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

HQ 1.5/11 AIRFOIL (hq1511-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/11 AIRFOIL (hq1511-il)
Reynolds number: 100,000
Max Cl/Cd: 50.5 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1511-il-100000.txt
Download as CSV file: xf-hq1511-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/11 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4831   0.11132   0.10622  -0.0193   1.0000   0.1061
 -10.250  -0.5030   0.10729   0.10230  -0.0252   1.0000   0.1114
 -10.000  -0.5149   0.10199   0.09710  -0.0291   1.0000   0.1130
  -9.750  -0.4856   0.09924   0.09430  -0.0245   1.0000   0.1169
  -9.500  -0.4854   0.09578   0.09088  -0.0260   1.0000   0.1226
  -9.250  -0.5289   0.08977   0.08505  -0.0365   1.0000   0.1261
  -9.000  -0.4873   0.08808   0.08329  -0.0284   1.0000   0.1337
  -8.750  -0.5043   0.08343   0.07876  -0.0329   1.0000   0.1385
  -8.500  -0.4996   0.08006   0.07543  -0.0321   1.0000   0.1459
  -8.250  -0.5984   0.05845   0.05333  -0.0483   1.0000   0.0677
  -8.000  -0.6221   0.05126   0.04541  -0.0466   1.0000   0.0567
  -7.750  -0.6194   0.04722   0.04125  -0.0450   1.0000   0.0553
  -7.500  -0.6178   0.04360   0.03737  -0.0428   1.0000   0.0543
  -7.250  -0.6153   0.04041   0.03383  -0.0403   1.0000   0.0544
  -7.000  -0.6106   0.03762   0.03063  -0.0376   1.0000   0.0551
  -6.750  -0.6035   0.03533   0.02781  -0.0348   1.0000   0.0567
  -6.500  -0.5943   0.03254   0.02457  -0.0325   1.0000   0.0582
  -6.250  -0.5813   0.02994   0.02185  -0.0308   1.0000   0.0608
  -6.000  -0.5658   0.02808   0.01979  -0.0289   1.0000   0.0654
  -5.750  -0.5488   0.02588   0.01718  -0.0271   1.0000   0.0723
  -5.500  -0.5329   0.02527   0.01637  -0.0255   1.0000   0.0842
  -5.250  -0.5162   0.02424   0.01532  -0.0244   1.0000   0.0973
  -5.000  -0.4969   0.02272   0.01384  -0.0234   1.0000   0.1062
  -4.750  -0.4760   0.02173   0.01264  -0.0221   1.0000   0.1126
  -4.500  -0.4565   0.02063   0.01168  -0.0210   1.0000   0.1222
  -4.250  -0.4371   0.01973   0.01085  -0.0199   1.0000   0.1329
  -4.000  -0.4181   0.01897   0.01018  -0.0187   1.0000   0.1463
  -3.750  -0.3989   0.01833   0.00965  -0.0176   1.0000   0.1699
  -3.500  -0.3801   0.01732   0.00905  -0.0167   1.0000   0.2245
  -3.250  -0.3690   0.01539   0.00897  -0.0143   1.0000   0.5313
  -3.000  -0.3449   0.01574   0.00957  -0.0130   0.9959   0.6495
  -2.750  -0.3146   0.01626   0.01005  -0.0129   0.9890   0.7072
  -2.500  -0.2823   0.01678   0.01051  -0.0131   0.9824   0.7489
  -2.250  -0.2539   0.01714   0.01086  -0.0124   0.9737   0.7834
  -2.000  -0.2266   0.01747   0.01116  -0.0112   0.9649   0.8188
  -1.750  -0.1979   0.01777   0.01143  -0.0101   0.9569   0.8543
  -1.500  -0.1754   0.01785   0.01147  -0.0084   0.9481   0.8815
  -1.250  -0.1348   0.01804   0.01155  -0.0106   0.9425   0.9011
  -1.000  -0.1016   0.01815   0.01159  -0.0115   0.9351   0.9231
  -0.750  -0.0444   0.01844   0.01178  -0.0169   0.9306   0.9443
  -0.500   0.0262   0.01872   0.01193  -0.0252   0.9278   0.9584
  -0.250   0.0875   0.01887   0.01199  -0.0322   0.9226   0.9702
   0.000   0.1550   0.01886   0.01192  -0.0404   0.9171   0.9776
   0.250   0.2268   0.01869   0.01170  -0.0489   0.9126   0.9826
   0.500   0.2867   0.01841   0.01140  -0.0554   0.9027   0.9893
