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

NPL 9626 AIRFOIL (npl9626-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NPL 9626 AIRFOIL (npl9626-il)
Reynolds number: 50,000
Max Cl/Cd: 29.62 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-npl9626-il-50000-n5.txt
Download as CSV file: xf-npl9626-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NPL 9626 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.6626   0.09285   0.08516  -0.0215   1.0000   0.0566
 -10.750  -0.6979   0.08201   0.07424  -0.0299   1.0000   0.0550
 -10.500  -0.7408   0.07346   0.06549  -0.0352   1.0000   0.0536
 -10.250  -0.7725   0.06830   0.06009  -0.0352   1.0000   0.0530
 -10.000  -0.7838   0.06479   0.05641  -0.0341   1.0000   0.0535
  -9.750  -0.7893   0.06150   0.05294  -0.0329   1.0000   0.0542
  -9.500  -0.7927   0.05823   0.04942  -0.0315   1.0000   0.0552
  -9.250  -0.7939   0.05492   0.04579  -0.0298   1.0000   0.0562
  -9.000  -0.7922   0.05162   0.04211  -0.0281   1.0000   0.0573
  -8.750  -0.7866   0.04840   0.03846  -0.0263   1.0000   0.0582
  -8.500  -0.7776   0.04537   0.03495  -0.0245   1.0000   0.0593
  -8.250  -0.7647   0.04261   0.03176  -0.0228   1.0000   0.0607
  -8.000  -0.7468   0.04055   0.02966  -0.0219   1.0000   0.0629
  -7.750  -0.7289   0.03868   0.02760  -0.0207   1.0000   0.0660
  -7.500  -0.7108   0.03681   0.02530  -0.0192   1.0000   0.0699
  -7.250  -0.6909   0.03513   0.02369  -0.0183   1.0000   0.0739
  -7.000  -0.6705   0.03366   0.02203  -0.0171   1.0000   0.0792
  -6.750  -0.6491   0.03210   0.02041  -0.0161   1.0000   0.0838
  -6.500  -0.6278   0.03074   0.01901  -0.0149   1.0000   0.0888
  -6.250  -0.6069   0.02948   0.01765  -0.0136   1.0000   0.0951
  -6.000  -0.5878   0.02826   0.01644  -0.0121   1.0000   0.1019
  -5.750  -0.5692   0.02709   0.01524  -0.0105   1.0000   0.1100
  -5.500  -0.5514   0.02597   0.01414  -0.0089   1.0000   0.1222
  -5.250  -0.5346   0.02476   0.01309  -0.0073   1.0000   0.1393
  -5.000  -0.5183   0.02346   0.01209  -0.0057   1.0000   0.1733
  -4.750  -0.5041   0.02206   0.01134  -0.0040   1.0000   0.2592
  -4.500  -0.4903   0.02110   0.01093  -0.0018   1.0000   0.3658
  -4.250  -0.4753   0.02051   0.01064   0.0006   1.0000   0.4418
  -4.000  -0.4590   0.02008   0.01043   0.0031   1.0000   0.5010
  -3.750  -0.4423   0.01977   0.01032   0.0056   1.0000   0.5580
  -3.500  -0.4268   0.01957   0.01043   0.0088   1.0000   0.6246
  -3.250  -0.4100   0.01950   0.01055   0.0120   1.0000   0.6836
  -3.000  -0.3896   0.01949   0.01061   0.0143   1.0000   0.7262
  -2.750  -0.3671   0.01947   0.01058   0.0160   1.0000   0.7602
  -2.500  -0.3431   0.01951   0.01059   0.0172   1.0000   0.7939
  -2.250  -0.3163   0.01962   0.01063   0.0179   1.0000   0.8274
  -2.000  -0.2859   0.01978   0.01071   0.0177   1.0000   0.8601
  -1.750  -0.2518   0.01998   0.01081   0.0166   1.0000   0.8916
  -1.500  -0.2015   0.02022   0.01090   0.0121   0.9959   0.9175
  -1.250  -0.1386   0.02042   0.01093   0.0050   0.9852   0.9398
  -1.000  -0.0719   0.02051   0.01088  -0.0030   0.9743   0.9569
  -0.750  -0.0049   0.02048   0.01074  -0.0112   0.9622   0.9720
  -0.500   0.0615   0.02034   0.01053  -0.0193   0.9489   0.9853
  -0.250   0.1254   0.02010   0.01025  -0.0269   0.9324   0.9974
   0.000   0.1706   0.01988   0.00999  -0.0309   0.9105   1.0000
