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M6 (65%) (m665-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 50,000
Max Cl/Cd: 33.07 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m665-il-50000-n5.txt
Download as CSV file: xf-m665-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5425   0.11174   0.10453  -0.0007   1.0000   0.1106
  -9.250  -0.5445   0.10820   0.10107  -0.0029   1.0000   0.1130
  -8.750  -0.5574   0.09462   0.08758  -0.0150   1.0000   0.0596
  -8.500  -0.5526   0.09074   0.08372  -0.0151   1.0000   0.0584
  -8.250  -0.5552   0.08667   0.07970  -0.0164   1.0000   0.0577
  -7.750  -0.5583   0.07843   0.07144  -0.0186   1.0000   0.0552
  -7.500  -0.5592   0.07404   0.06701  -0.0198   1.0000   0.0538
  -7.250  -0.5599   0.06940   0.06228  -0.0208   1.0000   0.0522
  -6.750  -0.5607   0.05911   0.05126  -0.0212   1.0000   0.0482
  -6.500  -0.5530   0.05531   0.04720  -0.0201   1.0000   0.0480
  -6.000  -0.5326   0.04894   0.04060  -0.0178   1.0000   0.0505
  -5.750  -0.5207   0.04609   0.03752  -0.0163   1.0000   0.0522
  -5.500  -0.5081   0.04297   0.03404  -0.0146   1.0000   0.0530
  -5.250  -0.4941   0.04008   0.03077  -0.0127   1.0000   0.0546
  -5.000  -0.4788   0.03739   0.02760  -0.0108   1.0000   0.0571
  -4.750  -0.4615   0.03473   0.02444  -0.0088   1.0000   0.0584
  -4.500  -0.4421   0.03226   0.02142  -0.0069   1.0000   0.0594
  -4.250  -0.4208   0.03003   0.01879  -0.0054   1.0000   0.0608
  -4.000  -0.3989   0.02812   0.01674  -0.0043   1.0000   0.0631
  -3.750  -0.3763   0.02670   0.01517  -0.0032   1.0000   0.0677
  -3.500  -0.3515   0.02541   0.01351  -0.0022   1.0000   0.0734
  -3.250  -0.3273   0.02414   0.01231  -0.0016   1.0000   0.0813
  -3.000  -0.2992   0.02299   0.01091  -0.0012   1.0000   0.0887
  -2.750  -0.2737   0.02202   0.00990  -0.0007   1.0000   0.0991
  -2.500  -0.2509   0.02117   0.00903   0.0002   1.0000   0.1150
  -2.250  -0.0707   0.01668   0.00747  -0.0267   1.0000   1.0000
  -2.000  -0.0539   0.01659   0.00712  -0.0249   1.0000   1.0000
  -1.750  -0.0381   0.01654   0.00687  -0.0230   1.0000   1.0000
  -1.500  -0.0236   0.01654   0.00669  -0.0210   1.0000   1.0000
  -1.250  -0.0104   0.01659   0.00661  -0.0187   1.0000   1.0000
  -1.000   0.0006   0.01671   0.00661  -0.0162   1.0000   1.0000
  -0.750   0.0091   0.01690   0.00671  -0.0134   1.0000   1.0000
  -0.500   0.0213   0.01714   0.00685  -0.0115   0.9973   1.0000
  -0.250   0.0677   0.01727   0.00682  -0.0160   0.9785   1.0000
   0.000   0.1133   0.01738   0.00681  -0.0201   0.9607   1.0000
   0.250   0.1586   0.01745   0.00681  -0.0241   0.9434   1.0000
   0.500   0.2036   0.01750   0.00681  -0.0278   0.9265   1.0000
   0.750   0.2468   0.01753   0.00682  -0.0310   0.9091   1.0000
   1.000   0.2858   0.01757   0.00686  -0.0333   0.8903   1.0000
   1.250   0.3209   0.01764   0.00694  -0.0347   0.8707   1.0000
   1.500   0.3542   0.01772   0.00703  -0.0356   0.8519   1.0000
   1.750   0.3835   0.01785   0.00721  -0.0356   0.8324   1.0000
   2.000   0.4101   0.01802   0.00740  -0.0351   0.8128   1.0000
   2.250   0.4361   0.01819   0.00762  -0.0344   0.7941   1.0000
   2.500   0.4608   0.01839   0.00786  -0.0334   0.7757   1.0000
   2.750   0.4836   0.01864   0.00821  -0.0321   0.7562   1.0000
   3.000   0.5067   0.01887   0.00851  -0.0308   0.7372   1.0000
   3.250   0.5298   0.01908   0.00879  -0.0293   0.7183   1.0000
   3.500   0.5511   0.01935   0.00917  -0.0277   0.6968   1.0000
   3.750   0.5732   0.01956   0.00952  -0.0260   0.6758   1.0000
   4.000   0.5948   0.01982   0.00991  -0.0243   0.6535   1.0000
   4.250   0.6170   0.02005   0.01027  -0.0226   0.6313   1.0000
   4.500   0.6377   0.02027   0.01063  -0.0206   0.6044   1.0000
   4.750   0.6554   0.02033   0.01077  -0.0177   0.5642   1.0000
   5.000   0.6696   0.02034   0.01064  -0.0140   0.5017   1.0000
   5.250   0.6825   0.02064   0.01065  -0.0105   0.4163   1.0000
   5.500   0.6937   0.02157   0.01099  -0.0073   0.2920   1.0000
   6.000   0.7122   0.02579   0.01377  -0.0027   0.1041   1.0000
   6.250   0.7251   0.02746   0.01532  -0.0006   0.0867   1.0000
   6.500   0.7395   0.02892   0.01682   0.0013   0.0753   1.0000
   6.750   0.7542   0.03051   0.01844   0.0034   0.0693   1.0000
   7.000   0.7727   0.03195   0.02012   0.0052   0.0643   1.0000
   7.250   0.7906   0.03352   0.02172   0.0066   0.0594   1.0000
   7.500   0.8110   0.03524   0.02360   0.0080   0.0554   1.0000
   7.750   0.8328   0.03715   0.02579   0.0092   0.0529   1.0000
   8.000   0.8534   0.03935   0.02827   0.0105   0.0511   1.0000
   8.250   0.8719   0.04176   0.03102   0.0118   0.0498   1.0000
   8.500   0.8877   0.04438   0.03396   0.0132   0.0487   1.0000
   8.750   0.9007   0.04722   0.03713   0.0147   0.0480   1.0000
   9.000   0.9108   0.05017   0.04037   0.0163   0.0471   1.0000
   9.500   0.9182   0.05680   0.04759   0.0198   0.0453   1.0000
   9.750   0.9129   0.06013   0.05135   0.0219   0.0447   1.0000
  10.000   0.9036   0.06366   0.05522   0.0238   0.0443   1.0000
  10.250   0.8904   0.06722   0.05903   0.0255   0.0441   1.0000
  10.500   0.8735   0.07074   0.06272   0.0270   0.0442   1.0000
  10.750   0.8551   0.07481   0.06693   0.0270   0.0443   1.0000
  11.000   0.8359   0.07968   0.07192   0.0253   0.0445   1.0000
  11.250   0.8176   0.08529   0.07761   0.0224   0.0448   1.0000
  11.500   0.8015   0.09143   0.08379   0.0189   0.0451   1.0000
  11.750   0.7882   0.09785   0.09023   0.0152   0.0455   1.0000
  12.000   0.7753   0.10486   0.09723   0.0111   0.0457   1.0000
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