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

NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA M9 AIRFOIL (m9-il)
Reynolds number: 500,000
Max Cl/Cd: 80.37 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m9-il-500000.txt
Download as CSV file: xf-m9-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M9 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.2421   0.08898   0.08592  -0.0073   0.6866   0.0276
  -7.500  -0.2353   0.08654   0.08350  -0.0081   0.6784   0.0283
  -7.250  -0.2331   0.08399   0.08093  -0.0090   0.6711   0.0299
  -6.500  -0.2720   0.02335   0.01877  -0.0598   0.6695   0.0258
  -6.250  -0.2472   0.02000   0.01501  -0.0610   0.6608   0.0270
  -6.000  -0.2200   0.01961   0.01457  -0.0608   0.6518   0.0280
  -5.750  -0.1932   0.01847   0.01322  -0.0607   0.6413   0.0294
  -5.500  -0.1665   0.01725   0.01167  -0.0605   0.6313   0.0307
  -5.250  -0.1409   0.01603   0.01023  -0.0601   0.6204   0.0323
  -5.000  -0.1133   0.01570   0.00985  -0.0598   0.6087   0.0337
  -4.750  -0.0856   0.01529   0.00927  -0.0594   0.5970   0.0357
  -4.500  -0.0589   0.01449   0.00823  -0.0590   0.5854   0.0377
  -4.250  -0.0313   0.01419   0.00789  -0.0586   0.5732   0.0394
  -4.000  -0.0036   0.01389   0.00746  -0.0582   0.5617   0.0415
  -3.750   0.0244   0.01369   0.00707  -0.0578   0.5507   0.0434
  -3.500   0.0513   0.01307   0.00638  -0.0573   0.5402   0.0460
  -3.250   0.0792   0.01287   0.00610  -0.0570   0.5310   0.0483
  -3.000   0.1073   0.01272   0.00582  -0.0566   0.5222   0.0506
  -2.750   0.1347   0.01221   0.00521  -0.0561   0.5141   0.0533
  -2.500   0.1626   0.01206   0.00501  -0.0558   0.5055   0.0560
  -2.250   0.1907   0.01192   0.00478  -0.0554   0.4976   0.0589
  -2.000   0.2185   0.01160   0.00438  -0.0550   0.4900   0.0619
  -1.750   0.2464   0.01149   0.00422  -0.0546   0.4835   0.0657
  -1.500   0.2747   0.01136   0.00407  -0.0543   0.4769   0.0695
  -1.250   0.3026   0.01116   0.00380  -0.0539   0.4708   0.0740
  -1.000   0.3307   0.01111   0.00373  -0.0536   0.4651   0.0793
  -0.750   0.3589   0.01099   0.00358  -0.0533   0.4591   0.0846
  -0.500   0.3868   0.01090   0.00345  -0.0529   0.4534   0.0915
  -0.250   0.4149   0.01085   0.00336  -0.0526   0.4486   0.0980
   0.000   0.4429   0.01069   0.00324  -0.0523   0.4440   0.1074
   0.250   0.4707   0.01061   0.00313  -0.0520   0.4394   0.1163
   0.500   0.4983   0.01059   0.00307  -0.0516   0.4343   0.1266
   0.750   0.5263   0.01048   0.00305  -0.0513   0.4291   0.1430
   1.000   0.5555   0.01110   0.00386  -0.0511   0.4239   0.2091
   1.250   0.5837   0.01165   0.00430  -0.0507   0.4188   0.2231
   1.500   0.6119   0.01210   0.00471  -0.0503   0.4145   0.2326
   1.750   0.6398   0.01234   0.00503  -0.0500   0.4099   0.2378
   2.000   0.6678   0.01268   0.00528  -0.0496   0.4053   0.2451
   2.250   0.6948   0.01276   0.00540  -0.0493   0.4008   0.2510
   2.500   0.7231   0.01296   0.00558  -0.0491   0.3964   0.2581
   2.750   0.7503   0.01283   0.00553  -0.0489   0.3913   0.2633
   3.000   0.7778   0.01301   0.00563  -0.0486   0.3859   0.2685
   3.250   0.8054   0.01290   0.00554  -0.0484   0.3806   0.2722
   3.500   0.8327   0.01287   0.00555  -0.0482   0.3747   0.2769
   3.750   0.8600   0.01297   0.00559  -0.0479   0.3688   0.2800
   4.000   0.8879   0.01304   0.00568  -0.0477   0.3627   0.2820
   4.250   0.9149   0.01291   0.00554  -0.0475   0.3555   0.2855