   0.750   0.3504   0.01788   0.01088  -0.0620   0.8929   0.9943
   1.000   0.4134   0.01721   0.01021  -0.0682   0.8836   0.9995
   1.250   0.4378   0.01700   0.01002  -0.0678   0.8699   1.0000
   1.500   0.4571   0.01687   0.00992  -0.0664   0.8561   1.0000
   1.750   0.4734   0.01682   0.00988  -0.0645   0.8427   1.0000
   2.000   0.4877   0.01682   0.00990  -0.0623   0.8295   1.0000
   2.250   0.5017   0.01680   0.00991  -0.0599   0.8164   1.0000
   2.500   0.5172   0.01675   0.00987  -0.0575   0.8033   1.0000
   2.750   0.5357   0.01666   0.00980  -0.0555   0.7899   1.0000
   3.000   0.5571   0.01654   0.00969  -0.0538   0.7762   1.0000
   3.250   0.5802   0.01642   0.00958  -0.0523   0.7616   1.0000
   3.500   0.6039   0.01632   0.00953  -0.0509   0.7458   1.0000
   3.750   0.6280   0.01621   0.00945  -0.0496   0.7285   1.0000
   4.000   0.6533   0.01603   0.00929  -0.0482   0.7096   1.0000
   4.250   0.6799   0.01575   0.00901  -0.0469   0.6892   1.0000
   4.500   0.7036   0.01548   0.00875  -0.0451   0.6605   1.0000
   4.750   0.7257   0.01528   0.00853  -0.0431   0.6245   1.0000
   5.000   0.7485   0.01518   0.00834  -0.0413   0.5841   1.0000
   5.250   0.7702   0.01530   0.00834  -0.0395   0.5367   1.0000
   5.500   0.7913   0.01567   0.00846  -0.0379   0.4887   1.0000
   5.750   0.8121   0.01627   0.00882  -0.0364   0.4464   1.0000
   6.000   0.8330   0.01695   0.00929  -0.0351   0.4101   1.0000
   6.250   0.8535   0.01757   0.00985  -0.0339   0.3753   1.0000
   6.500   0.8732   0.01820   0.01038  -0.0326   0.3415   1.0000
   6.750   0.8914   0.01902   0.01098  -0.0311   0.3069   1.0000
   7.000   0.9083   0.01995   0.01179  -0.0296   0.2676   1.0000
   7.250   0.9234   0.02098   0.01266  -0.0278   0.2267   1.0000
   7.500   0.9377   0.02207   0.01360  -0.0260   0.1868   1.0000
   7.750   0.9522   0.02307   0.01445  -0.0242   0.1511   1.0000
   8.000   0.9660   0.02460   0.01574  -0.0224   0.1246   1.0000
   8.250   0.9820   0.02654   0.01747  -0.0209   0.1062   1.0000
   8.500   1.0015   0.02843   0.01928  -0.0199   0.0929   1.0000
   8.750   1.0221   0.03009   0.02096  -0.0189   0.0833   1.0000
   9.000   1.0464   0.03239   0.02345  -0.0184   0.0769   1.0000
   9.250   1.0674   0.03420   0.02532  -0.0176   0.0710   1.0000
   9.500   1.0894   0.03698   0.02826  -0.0171   0.0669   1.0000
   9.750   1.1066   0.03955   0.03125  -0.0157   0.0643   1.0000
  10.000   1.1213   0.04220   0.03423  -0.0142   0.0620   1.0000
  10.250   1.1357   0.04454   0.03671  -0.0131   0.0592   1.0000
  10.500   1.1476   0.04912   0.04142  -0.0125   0.0568   1.0000
  10.750   1.1474   0.05212   0.04490  -0.0099   0.0563   1.0000
  11.000   1.1431   0.05559   0.04879  -0.0072   0.0560   1.0000
  11.250   1.1348   0.05932   0.05287  -0.0047   0.0560   1.0000
  11.500   1.1215   0.06299   0.05683  -0.0019   0.0561   1.0000
  11.750   1.1057   0.06688   0.06095   0.0004   0.0563   1.0000
  12.000   1.0878   0.07110   0.06538   0.0018   0.0565   1.0000
  12.250   1.0714   0.07614   0.07059   0.0022   0.0568   1.0000
  12.500   1.0480   0.07964   0.07434   0.0021   0.0572   1.0000
  12.750   0.9370   0.09476   0.09002  -0.0097   0.0618   1.0000
<< Back to HQ 1.5/11 AIRFOIL (hq1511-il)

Polar data table (+)

Polar graphs


<< Back to HQ 1.5/11 AIRFOIL (hq1511-il)