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

HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il)
Reynolds number: 200,000
Max Cl/Cd: 67.53 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq2195-il-200000-n5.txt
Download as CSV file: xf-hq2195-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 2.1/9.5 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4884   0.08476   0.08108  -0.0356   1.0000   0.0189
  -9.500  -0.4941   0.07942   0.07578  -0.0385   1.0000   0.0188
  -9.250  -0.5027   0.07340   0.06982  -0.0422   1.0000   0.0187
  -9.000  -0.5149   0.06666   0.06313  -0.0475   1.0000   0.0185
  -8.750  -0.5375   0.06043   0.05688  -0.0518   1.0000   0.0184
  -8.500  -0.5523   0.05522   0.05156  -0.0534   1.0000   0.0183
  -8.250  -0.5657   0.04987   0.04603  -0.0536   1.0000   0.0183
  -8.000  -0.5767   0.04474   0.04062  -0.0525   1.0000   0.0184
  -7.750  -0.5751   0.03686   0.03193  -0.0543   0.9955   0.0192
  -7.500  -0.5505   0.03382   0.02867  -0.0569   0.9894   0.0200
  -7.250  -0.5220   0.03128   0.02583  -0.0592   0.9843   0.0207
  -7.000  -0.4948   0.02836   0.02251  -0.0609   0.9785   0.0213
  -6.750  -0.4655   0.02577   0.01949  -0.0625   0.9735   0.0220
  -6.500  -0.4334   0.02368   0.01701  -0.0642   0.9700   0.0231
  -6.250  -0.4042   0.02214   0.01511  -0.0649   0.9639   0.0245
  -6.000  -0.3727   0.02067   0.01333  -0.0660   0.9591   0.0251
  -5.750  -0.3406   0.01892   0.01138  -0.0673   0.9555   0.0259
  -5.500  -0.3126   0.01784   0.01022  -0.0677   0.9487   0.0268
  -5.250  -0.2815   0.01696   0.00928  -0.0686   0.9439   0.0279
  -5.000  -0.2495   0.01629   0.00854  -0.0696   0.9397   0.0298
  -4.750  -0.2213   0.01563   0.00779  -0.0698   0.9327   0.0315
  -4.500  -0.1905   0.01492   0.00699  -0.0705   0.9277   0.0328
  -4.250  -0.1625   0.01415   0.00619  -0.0708   0.9214   0.0346
  -4.000  -0.1338   0.01361   0.00562  -0.0711   0.9154   0.0372
  -3.750  -0.1031   0.01318   0.00512  -0.0718   0.9111   0.0413
  -3.500  -0.0764   0.01277   0.00471  -0.0717   0.9041   0.0478
  -3.250  -0.0475   0.01238   0.00430  -0.0719   0.8987   0.0581
  -3.000  -0.0194   0.01200   0.00397  -0.0721   0.8931   0.0786
  -2.750   0.0075   0.01155   0.00371  -0.0722   0.8871   0.1277
  -2.500   0.0335   0.01057   0.00341  -0.0726   0.8823   0.3032
  -2.250   0.0570   0.00983   0.00341  -0.0721   0.8761   0.5054
  -2.000   0.0830   0.00964   0.00344  -0.0715   0.8707   0.5814
  -1.750   0.1101   0.00956   0.00344  -0.0711   0.8651   0.6314
  -1.500   0.1353   0.00952   0.00345  -0.0701   0.8553   0.6744
  -1.250   0.1600   0.00946   0.00345  -0.0689   0.8440   0.7120
  -1.000   0.1855   0.00941   0.00339  -0.0679   0.8328   0.7372
  -0.750   0.2122   0.00937   0.00330  -0.0673   0.8228   0.7505
  -0.500   0.2386   0.00934   0.00326  -0.0668   0.8142   0.7610
  -0.250   0.2659   0.00932   0.00321  -0.0664   0.8075   0.7716
   0.000   0.2926   0.00931   0.00320  -0.0660   0.7996   0.7825
   0.250   0.3196   0.00930   0.00317  -0.0656   0.7925   0.7937
   0.500   0.3459   0.00929   0.00319  -0.0651   0.7840   0.8047
   0.750   0.3727   0.00928   0.00317  -0.0646   0.7762   0.8163
   1.000   0.3989   0.00928   0.00319  -0.0640   0.7668   0.8289
   1.250   0.4252   0.00927   0.00320  -0.0635   0.7573   0.8420
   1.500   0.4513   0.00926   0.00320  -0.0628   0.7476   0.8562
   1.750   0.4770   0.00925   0.00323  -0.0621   0.7358   0.8717
   2.000   0.5030   0.00923   0.00325  -0.0615   0.7234   0.8889
   2.250   0.5298   0.00921   0.00326  -0.0610   0.7090   0.9094
   2.500   0.5600   0.00918   0.00327  -0.0612   0.6915   0.9348
