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NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M2 AIRFOIL (m2-il)
Reynolds number: 200,000
Max Cl/Cd: 31.98 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m2-il-200000.txt
Download as CSV file: xf-m2-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M2 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5685   0.09901   0.09561   0.0025   1.0000   0.0643
  -9.750  -0.5802   0.09328   0.08991  -0.0009   1.0000   0.0668
  -9.500  -0.7164   0.09353   0.09008   0.0039   1.0000   0.0592
  -9.250  -0.7127   0.09008   0.08663   0.0032   1.0000   0.0607
  -9.000  -0.7188   0.08422   0.08081  -0.0020   1.0000   0.0618
  -8.750  -0.7317   0.07823   0.07477  -0.0070   1.0000   0.0626
  -8.500  -0.7396   0.07299   0.06943  -0.0099   1.0000   0.0644
  -8.250  -0.7630   0.06890   0.06464  -0.0132   1.0000   0.0688
  -8.000  -0.7615   0.06112   0.05683  -0.0139   1.0000   0.0702
  -7.750  -0.7464   0.05756   0.05336  -0.0135   1.0000   0.0718
  -7.500  -0.7322   0.05499   0.05076  -0.0131   1.0000   0.0752
  -7.250  -0.7433   0.03798   0.03193  -0.0101   1.0000   0.0459
  -7.000  -0.7283   0.03288   0.02648  -0.0091   1.0000   0.0448
  -6.750  -0.7098   0.02918   0.02227  -0.0079   1.0000   0.0450
  -6.500  -0.6902   0.02573   0.01833  -0.0069   1.0000   0.0469
  -6.250  -0.6672   0.02373   0.01618  -0.0063   1.0000   0.0488
  -6.000  -0.6428   0.02205   0.01427  -0.0057   1.0000   0.0505
  -5.750  -0.6179   0.02080   0.01284  -0.0051   1.0000   0.0533
  -5.500  -0.5925   0.01967   0.01150  -0.0044   1.0000   0.0559
  -5.250  -0.5676   0.01823   0.00990  -0.0037   1.0000   0.0576
  -5.000  -0.5435   0.01706   0.00876  -0.0031   1.0000   0.0609
  -4.750  -0.5187   0.01642   0.00808  -0.0025   1.0000   0.0648
  -4.500  -0.4942   0.01566   0.00724  -0.0017   1.0000   0.0680
  -4.250  -0.4710   0.01473   0.00635  -0.0009   1.0000   0.0720
  -4.000  -0.4469   0.01425   0.00589  -0.0001   1.0000   0.0788
  -3.750  -0.4234   0.01366   0.00534   0.0007   1.0000   0.0892
  -3.500  -0.4005   0.01292   0.00469   0.0016   1.0000   0.1132
  -3.250  -0.3772   0.01241   0.00423   0.0023   1.0000   0.1393
  -3.000  -0.3537   0.01194   0.00387   0.0030   1.0000   0.1582
  -2.750  -0.3300   0.01154   0.00355   0.0036   1.0000   0.1807
  -2.500  -0.3071   0.01074   0.00316   0.0041   1.0000   0.2512
  -2.250  -0.2896   0.00910   0.00303   0.0054   1.0000   0.5904
  -2.000  -0.2698   0.00857   0.00310   0.0076   1.0000   0.7276
  -1.750  -0.2504   0.00839   0.00323   0.0102   1.0000   0.8162
  -1.500  -0.2307   0.00849   0.00350   0.0131   1.0000   0.8841
  -1.250  -0.2063   0.00868   0.00370   0.0147   1.0000   0.9355
  -1.000  -0.1657   0.00883   0.00380   0.0126   1.0000   0.9661
  -0.750  -0.1119   0.00896   0.00386   0.0073   1.0000   0.9835
  -0.500  -0.0519   0.00901   0.00388   0.0006   1.0000   0.9931
  -0.250  -0.0057   0.00899   0.00383  -0.0036   1.0000   1.0000
   0.000   0.0000   0.00896   0.00380   0.0000   1.0000   1.0000
   0.250   0.0057   0.00899   0.00383   0.0036   1.0000   1.0000
   0.500   0.0520   0.00901   0.00387  -0.0006   0.9931   1.0000
   0.750   0.1119   0.00896   0.00386  -0.0073   0.9835   1.0000
   1.000   0.1657   0.00883   0.00380  -0.0126   0.9661   1.0000
   1.250   0.2062   0.00867   0.00370  -0.0147   0.9355   1.0000
   1.500   0.2308   0.00849   0.00350  -0.0132   0.8849   1.0000
   1.750   0.2503   0.00838   0.00323  -0.0102   0.8160   1.0000
   2.000   0.2698   0.00857   0.00310  -0.0076   0.7277   1.0000
   2.250   0.2896   0.00910   0.00303  -0.0054   0.5906   1.0000
   2.500   0.3071   0.01074   0.00316  -0.0041   0.2518   1.0000
   2.750   0.3299   0.01154   0.00355  -0.0036   0.1807   1.0000
   3.000   0.3537   0.01194   0.00387  -0.0030   0.1582   1.0000
   3.250   0.3771   0.01241   0.00423  -0.0023   0.1393   1.0000
   3.500   0.4005   0.01292   0.00469  -0.0016   0.1132   1.0000
   3.750   0.4234   0.01365   0.00534  -0.0007   0.0893   1.0000
   4.000   0.4468   0.01425   0.00589   0.0001   0.0788   1.0000
   4.250   0.4710   0.01473   0.00635   0.0009   0.0720   1.0000
   4.500   0.4942   0.01566   0.00724   0.0017   0.0680   1.0000
   4.750   0.5187   0.01642   0.00808   0.0025   0.0648   1.0000
   5.000   0.5435   0.01706   0.00876   0.0031   0.0609   1.0000
   5.250   0.5676   0.01823   0.00990   0.0037   0.0576   1.0000
   5.500   0.5925   0.01967   0.01150   0.0044   0.0559   1.0000
   5.750   0.6179   0.02081   0.01285   0.0051   0.0533   1.0000
   6.000   0.6429   0.02205   0.01427   0.0057   0.0505   1.0000
   6.250   0.6672   0.02373   0.01617   0.0063   0.0488   1.0000
   6.500   0.6902   0.02573   0.01833   0.0069   0.0469   1.0000
   6.750   0.7098   0.02918   0.02227   0.0079   0.0450   1.0000
   7.000   0.7283   0.03288   0.02648   0.0091   0.0448   1.0000
   7.250   0.7434   0.03796   0.03191   0.0101   0.0459   1.0000
   7.500   0.7322   0.05500   0.05078   0.0131   0.0753   1.0000
   7.750   0.7465   0.05757   0.05337   0.0135   0.0718   1.0000
   8.000   0.7617   0.06112   0.05683   0.0138   0.0702   1.0000
   8.250   0.7631   0.06890   0.06464   0.0132   0.0688   1.0000
   8.500   0.7399   0.07301   0.06945   0.0098   0.0644   1.0000
   8.750   0.7321   0.07825   0.07479   0.0069   0.0626   1.0000
   9.000   0.7194   0.08423   0.08082   0.0019   0.0617   1.0000
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