   0.250   0.2058   0.01969   0.00977  -0.0328   0.8852   1.0000
   0.500   0.2387   0.01952   0.00956  -0.0339   0.8599   1.0000
   0.750   0.2675   0.01941   0.00941  -0.0341   0.8343   1.0000
   1.000   0.2941   0.01934   0.00929  -0.0338   0.8094   1.0000
   1.250   0.3189   0.01933   0.00921  -0.0330   0.7855   1.0000
   1.500   0.3418   0.01938   0.00921  -0.0319   0.7618   1.0000
   1.750   0.3648   0.01945   0.00923  -0.0308   0.7402   1.0000
   2.000   0.3876   0.01956   0.00927  -0.0296   0.7200   1.0000
   2.250   0.4094   0.01973   0.00941  -0.0284   0.6998   1.0000
   2.500   0.4315   0.01994   0.00960  -0.0272   0.6809   1.0000
   2.750   0.4536   0.02017   0.00981  -0.0260   0.6633   1.0000
   3.000   0.4760   0.02043   0.01006  -0.0249   0.6469   1.0000
   3.250   0.4988   0.02071   0.01036  -0.0238   0.6314   1.0000
   3.500   0.5223   0.02100   0.01067  -0.0229   0.6167   1.0000
   3.750   0.5449   0.02135   0.01108  -0.0219   0.6014   1.0000
   4.000   0.5675   0.02171   0.01152  -0.0209   0.5865   1.0000
   4.250   0.5897   0.02206   0.01194  -0.0196   0.5704   1.0000
   4.500   0.6093   0.02219   0.01202  -0.0175   0.5444   1.0000
   4.750   0.6251   0.02223   0.01192  -0.0146   0.5065   1.0000
   5.000   0.6399   0.02231   0.01181  -0.0116   0.4638   1.0000
   5.250   0.6566   0.02257   0.01201  -0.0093   0.4279   1.0000
   5.500   0.6742   0.02291   0.01235  -0.0074   0.3941   1.0000
   5.750   0.6907   0.02332   0.01278  -0.0053   0.3496   1.0000
   6.000   0.7049   0.02393   0.01320  -0.0030   0.2773   1.0000
   6.250   0.7143   0.02532   0.01392  -0.0005   0.1929   1.0000
   6.500   0.7254   0.02696   0.01514   0.0015   0.1533   1.0000
   6.750   0.7387   0.02840   0.01645   0.0033   0.1353   1.0000
   7.000   0.7527   0.02976   0.01774   0.0051   0.1238   1.0000
   7.250   0.7665   0.03112   0.01901   0.0068   0.1152   1.0000
   7.500   0.7834   0.03233   0.02034   0.0083   0.1081   1.0000
   7.750   0.7992   0.03371   0.02166   0.0098   0.1026   1.0000
   8.000   0.8183   0.03498   0.02309   0.0110   0.0969   1.0000
   8.250   0.8373   0.03635   0.02449   0.0122   0.0920   1.0000
   8.500   0.8578   0.03789   0.02603   0.0130   0.0878   1.0000
   8.750   0.8782   0.03945   0.02783   0.0140   0.0837   1.0000
   9.000   0.8986   0.04113   0.02962   0.0148   0.0804   1.0000
   9.250   0.9201   0.04300   0.03143   0.0153   0.0772   1.0000
   9.500   0.9346   0.04505   0.03387   0.0166   0.0745   1.0000
   9.750   0.9490   0.04731   0.03644   0.0178   0.0723   1.0000
  10.000   0.9621   0.04975   0.03914   0.0190   0.0707   1.0000
  10.250   0.9733   0.05219   0.04180   0.0202   0.0691   1.0000
  10.500   0.9849   0.05460   0.04431   0.0212   0.0674   1.0000
  10.750   0.9884   0.05748   0.04742   0.0226   0.0661   1.0000
  11.000   0.9793   0.06070   0.05105   0.0247   0.0652   1.0000
  11.250   0.9650   0.06398   0.05464   0.0269   0.0646   1.0000
  11.500   0.9474   0.06774   0.05868   0.0282   0.0643   1.0000
  11.750   0.9267   0.07221   0.06340   0.0281   0.0641   1.0000
  12.000   0.9021   0.07769   0.06911   0.0263   0.0642   1.0000
  12.250   0.8737   0.08452   0.07614   0.0228   0.0645   1.0000
  12.500   0.8421   0.09303   0.08480   0.0174   0.0650   1.0000
<< Back to NPL 9626 AIRFOIL (npl9626-il)

Polar data table (+)

Polar graphs


<< Back to NPL 9626 AIRFOIL (npl9626-il)