   2.750   0.5965   0.00917   0.00327  -0.0630   0.6672   0.9914
   3.000   0.6225   0.00929   0.00329  -0.0625   0.6320   1.0000
   3.250   0.6456   0.00956   0.00327  -0.0614   0.5621   1.0000
   3.500   0.6654   0.01018   0.00339  -0.0598   0.4822   1.0000
   3.750   0.6876   0.01072   0.00364  -0.0589   0.4303   1.0000
   4.000   0.7114   0.01116   0.00392  -0.0583   0.3962   1.0000
   4.250   0.7359   0.01155   0.00421  -0.0579   0.3681   1.0000
   4.500   0.7603   0.01194   0.00451  -0.0574   0.3412   1.0000
   4.750   0.7849   0.01232   0.00482  -0.0569   0.3173   1.0000
   5.000   0.8093   0.01272   0.00514  -0.0564   0.2916   1.0000
   5.250   0.8331   0.01316   0.00550  -0.0559   0.2589   1.0000
   5.500   0.8553   0.01376   0.00589  -0.0552   0.2113   1.0000
   5.750   0.8759   0.01457   0.00638  -0.0544   0.1564   1.0000
   6.000   0.8971   0.01535   0.00694  -0.0536   0.1193   1.0000
   6.250   0.9196   0.01598   0.00750  -0.0529   0.0985   1.0000
   6.500   0.9420   0.01661   0.00807  -0.0522   0.0830   1.0000
   6.750   0.9644   0.01722   0.00866  -0.0516   0.0708   1.0000
   7.000   0.9868   0.01783   0.00929  -0.0509   0.0615   1.0000
   7.250   1.0082   0.01853   0.01001  -0.0500   0.0536   1.0000
   7.500   1.0297   0.01920   0.01073  -0.0492   0.0470   1.0000
   7.750   1.0499   0.02000   0.01156  -0.0482   0.0421   1.0000
   8.000   1.0706   0.02071   0.01235  -0.0472   0.0380   1.0000
   8.250   1.0891   0.02164   0.01329  -0.0460   0.0347   1.0000
   8.500   1.1077   0.02252   0.01428  -0.0448   0.0329   1.0000
   8.750   1.1260   0.02343   0.01529  -0.0436   0.0312   1.0000
   9.000   1.1439   0.02433   0.01629  -0.0423   0.0295   1.0000
   9.250   1.1609   0.02527   0.01728  -0.0410   0.0279   1.0000
   9.500   1.1738   0.02663   0.01867  -0.0393   0.0264   1.0000
   9.750   1.1898   0.02771   0.01993  -0.0378   0.0255   1.0000
  10.000   1.2042   0.02885   0.02122  -0.0362   0.0245   1.0000
  10.250   1.2172   0.02997   0.02247  -0.0344   0.0233   1.0000
  10.500   1.2289   0.03108   0.02368  -0.0326   0.0222   1.0000
  10.750   1.2390   0.03228   0.02497  -0.0307   0.0213   1.0000
  11.000   1.2469   0.03388   0.02665  -0.0288   0.0206   1.0000
  11.250   1.2539   0.03581   0.02871  -0.0269   0.0199   1.0000
  11.500   1.2616   0.03745   0.03060  -0.0251   0.0193   1.0000
  11.750   1.2672   0.03928   0.03265  -0.0234   0.0186   1.0000
  12.000   1.2715   0.04110   0.03467  -0.0219   0.0178   1.0000
  12.250   1.2743   0.04308   0.03682  -0.0205   0.0171   1.0000
  12.500   1.2757   0.04527   0.03916  -0.0193   0.0167   1.0000
  12.750   1.2755   0.04774   0.04177  -0.0183   0.0163   1.0000
  13.000   1.2739   0.05042   0.04459  -0.0176   0.0159   1.0000
  13.250   1.2699   0.05356   0.04786  -0.0172   0.0156   1.0000
  13.500   1.2624   0.05737   0.05184  -0.0172   0.0153   1.0000
  13.750   1.2525   0.06163   0.05633  -0.0177   0.0152   1.0000
  14.000   1.2407   0.06635   0.06130  -0.0189   0.0150   1.0000
  14.250   1.2261   0.07187   0.06708  -0.0210   0.0148   1.0000
  14.500   1.2099   0.07811   0.07356  -0.0239   0.0148   1.0000
  14.750   1.1916   0.08533   0.08102  -0.0280   0.0147   1.0000
  15.000   1.1716   0.09350   0.08941  -0.0330   0.0147   1.0000
  15.250   1.1506   0.10253   0.09864  -0.0387   0.0147   1.0000
  15.500   1.1269   0.11273   0.10903  -0.0453   0.0148   1.0000
  15.750   1.1012   0.12405   0.12051  -0.0526   0.0150   1.0000
  16.000   1.0719   0.13715   0.13374  -0.0610   0.0152   1.0000
<< Back to HQ 2.1/9.5 AIRFOIL (hq2195-il)

Polar data table (+)

Polar graphs


<< Back to HQ 2.1/9.5 AIRFOIL (hq2195-il)