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

HQ 3.0/12 AIRFOIL (hq3012-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 3.0/12 AIRFOIL (hq3012-il)
Reynolds number: 100,000
Max Cl/Cd: 57.42 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq3012-il-100000.txt
Download as CSV file: xf-hq3012-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.0/12 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3535   0.10183   0.09702  -0.0424   1.0000   0.1058
  -9.250  -0.3831   0.09965   0.09502  -0.0463   1.0000   0.1084
  -9.000  -0.4149   0.09758   0.09313  -0.0480   1.0000   0.1088
  -8.750  -0.3702   0.09287   0.08834  -0.0422   1.0000   0.1128
  -8.500  -0.3732   0.09070   0.08622  -0.0403   1.0000   0.1163
  -8.250  -0.3895   0.08884   0.08448  -0.0386   1.0000   0.1190
  -8.000  -0.4188   0.08753   0.08333  -0.0363   1.0000   0.1209
  -7.750  -0.4499   0.08643   0.08237  -0.0333   1.0000   0.1214
  -7.500  -0.4869   0.08408   0.08013  -0.0340   1.0000   0.1222
  -7.250  -0.5303   0.08116   0.07712  -0.0375   1.0000   0.1232
  -7.000  -0.5198   0.07738   0.07349  -0.0334   1.0000   0.1255
  -6.750  -0.5143   0.07563   0.07181  -0.0297   1.0000   0.1285
  -6.500  -0.5224   0.07270   0.06887  -0.0297   1.0000   0.1332
  -6.250  -0.5384   0.06794   0.06395  -0.0333   1.0000   0.1402
  -6.000  -0.5258   0.06550   0.06156  -0.0317   0.9988   0.1452
  -5.750  -0.4958   0.04747   0.04192  -0.0474   0.9928   0.0826
  -5.500  -0.4658   0.04236   0.03639  -0.0499   0.9882   0.0745
  -5.250  -0.4274   0.03690   0.02952  -0.0526   0.9841   0.0671
  -5.000  -0.3954   0.03309   0.02520  -0.0544   0.9802   0.0677
  -4.750  -0.3648   0.03094   0.02298  -0.0559   0.9753   0.0729
  -4.500  -0.3259   0.02947   0.02084  -0.0578   0.9708   0.0806
  -4.250  -0.2953   0.02765   0.01901  -0.0590   0.9661   0.0911
  -4.000  -0.2636   0.02640   0.01765  -0.0601   0.9607   0.1033
  -3.750  -0.2259   0.02536   0.01652  -0.0622   0.9565   0.1173
  -3.500  -0.1976   0.02447   0.01562  -0.0624   0.9505   0.1280
  -3.250  -0.1655   0.02375   0.01500  -0.0635   0.9450   0.1444
  -3.000  -0.1265   0.02306   0.01444  -0.0657   0.9409   0.1693
  -2.750  -0.1051   0.02225   0.01398  -0.0652   0.9335   0.2183
  -2.500  -0.0820   0.02088   0.01453  -0.0639   0.9285   0.6263
  -2.250  -0.0625   0.02139   0.01506  -0.0614   0.9210   0.6921
  -2.000  -0.0366   0.02181   0.01543  -0.0600   0.9144   0.7358
  -1.750  -0.0158   0.02212   0.01571  -0.0580   0.9072   0.7671
  -1.500   0.0082   0.02231   0.01585  -0.0564   0.9001   0.7970
  -1.250   0.0263   0.02244   0.01598  -0.0537   0.8929   0.8277
  -1.000   0.0426   0.02243   0.01598  -0.0503   0.8853   0.8660
  -0.750   0.0588   0.02239   0.01597  -0.0470   0.8779   0.9031
  -0.500   0.0968   0.02241   0.01595  -0.0483   0.8714   0.9367
  -0.250   0.1634   0.02249   0.01589  -0.0555   0.8681   0.9573
   0.000   0.2122   0.02270   0.01600  -0.0606   0.8598   0.9740
   0.250   0.2769   0.02267   0.01585  -0.0680   0.8554   0.9866
   0.500   0.3421   0.02258   0.01567  -0.0754   0.8518   1.0000
   0.750   0.3454   0.02265   0.01568  -0.0725   0.8399   1.0000
   1.000   0.3634   0.02285   0.01581  -0.0719   0.8288   1.0000
   1.250   0.4208   0.02237   0.01526  -0.0768   0.8242   1.0000
   1.500   0.4474   0.02240   0.01524  -0.0770   0.8124   1.0000
   1.750   0.4845   0.02221   0.01500  -0.0786   0.8034   1.0000
   2.000   0.5267   0.02182   0.01458  -0.0807   0.7962   1.0000
   2.250   0.5551   0.02187   0.01460  -0.0809   0.7855   1.0000
   2.500   0.6014   0.02124   0.01396  -0.0832   0.7801   1.0000
   2.750   0.6282   0.02125   0.01396  -0.0829   0.7684   1.0000
   3.000   0.6598   0.02104   0.01377  -0.0830   0.7579   1.0000
