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M6 (65%) (m665-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 50,000
Max Cl/Cd: 33.8 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m665-il-50000.txt
Download as CSV file: xf-m665-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: M6 (65%)                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5301   0.09891   0.09206   0.0043   1.0000   0.2580
  -8.000  -0.5302   0.09578   0.08901   0.0045   1.0000   0.2732
  -7.750  -0.5361   0.09301   0.08633   0.0046   1.0000   0.2891
  -7.500  -0.5202   0.08871   0.08205   0.0066   1.0000   0.3100
  -7.250  -0.5204   0.08583   0.07925   0.0077   1.0000   0.3310
  -7.000  -0.5145   0.08240   0.07587   0.0094   1.0000   0.3537
  -6.750  -0.5338   0.08074   0.07436   0.0112   1.0000   0.3780
  -6.500  -0.5007   0.07640   0.06997   0.0149   1.0000   0.4186
  -6.250  -0.5047   0.07448   0.06816   0.0198   1.0000   0.4661
  -6.000  -0.4369   0.06959   0.06304   0.0251   1.0000   0.5637
  -4.750  -0.4779   0.04334   0.03483  -0.0104   1.0000   0.1600
  -4.500  -0.4598   0.03981   0.03105  -0.0088   1.0000   0.1511
  -4.250  -0.4412   0.03666   0.02726  -0.0068   1.0000   0.1410
  -4.000  -0.4221   0.03382   0.02402  -0.0050   1.0000   0.1371
  -3.750  -0.4024   0.03156   0.02131  -0.0031   1.0000   0.1390
  -3.500  -0.3813   0.02960   0.01887  -0.0013   1.0000   0.1427
  -3.250  -0.3579   0.02760   0.01648   0.0002   1.0000   0.1449
  -3.000  -0.3335   0.02570   0.01453   0.0012   1.0000   0.1517
  -2.750  -0.1073   0.01694   0.00865  -0.0300   1.0000   1.0000
  -2.500  -0.0879   0.01680   0.00794  -0.0284   1.0000   1.0000
  -2.250  -0.0702   0.01668   0.00745  -0.0267   1.0000   1.0000
  -2.000  -0.0535   0.01660   0.00712  -0.0250   1.0000   1.0000
  -1.750  -0.0379   0.01655   0.00687  -0.0230   1.0000   1.0000
  -1.500  -0.0234   0.01655   0.00671  -0.0210   1.0000   1.0000
  -1.250  -0.0104   0.01661   0.00663  -0.0187   1.0000   1.0000
  -1.000   0.0005   0.01673   0.00663  -0.0162   1.0000   1.0000
  -0.750   0.0089   0.01692   0.00673  -0.0133   1.0000   1.0000
  -0.500   0.0151   0.01720   0.00691  -0.0103   1.0000   1.0000
  -0.250   0.0207   0.01755   0.00716  -0.0073   1.0000   1.0000
   0.000   0.0265   0.01794   0.00746  -0.0045   1.0000   1.0000
   0.250   0.0332   0.01838   0.00780  -0.0020   1.0000   1.0000
   0.500   0.0410   0.01885   0.00818   0.0003   1.0000   1.0000
   0.750   0.0496   0.01936   0.00861   0.0022   1.0000   1.0000
   1.000   0.0590   0.01991   0.00909   0.0040   1.0000   1.0000
   1.250   0.0691   0.02049   0.00962   0.0055   1.0000   1.0000
   1.500   0.0936   0.02123   0.01033   0.0041   0.9953   1.0000
   1.750   0.1534   0.02219   0.01133  -0.0040   0.9763   1.0000
   2.000   0.2128   0.02308   0.01229  -0.0115   0.9576   1.0000
   2.250   0.2662   0.02381   0.01314  -0.0177   0.9381   1.0000
   2.500   0.3188   0.02444   0.01392  -0.0234   0.9178   1.0000
   2.750   0.3822   0.02492   0.01466  -0.0306   0.8985   1.0000
   3.000   0.4305   0.02535   0.01530  -0.0348   0.8766   1.0000
   3.250   0.4942   0.02538   0.01564  -0.0408   0.8549   1.0000
   3.500   0.5329   0.02562   0.01616  -0.0422   0.8300   1.0000
   3.750   0.5731   0.02570   0.01650  -0.0431   0.8053   1.0000
   4.000   0.6127   0.02557   0.01665  -0.0432   0.7810   1.0000
   4.250   0.6365   0.02583   0.01717  -0.0411   0.7543   1.0000
   4.500   0.6629   0.02558   0.01714  -0.0380   0.7240   1.0000
   4.750   0.6843   0.02452   0.01618  -0.0322   0.6810   1.0000
   5.000   0.7004   0.02319   0.01486  -0.0253   0.6280   1.0000
   5.250   0.7140   0.02208   0.01375  -0.0189   0.5676   1.0000
   5.500   0.7195   0.02129   0.01272  -0.0115   0.4631   1.0000
   5.750   0.7096   0.02384   0.01324  -0.0039   0.2331   1.0000
   6.000   0.7223   0.02645   0.01519  -0.0011   0.1732   1.0000
   6.250   0.7428   0.02851   0.01702   0.0007   0.1469   1.0000
   6.500   0.7684   0.03061   0.01909   0.0019   0.1326   1.0000
   6.750   0.7933   0.03284   0.02126   0.0028   0.1213   1.0000
   7.000   0.8170   0.03516   0.02404   0.0042   0.1160   1.0000
   7.250   0.8394   0.03787   0.02710   0.0056   0.1132   1.0000
   7.500   0.8589   0.04090   0.03051   0.0071   0.1119   1.0000
   7.750   0.8759   0.04407   0.03401   0.0087   0.1103   1.0000
   8.000   0.8919   0.04786   0.03791   0.0099   0.1081   1.0000
   8.250   0.9007   0.05135   0.04190   0.0119   0.1077   1.0000
   8.500   0.9057   0.05502   0.04613   0.0141   0.1090   1.0000
   8.750   0.8905   0.05981   0.05175   0.0168   0.1136   1.0000
   9.000   0.8816   0.06473   0.05702   0.0182   0.1171   1.0000
   9.250   0.8830   0.06962   0.06206   0.0191   0.1203   1.0000
   9.500   0.8458   0.07495   0.06771   0.0194   0.1251   1.0000
   9.750   0.8063   0.08077   0.07357   0.0184   0.1287   1.0000
  10.000   0.8129   0.08634   0.07918   0.0181   0.1359   1.0000
  10.250   0.7612   0.09784   0.09058   0.0079   0.1501   1.0000
  10.500   0.6775   0.12250   0.11489  -0.0192   0.3049   1.0000
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