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

HQ 3.0/9 AIRFOIL (hq309-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HQ 3.0/9 AIRFOIL (hq309-il)
Reynolds number: 200,000
Max Cl/Cd: 84.76 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq309-il-200000.txt
Download as CSV file: xf-hq309-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.0/9 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3649   0.08957   0.08619  -0.0359   1.0000   0.0365
  -8.000  -0.3691   0.08679   0.08349  -0.0362   1.0000   0.0375
  -7.750  -0.3775   0.08445   0.08123  -0.0358   1.0000   0.0381
  -7.500  -0.3928   0.08275   0.07962  -0.0337   1.0000   0.0384
  -7.250  -0.4117   0.08128   0.07825  -0.0312   1.0000   0.0386
  -7.000  -0.4256   0.07904   0.07608  -0.0309   1.0000   0.0389
  -6.750  -0.4360   0.07587   0.07295  -0.0334   1.0000   0.0393
  -6.500  -0.4408   0.07248   0.06948  -0.0367   1.0000   0.0399
  -6.250  -0.4096   0.06682   0.06348  -0.0476   0.9949   0.0405
  -6.000  -0.3934   0.05829   0.05487  -0.0534   0.9904   0.0415
  -5.750  -0.3710   0.05414   0.05076  -0.0560   0.9864   0.0430
  -5.500  -0.3408   0.05011   0.04662  -0.0607   0.9830   0.0453
  -5.250  -0.3119   0.04598   0.04224  -0.0648   0.9773   0.0486
  -5.000  -0.2724   0.04090   0.03641  -0.0709   0.9721   0.0547
  -4.750  -0.2410   0.03704   0.03257  -0.0741   0.9697   0.0571
  -4.500  -0.2126   0.03431   0.02962  -0.0757   0.9639   0.0614
  -4.250  -0.1776   0.03106   0.02589  -0.0785   0.9595   0.0703
  -4.000  -0.1309   0.02415   0.01768  -0.0784   0.9577   0.0288
  -3.750  -0.0941   0.02087   0.01405  -0.0808   0.9558   0.0317
  -3.500  -0.0654   0.01946   0.01243  -0.0809   0.9498   0.0336
  -3.250  -0.0296   0.01779   0.01053  -0.0820   0.9462   0.0344
  -3.000   0.0088   0.01649   0.00911  -0.0837   0.9436   0.0381
  -2.750   0.0490   0.01529   0.00790  -0.0861   0.9416   0.0519
  -2.500   0.0751   0.01439   0.00704  -0.0859   0.9342   0.0829
  -2.250   0.1138   0.01357   0.00645  -0.0885   0.9309   0.1396
  -2.000   0.1453   0.01170   0.00642  -0.0898   0.9284   0.6425
  -1.750   0.1699   0.01167   0.00645  -0.0887   0.9207   0.7014
  -1.500   0.2040   0.01154   0.00632  -0.0893   0.9166   0.7497
  -1.250   0.2317   0.01146   0.00626  -0.0886   0.9102   0.7875
  -1.000   0.2579   0.01127   0.00614  -0.0873   0.9039   0.8307
  -0.750   0.2794   0.01100   0.00597  -0.0848   0.8977   0.8750
  -0.500   0.3045   0.01077   0.00575  -0.0835   0.8906   0.9065
  -0.250   0.3432   0.01051   0.00544  -0.0852   0.8869   0.9297
   0.000   0.3848   0.01040   0.00531  -0.0881   0.8794   0.9641
   0.250   0.4293   0.01019   0.00502  -0.0916   0.8743   1.0000
   0.500   0.4568   0.01014   0.00488  -0.0917   0.8638   1.0000
   0.750   0.4858   0.01004   0.00470  -0.0918   0.8537   1.0000
   1.000   0.5165   0.00992   0.00449  -0.0920   0.8456   1.0000
   1.250   0.5431   0.00995   0.00448  -0.0918   0.8348   1.0000
   1.500   0.5704   0.00995   0.00444  -0.0915   0.8245   1.0000
   1.750   0.5981   0.00991   0.00435  -0.0912   0.8138   1.0000
   2.000   0.6257   0.00987   0.00425  -0.0908   0.8024   1.0000
   2.250   0.6531   0.00983   0.00420  -0.0904   0.7904   1.0000
