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HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il)
Reynolds number: 500,000
Max Cl/Cd: 108.1 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq258-il-500000.txt
Download as CSV file: xf-hq258-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3217   0.08373   0.08162  -0.0320   1.0000   0.0138
  -8.500  -0.3209   0.07993   0.07784  -0.0330   1.0000   0.0139
  -8.250  -0.3209   0.07617   0.07411  -0.0339   1.0000   0.0139
  -8.000  -0.3221   0.07244   0.07041  -0.0348   1.0000   0.0139
  -7.750  -0.3252   0.06882   0.06683  -0.0354   1.0000   0.0139
  -7.500  -0.3312   0.06547   0.06352  -0.0355   1.0000   0.0140
  -7.250  -0.3453   0.06317   0.06128  -0.0337   1.0000   0.0140
  -7.000  -0.3580   0.06048   0.05865  -0.0336   0.9983   0.0140
  -6.750  -0.3409   0.05264   0.05075  -0.0459   0.9942   0.0140
  -6.500  -0.3247   0.04609   0.04407  -0.0534   0.9893   0.0140
  -6.250  -0.3045   0.03960   0.03743  -0.0600   0.9853   0.0140
  -6.000  -0.2913   0.02821   0.02574  -0.0686   0.9820   0.0147
  -5.750  -0.2709   0.02325   0.02059  -0.0721   0.9770   0.0155
  -5.500  -0.2442   0.01987   0.01705  -0.0752   0.9736   0.0165
  -5.250  -0.2149   0.01651   0.01347  -0.0781   0.9706   0.0180
  -5.000  -0.1858   0.01350   0.01018  -0.0800   0.9668   0.0200
  -4.750  -0.1534   0.01335   0.00974  -0.0797   0.9608   0.0239
  -4.500  -0.1253   0.01090   0.00695  -0.0806   0.9563   0.0240
  -4.250  -0.1166   0.01651   0.01156  -0.0816   0.9605   0.0120
  -4.000  -0.0843   0.01369   0.00833  -0.0821   0.9569   0.0114
  -3.750  -0.0552   0.01222   0.00666  -0.0820   0.9517   0.0122
  -3.500  -0.0262   0.01165   0.00596  -0.0820   0.9457   0.0137
  -3.250   0.0013   0.01030   0.00446  -0.0819   0.9405   0.0155
  -3.000   0.0273   0.00950   0.00356  -0.0813   0.9328   0.0165
  -2.750   0.0552   0.00896   0.00291  -0.0810   0.9269   0.0202
  -2.500   0.0816   0.00840   0.00242  -0.0806   0.9191   0.0594
  -2.250   0.1082   0.00794   0.00219  -0.0804   0.9124   0.1304
  -2.000   0.1309   0.00656   0.00197  -0.0803   0.9039   0.4917
  -1.750   0.1558   0.00624   0.00195  -0.0796   0.8952   0.6100
  -1.500   0.1822   0.00615   0.00189  -0.0790   0.8874   0.6560
  -1.250   0.2083   0.00607   0.00186  -0.0784   0.8786   0.6930
  -1.000   0.2348   0.00603   0.00183  -0.0780   0.8708   0.7214
  -0.750   0.2610   0.00599   0.00178  -0.0774   0.8625   0.7513
  -0.500   0.2855   0.00592   0.00179  -0.0763   0.8520   0.7931
  -0.250   0.3104   0.00585   0.00175  -0.0754   0.8405   0.8199
   0.000   0.3363   0.00580   0.00169  -0.0747   0.8297   0.8364
   0.250   0.3624   0.00574   0.00164  -0.0741   0.8200   0.8534
   0.500   0.3879   0.00567   0.00160  -0.0735   0.8093   0.8750
   0.750   0.4127   0.00555   0.00156  -0.0725   0.7979   0.9085
   1.000   0.4512   0.00545   0.00149  -0.0747   0.7859   0.9833
   1.250   0.4803   0.00550   0.00147  -0.0750   0.7736   1.0000
   1.500   0.5078   0.00557   0.00149  -0.0749   0.7607   1.0000
   1.750   0.5350   0.00564   0.00151  -0.0748   0.7466   1.0000
   2.000   0.5620   0.00573   0.00153  -0.0746   0.7306   1.0000
   2.250   0.5888   0.00583   0.00157  -0.0743   0.7127   1.0000
   2.500   0.6155   0.00594   0.00162  -0.0741   0.6928   1.0000
   2.750   0.6419   0.00608   0.00172  -0.0738   0.6721   1.0000
   3.000   0.6681   0.00624   0.00181  -0.0734   0.6470   1.0000
   3.250   0.6940   0.00642   0.00191  -0.0730   0.6179   1.0000
   3.500   0.7189   0.00668   0.00203  -0.0724   0.5762   1.0000
   3.750   0.7429   0.00704   0.00220  -0.0717   0.5253   1.0000
   4.000   0.7663   0.00749   0.00242  -0.0710   0.4678   1.0000
   4.250   0.7887   0.00808   0.00274  -0.0702   0.3975   1.0000
   4.500   0.8094   0.00889   0.00310  -0.0692   0.3062   1.0000
   4.750   0.8330   0.00943   0.00343  -0.0687   0.2594   1.0000
   5.000   0.8556   0.01009   0.00379  -0.0681   0.2004   1.0000
   5.250   0.8785   0.01073   0.00418  -0.0675   0.1481   1.0000
   5.500   0.8978   0.01185   0.00480  -0.0664   0.0631   1.0000
   5.750   0.9168   0.01311   0.00574  -0.0651   0.0099   1.0000
   6.000   0.9408   0.01371   0.00648  -0.0643   0.0079   1.0000
   6.250   0.9641   0.01437   0.00728  -0.0634   0.0070   1.0000
   6.500   0.9861   0.01525   0.00829  -0.0623   0.0062   1.0000
   6.750   1.0067   0.01627   0.00945  -0.0611   0.0059   1.0000
   7.000   1.0257   0.01752   0.01086  -0.0595   0.0059   1.0000
   7.250   1.0437   0.01900   0.01249  -0.0577   0.0060   1.0000
   7.500   1.0613   0.02084   0.01452  -0.0559   0.0062   1.0000
   7.750   1.0793   0.02330   0.01722  -0.0540   0.0068   1.0000
   8.000   1.0969   0.02653   0.02075  -0.0523   0.0075   1.0000
  15.000   0.7692   0.18296   0.18122  -0.0884   0.0157   1.0000
  15.250   0.7714   0.18733   0.18559  -0.0906   0.0157   1.0000
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