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M6 (65%) (m665-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 100,000
Max Cl/Cd: 44.76 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m665-il-100000-n5.txt
Download as CSV file: xf-m665-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5483   0.08928   0.08423  -0.0127   1.0000   0.0341
  -8.500  -0.5497   0.08524   0.08024  -0.0144   1.0000   0.0328
  -8.250  -0.5561   0.08075   0.07579  -0.0169   1.0000   0.0319
  -8.000  -0.5625   0.07648   0.07153  -0.0183   1.0000   0.0309
  -7.750  -0.5666   0.07165   0.06667  -0.0201   1.0000   0.0299
  -7.500  -0.5703   0.06624   0.06117  -0.0215   1.0000   0.0286
  -7.250  -0.5776   0.05743   0.05197  -0.0227   1.0000   0.0261
  -6.750  -0.5659   0.04941   0.04343  -0.0204   1.0000   0.0255
  -6.500  -0.5596   0.04552   0.03926  -0.0187   1.0000   0.0254
  -6.250  -0.5518   0.04183   0.03522  -0.0167   1.0000   0.0254
  -6.000  -0.5422   0.03841   0.03141  -0.0144   1.0000   0.0255
  -5.750  -0.5320   0.03501   0.02757  -0.0120   1.0000   0.0258
  -5.500  -0.5198   0.03205   0.02428  -0.0098   1.0000   0.0264
  -5.250  -0.5044   0.03029   0.02236  -0.0082   1.0000   0.0279
  -5.000  -0.4872   0.02876   0.02063  -0.0065   1.0000   0.0301
  -4.750  -0.4693   0.02687   0.01837  -0.0045   1.0000   0.0323
  -4.500  -0.4498   0.02532   0.01636  -0.0026   1.0000   0.0353
  -4.250  -0.4313   0.02351   0.01435  -0.0011   1.0000   0.0385
  -4.000  -0.4111   0.02227   0.01296   0.0004   1.0000   0.0409
  -3.750  -0.3899   0.02092   0.01143   0.0018   1.0000   0.0424
  -3.500  -0.3690   0.01979   0.01015   0.0033   1.0000   0.0440
  -3.250  -0.3492   0.01883   0.00909   0.0050   1.0000   0.0460
  -3.000  -0.3289   0.01804   0.00820   0.0064   0.9992   0.0481
  -2.750  -0.2926   0.01707   0.00725   0.0043   0.9893   0.0530
  -2.500  -0.2560   0.01643   0.00659   0.0022   0.9786   0.0631
  -2.250  -0.2202   0.01579   0.00594   0.0005   0.9674   0.0758
  -2.000  -0.1850   0.01505   0.00534  -0.0012   0.9558   0.1057
  -1.750  -0.1504   0.01411   0.00490  -0.0029   0.9443   0.2158
  -1.500  -0.0445   0.01238   0.00563  -0.0165   0.9751   0.9732
  -1.250   0.0261   0.01231   0.00529  -0.0253   0.9791   1.0000
  -1.000   0.0690   0.01223   0.00505  -0.0287   0.9572   1.0000
  -0.750   0.1121   0.01215   0.00481  -0.0321   0.9354   1.0000
  -0.500   0.1530   0.01206   0.00460  -0.0349   0.9120   1.0000
  -0.250   0.1905   0.01200   0.00441  -0.0369   0.8880   1.0000
   0.000   0.2233   0.01196   0.00426  -0.0378   0.8641   1.0000
   0.250   0.2509   0.01197   0.00416  -0.0376   0.8404   1.0000
   0.500   0.2768   0.01200   0.00410  -0.0371   0.8191   1.0000
   0.750   0.3013   0.01205   0.00408  -0.0362   0.7984   1.0000
   1.000   0.3254   0.01212   0.00407  -0.0353   0.7796   1.0000
   1.250   0.3499   0.01219   0.00409  -0.0345   0.7620   1.0000
   1.500   0.3738   0.01228   0.00414  -0.0336   0.7442   1.0000
   1.750   0.3975   0.01238   0.00423  -0.0327   0.7271   1.0000
   2.000   0.4210   0.01249   0.00432  -0.0317   0.7103   1.0000
   2.250   0.4443   0.01261   0.00444  -0.0306   0.6934   1.0000
   2.500   0.4674   0.01275   0.00455  -0.0294   0.6756   1.0000
   2.750   0.4901   0.01288   0.00471  -0.0282   0.6564   1.0000
   3.000   0.5129   0.01303   0.00487  -0.0271   0.6368   1.0000
   3.250   0.5357   0.01319   0.00505  -0.0259   0.6181   1.0000
   3.500   0.5585   0.01336   0.00525  -0.0247   0.5989   1.0000
   3.750   0.5808   0.01353   0.00548  -0.0234   0.5754   1.0000
   4.000   0.6021   0.01370   0.00565  -0.0218   0.5446   1.0000
   4.250   0.6220   0.01393   0.00576  -0.0200   0.4984   1.0000
   4.500   0.6400   0.01430   0.00587  -0.0179   0.4264   1.0000
   4.750   0.6574   0.01492   0.00614  -0.0159   0.3402   1.0000
   5.000   0.6717   0.01605   0.00663  -0.0139   0.2158   1.0000
   5.250   0.6839   0.01773   0.00753  -0.0119   0.0945   1.0000
   5.500   0.7002   0.01894   0.00855  -0.0101   0.0629   1.0000
   5.750   0.7180   0.01988   0.00951  -0.0084   0.0507   1.0000
   6.000   0.7352   0.02087   0.01052  -0.0067   0.0443   1.0000
   6.250   0.7527   0.02180   0.01159  -0.0049   0.0409   1.0000
   6.500   0.7696   0.02279   0.01268  -0.0031   0.0384   1.0000
   6.750   0.7857   0.02393   0.01389  -0.0012   0.0365   1.0000
   7.000   0.8013   0.02536   0.01533   0.0007   0.0351   1.0000
   7.250   0.8192   0.02684   0.01688   0.0023   0.0342   1.0000
   7.500   0.8387   0.02816   0.01837   0.0037   0.0330   1.0000
   7.750   0.8579   0.02952   0.01993   0.0051   0.0313   1.0000
   8.000   0.8763   0.03102   0.02165   0.0065   0.0297   1.0000
   8.250   0.8943   0.03286   0.02371   0.0080   0.0290   1.0000
   8.500   0.9109   0.03489   0.02600   0.0095   0.0284   1.0000
   8.750   0.9256   0.03708   0.02848   0.0112   0.0279   1.0000
   9.000   0.9378   0.03947   0.03119   0.0130   0.0275   1.0000
   9.250   0.9472   0.04205   0.03412   0.0150   0.0272   1.0000
   9.500   0.9533   0.04484   0.03726   0.0172   0.0270   1.0000
   9.750   0.9556   0.04782   0.04059   0.0195   0.0269   1.0000
  10.000   0.9537   0.05099   0.04409   0.0220   0.0269   1.0000
  10.250   0.9471   0.05424   0.04766   0.0245   0.0269   1.0000
  10.500   0.9344   0.05755   0.05125   0.0272   0.0270   1.0000
  10.750   0.9150   0.06082   0.05474   0.0298   0.0271   1.0000
  11.000   0.8927   0.06486   0.05900   0.0305   0.0273   1.0000
  11.250   0.8678   0.07003   0.06437   0.0289   0.0275   1.0000
  11.500   0.8405   0.07676   0.07126   0.0249   0.0279   1.0000
  11.750   0.8095   0.08600   0.08061   0.0181   0.0284   1.0000
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