   3.250   0.7027   0.02029   0.01300  -0.0844   0.7505   1.0000
   3.500   0.7315   0.02004   0.01276  -0.0839   0.7376   1.0000
   3.750   0.7616   0.01971   0.01247  -0.0836   0.7246   1.0000
   4.000   0.7922   0.01936   0.01213  -0.0832   0.7113   1.0000
   4.250   0.8228   0.01902   0.01180  -0.0829   0.6975   1.0000
   4.500   0.8528   0.01869   0.01150  -0.0824   0.6826   1.0000
   4.750   0.8825   0.01837   0.01119  -0.0819   0.6662   1.0000
   5.000   0.9123   0.01805   0.01086  -0.0814   0.6486   1.0000
   5.250   0.9373   0.01794   0.01078  -0.0802   0.6273   1.0000
   5.500   0.9642   0.01775   0.01056  -0.0793   0.6044   1.0000
   5.750   0.9873   0.01774   0.01054  -0.0779   0.5776   1.0000
   6.000   1.0105   0.01780   0.01053  -0.0765   0.5483   1.0000
   6.250   1.0329   0.01799   0.01062  -0.0751   0.5166   1.0000
   6.500   1.0533   0.01838   0.01087  -0.0735   0.4832   1.0000
   6.750   1.0726   0.01893   0.01126  -0.0719   0.4498   1.0000
   7.000   1.0917   0.01961   0.01177  -0.0703   0.4187   1.0000
   7.250   1.1102   0.02037   0.01240  -0.0687   0.3898   1.0000
   7.500   1.1287   0.02119   0.01308  -0.0673   0.3636   1.0000
   7.750   1.1473   0.02203   0.01377  -0.0658   0.3398   1.0000
   8.000   1.1643   0.02288   0.01459  -0.0643   0.3171   1.0000
   8.250   1.1825   0.02379   0.01538  -0.0629   0.2968   1.0000
   8.500   1.1994   0.02473   0.01631  -0.0614   0.2775   1.0000
   8.750   1.2153   0.02565   0.01724  -0.0598   0.2591   1.0000
   9.000   1.2309   0.02659   0.01816  -0.0581   0.2420   1.0000
   9.250   1.2455   0.02754   0.01910  -0.0564   0.2258   1.0000
   9.500   1.2593   0.02854   0.02008  -0.0545   0.2103   1.0000
   9.750   1.2719   0.02960   0.02114  -0.0526   0.1951   1.0000
  10.000   1.2843   0.03079   0.02229  -0.0507   0.1803   1.0000
  10.250   1.2936   0.03198   0.02349  -0.0483   0.1664   1.0000
  10.500   1.3006   0.03320   0.02478  -0.0457   0.1536   1.0000
  10.750   1.3085   0.03460   0.02627  -0.0433   0.1416   1.0000
  11.000   1.3182   0.03618   0.02795  -0.0413   0.1309   1.0000
  11.250   1.3295   0.03780   0.02958  -0.0396   0.1214   1.0000
  11.500   1.3415   0.03929   0.03105  -0.0381   0.1129   1.0000
  11.750   1.3493   0.04125   0.03330  -0.0361   0.1059   1.0000
  12.000   1.3637   0.04293   0.03482  -0.0352   0.0981   1.0000
  12.250   1.3583   0.04469   0.03696  -0.0321   0.0931   1.0000
  12.500   1.3662   0.04622   0.03828  -0.0309   0.0858   1.0000
  12.750   1.3576   0.04833   0.04081  -0.0282   0.0819   1.0000
  13.000   1.3547   0.05012   0.04268  -0.0265   0.0766   1.0000
  13.250   1.3519   0.05256   0.04521  -0.0251   0.0714   1.0000
  13.500   1.3433   0.05514   0.04805  -0.0238   0.0669   1.0000
  13.750   1.3444   0.05770   0.05049  -0.0230   0.0614   1.0000
  14.000   1.3302   0.06134   0.05453  -0.0222   0.0579   1.0000
  14.250   1.3233   0.06451   0.05780  -0.0219   0.0537   1.0000
  14.500   1.3160   0.06858   0.06193  -0.0216   0.0495   1.0000
  14.750   1.3005   0.07345   0.06715  -0.0220   0.0470   1.0000
  15.000   1.2881   0.07816   0.07208  -0.0228   0.0446   1.0000
  15.250   1.2850   0.08194   0.07587  -0.0233   0.0420   1.0000
  15.500   1.2737   0.08767   0.08175  -0.0243   0.0404   1.0000
  15.750   1.2540   0.09403   0.08841  -0.0267   0.0401   1.0000
  16.000   1.2333   0.10102   0.09564  -0.0299   0.0400   1.0000
  16.250   1.2112   0.10872   0.10360  -0.0339   0.0399   1.0000
  16.500   1.1894   0.11690   0.11199  -0.0386   0.0401   1.0000
  16.750   1.1659   0.12597   0.12126  -0.0442   0.0404   1.0000
<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)

Polar data table (+)

Polar graphs


<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)