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HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il)
Reynolds number: 1,000,000
Max Cl/Cd: 122 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq259-il-1000000.txt
Download as CSV file: xf-hq259-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.5/9 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4089   0.08265   0.08110  -0.0336   1.0000   0.0073
  -8.250  -0.4106   0.07901   0.07749  -0.0349   1.0000   0.0072
  -8.000  -0.4161   0.07542   0.07394  -0.0359   1.0000   0.0073
  -7.750  -0.4306   0.07272   0.07130  -0.0348   1.0000   0.0073
  -7.500  -0.4191   0.06578   0.06436  -0.0457   0.9959   0.0073
  -7.250  -0.3968   0.05618   0.05465  -0.0610   0.9884   0.0075
  -7.000  -0.3719   0.04828   0.04657  -0.0706   0.9837   0.0080
  -6.750  -0.3442   0.04212   0.04021  -0.0764   0.9788   0.0087
  -6.500  -0.3141   0.03829   0.03620  -0.0791   0.9729   0.0092
  -6.250  -0.2902   0.03309   0.03070  -0.0819   0.9646   0.0093
  -6.000  -0.2922   0.01648   0.01253  -0.0812   0.9451   0.0055
  -5.750  -0.2711   0.01364   0.00930  -0.0804   0.9360   0.0062
  -5.500  -0.2455   0.01303   0.00858  -0.0801   0.9284   0.0068
  -5.250  -0.2201   0.01220   0.00761  -0.0797   0.9202   0.0073
  -5.000  -0.1945   0.01132   0.00657  -0.0792   0.9128   0.0076
  -4.750  -0.1689   0.01052   0.00564  -0.0787   0.9050   0.0079
  -4.500  -0.1427   0.00996   0.00497  -0.0784   0.8979   0.0084
  -4.250  -0.1162   0.00950   0.00441  -0.0780   0.8905   0.0089
  -4.000  -0.0900   0.00886   0.00364  -0.0777   0.8834   0.0091
  -3.750  -0.0632   0.00841   0.00309  -0.0774   0.8761   0.0093
  -3.500  -0.0366   0.00770   0.00220  -0.0770   0.8689   0.0128
  -3.250  -0.0094   0.00741   0.00189  -0.0767   0.8614   0.0226
  -3.000   0.0180   0.00720   0.00172  -0.0766   0.8532   0.0414
  -2.750   0.0452   0.00706   0.00157  -0.0765   0.8454   0.0564
  -2.500   0.0728   0.00688   0.00143  -0.0765   0.8374   0.0770
  -2.250   0.0998   0.00653   0.00129  -0.0765   0.8303   0.1476
  -2.000   0.1257   0.00582   0.00112  -0.0766   0.8227   0.3232
  -1.750   0.1519   0.00531   0.00102  -0.0766   0.8152   0.4756
  -1.500   0.1789   0.00512   0.00098  -0.0764   0.8065   0.5458
  -1.250   0.2064   0.00503   0.00095  -0.0763   0.7966   0.5878
  -1.000   0.2339   0.00498   0.00093  -0.0762   0.7864   0.6168
  -0.750   0.2612   0.00495   0.00091  -0.0760   0.7758   0.6487
  -0.500   0.2886   0.00494   0.00091  -0.0758   0.7654   0.6724
  -0.250   0.3161   0.00491   0.00092  -0.0757   0.7558   0.6996
   0.000   0.3433   0.00490   0.00095  -0.0755   0.7462   0.7265
   0.250   0.3707   0.00492   0.00096  -0.0753   0.7353   0.7422
   0.500   0.3981   0.00495   0.00097  -0.0751   0.7231   0.7555
   0.750   0.4256   0.00498   0.00098  -0.0750   0.7109   0.7661
   1.000   0.4532   0.00501   0.00099  -0.0749   0.6984   0.7767
   1.250   0.4805   0.00504   0.00103  -0.0748   0.6854   0.7876
   1.500   0.5075   0.00509   0.00106  -0.0745   0.6701   0.7992
   1.750   0.5343   0.00515   0.00110  -0.0743   0.6522   0.8118
   2.000   0.5610   0.00521   0.00114  -0.0740   0.6329   0.8259
   2.250   0.5873   0.00528   0.00120  -0.0737   0.6132   0.8425
   2.500   0.6127   0.00535   0.00128  -0.0732   0.5880   0.8637
   2.750   0.6363   0.00539   0.00135  -0.0722   0.5602   0.9025
   3.000   0.6698   0.00549   0.00143  -0.0735   0.5224   1.0000
   3.250   0.6951   0.00581   0.00156  -0.0731   0.4787   1.0000
   3.500   0.7200   0.00617   0.00174  -0.0727   0.4312   1.0000
   3.750   0.7449   0.00654   0.00192  -0.0723   0.3864   1.0000
   4.000   0.7702   0.00688   0.00211  -0.0720   0.3479   1.0000
   4.250   0.7941   0.00736   0.00234  -0.0715   0.2946   1.0000
   4.500   0.8182   0.00781   0.00259  -0.0710   0.2508   1.0000
   4.750   0.8430   0.00820   0.00284  -0.0706   0.2166   1.0000
   5.000   0.8672   0.00865   0.00310  -0.0702   0.1786   1.0000
   5.250   0.8915   0.00907   0.00337  -0.0697   0.1456   1.0000
   5.500   0.9151   0.00958   0.00369  -0.0692   0.1106   1.0000
   5.750   0.9346   0.01057   0.00427  -0.0680   0.0434   1.0000
   6.000   0.9549   0.01150   0.00497  -0.0669   0.0063   1.0000
   6.250   0.9799   0.01185   0.00539  -0.0664   0.0056   1.0000
   6.500   1.0044   0.01226   0.00585  -0.0659   0.0050   1.0000
   6.750   1.0282   0.01274   0.00639  -0.0653   0.0044   1.0000
   7.000   1.0494   0.01359   0.00737  -0.0641   0.0036   1.0000
   7.500   1.0883   0.01560   0.00965  -0.0613   0.0033   1.0000
   7.750   1.1069   0.01664   0.01080  -0.0598   0.0033   1.0000
   8.000   1.1225   0.01802   0.01232  -0.0579   0.0033   1.0000
   8.250   1.1375   0.01955   0.01396  -0.0559   0.0034   1.0000
   8.500   1.1570   0.02044   0.01494  -0.0547   0.0035   1.0000
   8.750   1.1763   0.02133   0.01591  -0.0534   0.0036   1.0000
   9.000   1.1954   0.02221   0.01690  -0.0521   0.0038   1.0000
  17.250   0.9703   0.22187   0.22037  -0.1097   0.0061   1.0000
  17.500   0.9783   0.22607   0.22457  -0.1120   0.0059   1.0000
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