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

HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il)
Reynolds number: 200,000
Max Cl/Cd: 67.18 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1585-il-200000.txt
Download as CSV file: xf-hq1585-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/8.5 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5269   0.08386   0.08047  -0.0253   1.0000   0.0501
  -8.500  -0.5378   0.07760   0.07427  -0.0328   1.0000   0.0505
  -8.250  -0.5547   0.07205   0.06859  -0.0393   1.0000   0.0509
  -8.000  -0.5602   0.06822   0.06456  -0.0418   1.0000   0.0512
  -7.750  -0.5679   0.06187   0.05801  -0.0439   1.0000   0.0518
  -7.500  -0.5575   0.05644   0.05278  -0.0437   1.0000   0.0534
  -7.250  -0.5466   0.05358   0.04992  -0.0434   1.0000   0.0551
  -7.000  -0.5378   0.05034   0.04659  -0.0436   1.0000   0.0577
  -6.500  -0.5252   0.04222   0.03780  -0.0438   1.0000   0.0671
  -6.250  -0.5109   0.04037   0.03602  -0.0426   1.0000   0.0719
  -6.000  -0.5038   0.03747   0.03277  -0.0414   1.0000   0.0811
  -5.750  -0.4884   0.02830   0.02208  -0.0370   1.0000   0.0397
  -5.500  -0.4732   0.02629   0.01982  -0.0349   1.0000   0.0386
  -5.250  -0.4580   0.02383   0.01702  -0.0331   1.0000   0.0385
  -5.000  -0.4411   0.02121   0.01406  -0.0317   1.0000   0.0405
  -4.750  -0.4223   0.01960   0.01234  -0.0304   1.0000   0.0425
  -4.500  -0.4019   0.01834   0.01092  -0.0292   1.0000   0.0437
  -4.250  -0.3679   0.01712   0.00957  -0.0305   0.9970   0.0467
  -4.000  -0.3306   0.01642   0.00869  -0.0324   0.9932   0.0509
  -3.750  -0.2975   0.01483   0.00713  -0.0337   0.9891   0.0551
  -3.500  -0.2608   0.01406   0.00635  -0.0357   0.9849   0.0626
  -3.250  -0.2249   0.01323   0.00551  -0.0377   0.9807   0.0780
  -3.000  -0.1943   0.01142   0.00468  -0.0394   0.9755   0.2623
  -2.750  -0.1630   0.01021   0.00464  -0.0409   0.9712   0.5449
  -2.500  -0.1312   0.01002   0.00479  -0.0414   0.9660   0.6558
  -2.250  -0.0993   0.01005   0.00491  -0.0418   0.9601   0.7182
  -2.000  -0.0624   0.01010   0.00498  -0.0432   0.9562   0.7621
  -1.750  -0.0322   0.01010   0.00501  -0.0432   0.9501   0.7943
  -1.500   0.0007   0.01004   0.00496  -0.0438   0.9448   0.8207
  -1.250   0.0380   0.00995   0.00483  -0.0453   0.9410   0.8413
  -1.000   0.0658   0.00986   0.00475  -0.0449   0.9320   0.8586
  -0.750   0.1013   0.00972   0.00461  -0.0458   0.9266   0.8764
  -0.500   0.1261   0.00965   0.00455  -0.0446   0.9169   0.8981
  -0.250   0.1570   0.00955   0.00446  -0.0445   0.9095   0.9190
   0.000   0.1910   0.00942   0.00432  -0.0451   0.9004   0.9385
   0.250   0.2308   0.00930   0.00419  -0.0469   0.8901   0.9580
   0.500   0.2745   0.00915   0.00401  -0.0497   0.8796   0.9740
   0.750   0.3170   0.00901   0.00383  -0.0524   0.8693   0.9871