   4.500   0.9421   0.01289   0.00555  -0.0473   0.3481   0.2880
   4.750   0.9692   0.01292   0.00557  -0.0471   0.3388   0.2894
   5.000   0.9963   0.01298   0.00563  -0.0468   0.3283   0.2913
   5.250   1.0230   0.01308   0.00569  -0.0466   0.3151   0.2920
   5.500   1.0492   0.01327   0.00582  -0.0463   0.2975   0.2938
   5.750   1.0746   0.01355   0.00598  -0.0460   0.2756   0.2942
   6.000   1.0995   0.01384   0.00617  -0.0457   0.2565   0.2931
   6.250   1.1242   0.01416   0.00641  -0.0453   0.2431   0.2920
   6.500   1.1491   0.01444   0.00665  -0.0449   0.2344   0.2910
   6.750   1.1737   0.01476   0.00694  -0.0445   0.2276   0.2909
   7.000   1.1988   0.01498   0.00719  -0.0441   0.2230   0.2902
   7.250   1.2232   0.01528   0.00749  -0.0437   0.2187   0.2892
   7.500   1.2465   0.01567   0.00786  -0.0431   0.2144   0.2883
   7.750   1.2713   0.01590   0.00814  -0.0427   0.2118   0.2875
   8.000   1.2958   0.01613   0.00844  -0.0423   0.2091   0.2869
   8.250   1.3197   0.01642   0.00876  -0.0418   0.2062   0.2863
   8.500   1.3427   0.01678   0.00913  -0.0412   0.2032   0.2859
   8.750   1.3644   0.01723   0.00959  -0.0405   0.2000   0.2855
   9.000   1.3866   0.01762   0.01003  -0.0398   0.1974   0.2852
   9.250   1.4102   0.01783   0.01034  -0.0393   0.1951   0.2850
   9.500   1.4332   0.01808   0.01065  -0.0388   0.1918   0.2848
   9.750   1.4545   0.01843   0.01103  -0.0380   0.1878   0.2847
  10.000   1.4732   0.01897   0.01158  -0.0370   0.1839   0.2845
  10.250   1.4943   0.01929   0.01199  -0.0363   0.1815   0.2845
  10.500   1.5157   0.01955   0.01234  -0.0356   0.1791   0.2846
  10.750   1.5358   0.01987   0.01275  -0.0347   0.1763   0.2846
  11.000   1.5538   0.02029   0.01322  -0.0336   0.1732   0.2847
  11.250   1.5678   0.02090   0.01384  -0.0321   0.1694   0.2849
  11.500   1.5823   0.02130   0.01434  -0.0305   0.1664   0.2852
  11.750   1.5956   0.02173   0.01488  -0.0288   0.1636   0.2854
  12.000   1.6059   0.02238   0.01560  -0.0270   0.1600   0.2857
  12.250   1.6132   0.02333   0.01658  -0.0253   0.1566   0.2864
  12.500   1.6218   0.02434   0.01766  -0.0240   0.1527   0.2870
  12.750   1.6334   0.02525   0.01865  -0.0231   0.1470   0.2877
  13.000   1.6382   0.02676   0.02019  -0.0222   0.1417   0.2882
  13.250   1.6468   0.02806   0.02157  -0.0214   0.1355   0.2888
  13.500   1.6483   0.03005   0.02355  -0.0207   0.1271   0.2887
  13.750   1.6459   0.03245   0.02594  -0.0201   0.1145   0.2893
  14.000   1.6316   0.03611   0.02951  -0.0196   0.0956   0.2892
  14.250   1.6069   0.04106   0.03437  -0.0195   0.0786   0.2889
  14.500   1.5906   0.04540   0.03873  -0.0197   0.0707   0.2892
  14.750   1.5781   0.04952   0.04289  -0.0201   0.0659   0.2894
  15.000   1.5645   0.05392   0.04735  -0.0209   0.0618   0.2895
  15.250   1.5546   0.05812   0.05163  -0.0217   0.0580   0.2896
  15.500   1.5444   0.06251   0.05609  -0.0228   0.0554   0.2897
  15.750   1.5300   0.06763   0.06128  -0.0243   0.0519   0.2901
  16.000   1.5229   0.07185   0.06558  -0.0255   0.0477   0.2903
  16.250   1.5134   0.07653   0.07036  -0.0270   0.0454   0.2905
  16.500   1.5017   0.08166   0.07555  -0.0288   0.0409   0.2909
  16.750   1.4820   0.08821   0.08213  -0.0314   0.0311   0.2911
  17.000   1.4621   0.09502   0.08897  -0.0341   0.0243   0.2907
<< Back to NACA M9 AIRFOIL (m9-il)

Polar data table (+)

Polar graphs


<< Back to NACA M9 AIRFOIL (m9-il)