
   7.500   1.0050   0.02096   0.01291  -0.0439   0.0074   1.0000
   7.750   1.0233   0.02229   0.01439  -0.0427   0.0063   1.0000
   8.000   1.0428   0.02339   0.01566  -0.0416   0.0054   1.0000
   8.250   1.0606   0.02476   0.01722  -0.0403   0.0050   1.0000
   8.500   1.0772   0.02634   0.01900  -0.0389   0.0046   1.0000
   8.750   1.0926   0.02809   0.02099  -0.0373   0.0045   1.0000
   9.000   1.1069   0.03003   0.02317  -0.0357   0.0043   1.0000
   9.250   1.1193   0.03221   0.02563  -0.0340   0.0041   1.0000
   9.500   1.1295   0.03461   0.02834  -0.0321   0.0041   1.0000
   9.750   1.1359   0.03733   0.03141  -0.0299   0.0040   1.0000
  10.000   1.1385   0.04017   0.03459  -0.0276   0.0040   1.0000
  10.250   1.1345   0.04304   0.03779  -0.0246   0.0039   1.0000
  10.500   1.1253   0.04617   0.04122  -0.0217   0.0039   1.0000
  10.750   1.1134   0.04965   0.04499  -0.0196   0.0039   1.0000
  11.000   1.0998   0.05354   0.04916  -0.0184   0.0039   1.0000
  11.250   1.0835   0.05810   0.05397  -0.0185   0.0039   1.0000
  11.500   1.0658   0.06347   0.05958  -0.0200   0.0040   1.0000
  11.750   1.0468   0.06980   0.06614  -0.0232   0.0040   1.0000
  12.000   1.0270   0.07747   0.07401  -0.0281   0.0040   1.0000
  12.250   1.0058   0.08713   0.08385  -0.0349   0.0041   1.0000
  12.500   0.9810   0.09972   0.09660  -0.0435   0.0043   1.0000
  12.750   0.9504   0.11507   0.11202  -0.0522   0.0045   1.0000
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