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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 200,000
Max Cl/Cd: 32.16 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m1-il-200000-n5.txt
Download as CSV file: xf-m1-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-M1 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7323   0.08747   0.08413   0.0130   1.0000   0.0114
  -8.750  -0.7396   0.08070   0.07741   0.0069   1.0000   0.0113
  -8.500  -0.7485   0.07302   0.06969  -0.0001   1.0000   0.0110
  -8.250  -0.7526   0.06570   0.06224  -0.0046   1.0000   0.0109
  -8.000  -0.7532   0.05884   0.05517  -0.0072   1.0000   0.0111
  -7.750  -0.7505   0.05221   0.04826  -0.0086   1.0000   0.0114
  -7.500  -0.7449   0.04560   0.04125  -0.0090   1.0000   0.0117
  -7.250  -0.7369   0.03905   0.03417  -0.0085   1.0000   0.0119
  -7.000  -0.7245   0.03335   0.02784  -0.0076   1.0000   0.0122
  -6.750  -0.7074   0.02890   0.02276  -0.0065   1.0000   0.0128
  -6.500  -0.6865   0.02567   0.01897  -0.0056   1.0000   0.0136
  -6.250  -0.6647   0.02296   0.01583  -0.0049   1.0000   0.0152
  -6.000  -0.6407   0.02189   0.01457  -0.0046   1.0000   0.0169
  -5.750  -0.6159   0.02038   0.01281  -0.0041   1.0000   0.0188
  -5.500  -0.5905   0.01903   0.01119  -0.0035   1.0000   0.0220
  -5.250  -0.5643   0.01829   0.01022  -0.0031   1.0000   0.0252
  -5.000  -0.5400   0.01679   0.00862  -0.0026   1.0000   0.0283
  -4.750  -0.5144   0.01618   0.00792  -0.0023   1.0000   0.0329
  -4.500  -0.4885   0.01552   0.00712  -0.0019   1.0000   0.0360
  -4.250  -0.4637   0.01462   0.00618  -0.0015   1.0000   0.0398
  -4.000  -0.4382   0.01399   0.00550  -0.0012   1.0000   0.0420
  -3.750  -0.4123   0.01351   0.00497  -0.0009   1.0000   0.0450
  -3.500  -0.3864   0.01301   0.00437  -0.0005   1.0000   0.0463
  -3.250  -0.3604   0.01257   0.00383  -0.0002   1.0000   0.0473
  -3.000  -0.3342   0.01220   0.00339   0.0002   1.0000   0.0487
  -2.750  -0.3082   0.01179   0.00294   0.0005   1.0000   0.0536
  -2.500  -0.2823   0.01141   0.00261   0.0008   1.0000   0.0676
  -2.250  -0.2575   0.01071   0.00229   0.0011   1.0000   0.1534
  -2.000  -0.2335   0.00990   0.00205   0.0013   1.0000   0.2920
  -1.750  -0.2113   0.00894   0.00189   0.0020   1.0000   0.4870
  -1.500  -0.1898   0.00825   0.00185   0.0033   1.0000   0.6374
  -1.250  -0.1697   0.00777   0.00191   0.0054   1.0000   0.7690
  -1.000  -0.1466   0.00757   0.00205   0.0075   1.0000   0.8900
  -0.750  -0.0899   0.00764   0.00219   0.0021   1.0000   0.9802
  -0.500  -0.0410   0.00767   0.00217  -0.0026   1.0000   1.0000
  -0.250  -0.0202   0.00763   0.00211  -0.0014   1.0000   1.0000
   0.000   0.0000   0.00762   0.00209   0.0000   1.0000   1.0000
   0.250   0.0202   0.00763   0.00211   0.0014   1.0000   1.0000
   0.500   0.0410   0.00767   0.00217   0.0026   1.0000   1.0000
   0.750   0.0904   0.00764   0.00218  -0.0022   0.9797   1.0000
   1.000   0.1467   0.00757   0.00205  -0.0075   0.8894   1.0000
   1.250   0.1697   0.00777   0.00190  -0.0054   0.7675   1.0000
   1.500   0.1899   0.00825   0.00185  -0.0033   0.6375   1.0000
   1.750   0.2114   0.00894   0.00189  -0.0020   0.4871   1.0000
   2.000   0.2336   0.00990   0.00205  -0.0014   0.2920   1.0000
   2.250   0.2575   0.01072   0.00229  -0.0011   0.1520   1.0000
   2.500   0.2823   0.01141   0.00261  -0.0008   0.0671   1.0000
   2.750   0.3083   0.01179   0.00294  -0.0005   0.0535   1.0000
   3.000   0.3343   0.01220   0.00339  -0.0002   0.0486   1.0000
   3.250   0.3605   0.01257   0.00384   0.0002   0.0472   1.0000
   3.500   0.3865   0.01301   0.00437   0.0005   0.0463   1.0000
   3.750   0.4123   0.01352   0.00497   0.0009   0.0450   1.0000
   4.000   0.4382   0.01400   0.00551   0.0012   0.0420   1.0000
   4.250   0.4637   0.01462   0.00618   0.0015   0.0398   1.0000
   4.500   0.4885   0.01552   0.00712   0.0019   0.0360   1.0000
   4.750   0.5144   0.01618   0.00793   0.0023   0.0329   1.0000
   5.000   0.5400   0.01679   0.00862   0.0026   0.0283   1.0000
   5.250   0.5644   0.01829   0.01021   0.0031   0.0252   1.0000
   5.500   0.5905   0.01903   0.01120   0.0035   0.0220   1.0000
   5.750   0.6159   0.02038   0.01280   0.0041   0.0188   1.0000
   6.000   0.6407   0.02188   0.01456   0.0046   0.0169   1.0000
   6.250   0.6647   0.02297   0.01584   0.0049   0.0152   1.0000
   6.500   0.6865   0.02567   0.01897   0.0056   0.0136   1.0000
   6.750   0.7074   0.02890   0.02276   0.0065   0.0128   1.0000
   7.000   0.7245   0.03334   0.02784   0.0076   0.0122   1.0000
   7.250   0.7369   0.03904   0.03416   0.0085   0.0119   1.0000
   7.500   0.7449   0.04560   0.04125   0.0090   0.0117   1.0000
   7.750   0.7505   0.05222   0.04826   0.0086   0.0114   1.0000
   8.000   0.7533   0.05884   0.05518   0.0072   0.0111   1.0000
   8.250   0.7529   0.06566   0.06220   0.0046   0.0109   1.0000
   8.500   0.7486   0.07305   0.06971   0.0001   0.0110   1.0000
   8.750   0.7399   0.08072   0.07743  -0.0070   0.0113   1.0000
   9.000   0.7326   0.08751   0.08418  -0.0131   0.0114   1.0000
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