   2.500   0.6798   0.00983   0.00418  -0.0899   0.7774   1.0000
   2.750   0.7065   0.00985   0.00417  -0.0894   0.7637   1.0000
   3.000   0.7330   0.00987   0.00418  -0.0889   0.7490   1.0000
   3.250   0.7588   0.00992   0.00423  -0.0883   0.7322   1.0000
   3.500   0.7848   0.00997   0.00431  -0.0876   0.7137   1.0000
   3.750   0.8109   0.01004   0.00434  -0.0870   0.6943   1.0000
   4.000   0.8361   0.01014   0.00443  -0.0862   0.6707   1.0000
   4.250   0.8610   0.01028   0.00453  -0.0854   0.6445   1.0000
   4.500   0.8853   0.01047   0.00469  -0.0845   0.6138   1.0000
   4.750   0.9086   0.01072   0.00486  -0.0834   0.5765   1.0000
   5.000   0.9308   0.01108   0.00507  -0.0822   0.5324   1.0000
   5.250   0.9515   0.01157   0.00536  -0.0808   0.4813   1.0000
   5.500   0.9709   0.01220   0.00575  -0.0793   0.4282   1.0000
   5.750   0.9898   0.01292   0.00624  -0.0778   0.3769   1.0000
   6.000   1.0091   0.01362   0.00675  -0.0765   0.3300   1.0000
   6.250   1.0286   0.01433   0.00729  -0.0753   0.2910   1.0000
   6.500   1.0488   0.01501   0.00786  -0.0742   0.2588   1.0000
   6.750   1.0684   0.01571   0.00843  -0.0731   0.2264   1.0000
   7.000   1.0885   0.01638   0.00899  -0.0720   0.1912   1.0000
   7.250   1.1074   0.01715   0.00960  -0.0709   0.1546   1.0000
   7.500   1.1266   0.01794   0.01034  -0.0698   0.1238   1.0000
   7.750   1.1440   0.01891   0.01117  -0.0684   0.0941   1.0000
   8.000   1.1603   0.02002   0.01214  -0.0669   0.0666   1.0000
   8.250   1.1740   0.02139   0.01338  -0.0650   0.0439   1.0000
   8.500   1.1796   0.02361   0.01552  -0.0618   0.0237   1.0000
   8.750   1.1939   0.02479   0.01685  -0.0598   0.0194   1.0000
   9.000   1.2027   0.02636   0.01851  -0.0571   0.0169   1.0000
   9.250   1.2023   0.02887   0.02115  -0.0533   0.0155   1.0000
   9.500   1.2130   0.03055   0.02297  -0.0510   0.0149   1.0000
   9.750   1.2244   0.03247   0.02506  -0.0489   0.0143   1.0000
  10.000   1.2367   0.03470   0.02754  -0.0470   0.0140   1.0000
  10.250   1.2486   0.03723   0.03032  -0.0452   0.0137   1.0000
  10.500   1.2579   0.04005   0.03342  -0.0432   0.0135   1.0000
  10.750   1.2629   0.04308   0.03677  -0.0409   0.0136   1.0000
  11.000   1.2625   0.04630   0.04031  -0.0384   0.0136   1.0000
  11.250   1.2575   0.04972   0.04404  -0.0360   0.0137   1.0000
  11.500   1.2482   0.05339   0.04802  -0.0337   0.0138   1.0000
  11.750   1.2361   0.05729   0.05221  -0.0320   0.0139   1.0000
  12.000   1.2213   0.06160   0.05679  -0.0309   0.0141   1.0000
  12.250   1.2046   0.06633   0.06177  -0.0306   0.0142   1.0000
  12.500   1.1871   0.07143   0.06711  -0.0313   0.0143   1.0000
  12.750   1.1680   0.07718   0.07307  -0.0329   0.0144   1.0000
  13.000   1.1484   0.08351   0.07961  -0.0356   0.0145   1.0000
  13.250   1.1281   0.09060   0.08689  -0.0393   0.0146   1.0000
  13.500   1.1077   0.09843   0.09489  -0.0441   0.0147   1.0000
  13.750   1.0876   0.10713   0.10375  -0.0498   0.0148   1.0000
  14.000   1.0673   0.11685   0.11360  -0.0564   0.0149   1.0000
<< Back to HQ 3.0/9 AIRFOIL (hq309-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.0/9 AIRFOIL (hq309-il)