   1.000   0.3555   0.00894   0.00375  -0.0548   0.8585   1.0000
   1.250   0.3721   0.00895   0.00374  -0.0529   0.8461   1.0000
   1.500   0.3930   0.00899   0.00374  -0.0516   0.8344   1.0000
   1.750   0.4168   0.00904   0.00375  -0.0507   0.8228   1.0000
   2.000   0.4417   0.00909   0.00378  -0.0499   0.8107   1.0000
   2.250   0.4670   0.00914   0.00381  -0.0492   0.7981   1.0000
   2.500   0.4923   0.00919   0.00386  -0.0484   0.7832   1.0000
   2.750   0.5170   0.00919   0.00385  -0.0474   0.7634   1.0000
   3.000   0.5416   0.00914   0.00380  -0.0462   0.7391   1.0000
   3.250   0.5662   0.00911   0.00374  -0.0451   0.7095   1.0000
   3.500   0.5909   0.00913   0.00368  -0.0440   0.6738   1.0000
   3.750   0.6152   0.00925   0.00367  -0.0429   0.6298   1.0000
   4.000   0.6389   0.00951   0.00378  -0.0418   0.5751   1.0000
   4.250   0.6610   0.00997   0.00392  -0.0406   0.4971   1.0000
   4.500   0.6818   0.01066   0.00421  -0.0393   0.4176   1.0000
   4.750   0.7037   0.01133   0.00459  -0.0385   0.3515   1.0000
   5.000   0.7260   0.01202   0.00502  -0.0377   0.2945   1.0000
   5.250   0.7485   0.01270   0.00551  -0.0371   0.2369   1.0000
   5.500   0.7676   0.01392   0.00617  -0.0361   0.1323   1.0000
   5.750   0.7871   0.01524   0.00716  -0.0349   0.0884   1.0000
   6.000   0.8095   0.01613   0.00805  -0.0340   0.0784   1.0000
   6.250   0.8306   0.01717   0.00905  -0.0331   0.0712   1.0000
   6.500   0.8543   0.01781   0.00978  -0.0325   0.0634   1.0000
   6.750   0.8761   0.01877   0.01078  -0.0317   0.0557   1.0000
   7.000   0.8957   0.02019   0.01215  -0.0307   0.0475   1.0000
   7.250   0.9181   0.02125   0.01335  -0.0298   0.0400   1.0000
   7.500   0.9379   0.02363   0.01582  -0.0285   0.0339   1.0000
   7.750   0.9597   0.02472   0.01701  -0.0278   0.0291   1.0000
   8.000   0.9796   0.02795   0.02037  -0.0269   0.0264   1.0000
   8.250   0.9997   0.03039   0.02317  -0.0257   0.0251   1.0000
   8.500   1.0184   0.03287   0.02602  -0.0243   0.0244   1.0000
   8.750   1.0334   0.03598   0.02957  -0.0227   0.0240   1.0000
   9.000   1.0437   0.03965   0.03370  -0.0208   0.0239   1.0000
   9.250   1.0491   0.04366   0.03815  -0.0188   0.0240   1.0000
   9.500   1.0503   0.04757   0.04247  -0.0166   0.0240   1.0000
   9.750   1.0468   0.05142   0.04669  -0.0145   0.0238   1.0000
  10.000   1.0386   0.05502   0.05061  -0.0122   0.0234   1.0000
  10.250   1.0224   0.05867   0.05450  -0.0096   0.0235   1.0000
  10.500   1.0024   0.06257   0.05859  -0.0079   0.0239   1.0000
  10.750   0.9822   0.06691   0.06310  -0.0080   0.0241   1.0000
  11.000   0.9598   0.07228   0.06863  -0.0098   0.0244   1.0000
  11.250   0.9369   0.07879   0.07526  -0.0133   0.0249   1.0000
<< Back to HQ 1.5/8.5 AIRFOIL (hq1585-il)

Polar data table (+)

Polar graphs


<< Back to HQ 1.5/8.5 AIRFOIL (hq1585-il)