Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

HQ 3.0/8 AIRFOIL (hq308-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 3.0/8 AIRFOIL (hq308-il)
Reynolds number: 200,000
Max Cl/Cd: 79.56 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq308-il-200000-n5.txt
Download as CSV file: xf-hq308-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.0/8 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3681   0.08788   0.08458  -0.0349   1.0000   0.0188
  -7.500  -0.3759   0.08567   0.08246  -0.0340   1.0000   0.0189
  -7.250  -0.3848   0.08350   0.08036  -0.0330   1.0000   0.0189
  -7.000  -0.3841   0.08030   0.07721  -0.0349   0.9987   0.0189
  -6.750  -0.3587   0.07403   0.07090  -0.0452   0.9924   0.0190
  -6.500  -0.3325   0.06796   0.06471  -0.0542   0.9868   0.0190
  -6.000  -0.2929   0.05330   0.04983  -0.0672   0.9754   0.0097
  -5.750  -0.2665   0.04749   0.04381  -0.0734   0.9692   0.0093
  -5.500  -0.2364   0.04202   0.03808  -0.0788   0.9642   0.0090
  -5.250  -0.2022   0.03656   0.03225  -0.0836   0.9610   0.0089
  -5.000  -0.1734   0.03238   0.02770  -0.0854   0.9545   0.0097
  -4.750  -0.1419   0.02775   0.02255  -0.0881   0.9501   0.0118
  -4.500  -0.1083   0.02409   0.01839  -0.0899   0.9469   0.0117
  -4.250  -0.0802   0.02126   0.01507  -0.0901   0.9405   0.0115
  -4.000  -0.0481   0.01886   0.01222  -0.0909   0.9366   0.0115
  -3.750  -0.0161   0.01695   0.00996  -0.0915   0.9330   0.0118
  -3.500   0.0114   0.01554   0.00835  -0.0913   0.9269   0.0124
  -3.250   0.0424   0.01441   0.00704  -0.0918   0.9226   0.0136
  -3.000   0.0715   0.01382   0.00636  -0.0921   0.9173   0.0170
  -2.750   0.1003   0.01304   0.00545  -0.0921   0.9115   0.0187
  -2.500   0.1313   0.01225   0.00451  -0.0927   0.9073   0.0213
  -2.250   0.1586   0.01182   0.00399  -0.0925   0.9001   0.0307
  -2.000   0.1884   0.01137   0.00363  -0.0929   0.8949   0.0702
  -1.750   0.2152   0.01066   0.00336  -0.0932   0.8879   0.2053
  -1.500   0.2390   0.00946   0.00339  -0.0929   0.8817   0.5802
  -1.250   0.2639   0.00928   0.00341  -0.0920   0.8743   0.6687
  -1.000   0.2885   0.00912   0.00339  -0.0907   0.8675   0.7431
  -0.750   0.3115   0.00898   0.00335  -0.0892   0.8584   0.7918
  -0.500   0.3371   0.00886   0.00319  -0.0883   0.8500   0.8196
  -0.250   0.3636   0.00878   0.00306  -0.0877   0.8411   0.8373
   0.000   0.3897   0.00870   0.00297  -0.0871   0.8320   0.8582
   0.250   0.4172   0.00859   0.00286  -0.0867   0.8235   0.8873
   0.500   0.4525   0.00845   0.00271  -0.0880   0.8126   0.9472
   0.750   0.4828   0.00845   0.00262  -0.0885   0.7994   1.0000
   1.000   0.5103   0.00850   0.00259  -0.0884   0.7871   1.0000
   1.250   0.5378   0.00856   0.00259  -0.0882   0.7753   1.0000
   1.500   0.5652   0.00862   0.00261  -0.0881   0.7627   1.0000
   1.750   0.5922   0.00869   0.00261  -0.0878   0.7472   1.0000
   2.000   0.6188   0.00878   0.00262  -0.0874   0.7284   1.0000
   2.250   0.6454   0.00888   0.00266  -0.0871   0.7088   1.0000
   2.500   0.6720   0.00900   0.00275  -0.0867   0.6905   1.0000
   2.750   0.6983   0.00914   0.00284  -0.0863   0.6696   1.0000
   3.000   0.7244   0.00930   0.00294  -0.0859   0.6468   1.0000
   3.250   0.7499   0.00950   0.00306  -0.0854   0.6188   1.0000
   3.500   0.7749   0.00974   0.00320  -0.0847   0.5858   1.0000
   3.750   0.7990   0.01005   0.00343  -0.0840   0.5464   1.0000
   4.000   0.8222   0.01044   0.00365  -0.0831   0.5020   1.0000
   4.250   0.8448   0.01091   0.00393  -0.0821   0.4555   1.0000
   4.500   0.8672   0.01142   0.00427  -0.0813   0.4108   1.0000
   4.750   0.8896   0.01195   0.00465  -0.0804   0.3693   1.0000
   5.000   0.9120   0.01250   0.00506  -0.0796   0.3282   1.0000
   5.250   0.9328   0.01320   0.00560  -0.0787   0.2780   1.0000
   5.500   0.9515   0.01416   0.00614  -0.0775   0.2063   1.0000
   5.750   0.9716   0.01500   0.00669  -0.0766   0.1546   1.0000
   6.000   0.9936   0.01566   0.00724  -0.0759   0.1228   1.0000
   6.250   1.0144   0.01647   0.00786  -0.0751   0.0847   1.0000
   6.500   1.0336   0.01749   0.00864  -0.0740   0.0461   1.0000
   6.750   1.0538   0.01840   0.00948  -0.0730   0.0254   1.0000
   7.000   1.0730   0.01945   0.01047  -0.0718   0.0093   1.0000
   7.250   1.0934   0.02035   0.01152  -0.0706   0.0069   1.0000
   7.500   1.1135   0.02126   0.01261  -0.0694   0.0061   1.0000
   7.750   1.1326   0.02227   0.01381  -0.0682   0.0056   1.0000
   8.000   1.1507   0.02336   0.01509  -0.0667   0.0054   1.0000
   8.250   1.1669   0.02463   0.01656  -0.0651   0.0051   1.0000
   8.500   1.1814   0.02604   0.01815  -0.0633   0.0049   1.0000
   8.750   1.1942   0.02759   0.01990  -0.0613   0.0048   1.0000
   9.000   1.2056   0.02932   0.02182  -0.0591   0.0047   1.0000
   9.250   1.2161   0.03117   0.02387  -0.0569   0.0047   1.0000
   9.500   1.2247   0.03316   0.02606  -0.0545   0.0046   1.0000
   9.750   1.2318   0.03534   0.02848  -0.0520   0.0046   1.0000
  10.000   1.2371   0.03772   0.03111  -0.0495   0.0045   1.0000
  10.250   1.2401   0.04025   0.03391  -0.0469   0.0045   1.0000
  10.500   1.2398   0.04304   0.03707  -0.0443   0.0045   1.0000
  10.750   1.2368   0.04597   0.04029  -0.0419   0.0045   1.0000
  11.000   1.2307   0.04918   0.04378  -0.0397   0.0045   1.0000
  11.250   1.2208   0.05280   0.04768  -0.0381   0.0044   1.0000
  11.500   1.2081   0.05691   0.05206  -0.0371   0.0043   1.0000
  11.750   1.1919   0.06173   0.05714  -0.0370   0.0043   1.0000
  12.000   1.1780   0.06658   0.06223  -0.0377   0.0043   1.0000
  12.250   1.1613   0.07232   0.06820  -0.0396   0.0042   1.0000
  12.500   1.1467   0.07827   0.07436  -0.0422   0.0043   1.0000
  12.750   1.1327   0.08480   0.08108  -0.0458   0.0043   1.0000
  13.000   1.1168   0.09237   0.08884  -0.0505   0.0043   1.0000
  13.250   1.1024   0.10053   0.09717  -0.0558   0.0044   1.0000
  13.500   1.0843   0.11084   0.10769  -0.0625   0.0046   1.0000
  13.750   1.0584   0.12455   0.12151  -0.0710   0.0049   1.0000
  14.000   1.0348   0.13838   0.13544  -0.0789   0.0050   1.0000
<< Back to HQ 3.0/8 AIRFOIL (hq308-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.0/8 AIRFOIL (hq308-il)