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NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA M2 AIRFOIL (m2-il)
Reynolds number: 200,000
Max Cl/Cd: 36.81 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m2-il-200000-n5.txt
Download as CSV file: xf-m2-il-200000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.7792   0.08418   0.08082   0.0020   1.0000   0.0158
 -10.000  -0.8379   0.06515   0.06167  -0.0136   1.0000   0.0149
  -9.750  -0.8940   0.05106   0.04707  -0.0150   1.0000   0.0146
  -9.500  -0.9019   0.04549   0.04111  -0.0142   1.0000   0.0149
  -9.250  -0.9011   0.04091   0.03610  -0.0131   1.0000   0.0153
  -9.000  -0.8947   0.03692   0.03165  -0.0118   1.0000   0.0160
  -8.750  -0.8841   0.03331   0.02755  -0.0106   1.0000   0.0169
  -8.500  -0.8688   0.03045   0.02419  -0.0095   1.0000   0.0183
  -8.250  -0.8492   0.02852   0.02181  -0.0086   1.0000   0.0197
  -8.000  -0.8309   0.02630   0.01936  -0.0078   1.0000   0.0210
  -7.750  -0.8090   0.02514   0.01808  -0.0073   1.0000   0.0223
  -7.500  -0.7862   0.02402   0.01674  -0.0068   1.0000   0.0242
  -7.250  -0.7627   0.02295   0.01543  -0.0062   1.0000   0.0267
  -7.000  -0.7384   0.02200   0.01422  -0.0057   1.0000   0.0284
  -6.750  -0.7161   0.02029   0.01236  -0.0051   1.0000   0.0307
  -6.500  -0.6923   0.01944   0.01146  -0.0047   1.0000   0.0332
  -6.250  -0.6680   0.01859   0.01047  -0.0042   1.0000   0.0354
  -6.000  -0.6433   0.01786   0.00961  -0.0037   1.0000   0.0377
  -5.750  -0.6186   0.01721   0.00882  -0.0032   1.0000   0.0393
  -5.500  -0.5953   0.01621   0.00778  -0.0025   1.0000   0.0411
  -5.250  -0.5713   0.01552   0.00709  -0.0020   1.0000   0.0438
  -5.000  -0.5469   0.01497   0.00647  -0.0015   1.0000   0.0463
  -4.750  -0.5225   0.01445   0.00588  -0.0009   1.0000   0.0484
  -4.500  -0.4981   0.01399   0.00535  -0.0003   1.0000   0.0506
  -4.250  -0.4740   0.01352   0.00490   0.0003   1.0000   0.0545
  -4.000  -0.4497   0.01314   0.00453   0.0009   1.0000   0.0609
  -3.750  -0.4255   0.01281   0.00425   0.0016   1.0000   0.0733
  -3.500  -0.4011   0.01257   0.00403   0.0022   1.0000   0.0895
  -3.250  -0.3768   0.01235   0.00381   0.0028   1.0000   0.1037
  -3.000  -0.3526   0.01209   0.00359   0.0033   1.0000   0.1151
  -2.750  -0.3284   0.01186   0.00338   0.0039   1.0000   0.1246
  -2.500  -0.3040   0.01167   0.00320   0.0045   1.0000   0.1344
  -2.250  -0.2799   0.01141   0.00302   0.0050   1.0000   0.1520
  -2.000  -0.2557   0.01111   0.00286   0.0055   1.0000   0.1825
  -1.750  -0.2315   0.00997   0.00265   0.0053   0.9991   0.4068
  -1.500  -0.1971   0.00935   0.00257   0.0035   0.9943   0.5324
  -1.250  -0.1626   0.00876   0.00259   0.0021   0.9876   0.6650
  -1.000  -0.1280   0.00844   0.00264   0.0011   0.9790   0.7590
  -0.750  -0.0937   0.00831   0.00272   0.0003   0.9676   0.8165
  -0.500  -0.0608   0.00830   0.00283   0.0000   0.9540   0.8626
  -0.250  -0.0293   0.00833   0.00290  -0.0002   0.9394   0.8969
   0.000   0.0000   0.00834   0.00291   0.0000   0.9209   0.9209
   0.250   0.0293   0.00833   0.00290   0.0002   0.8969   0.9394
   0.500   0.0609   0.00830   0.00283   0.0000   0.8629   0.9540
   0.750   0.0937   0.00831   0.00272  -0.0003   0.8165   0.9676
   1.000   0.1280   0.00844   0.00264  -0.0011   0.7591   0.9790
   1.250   0.1628   0.00876   0.00259  -0.0021   0.6650   0.9877
   1.500   0.1972   0.00935   0.00257  -0.0035   0.5329   0.9944
   1.750   0.2315   0.00997   0.00265  -0.0053   0.4069   0.9991
   2.000   0.2557   0.01111   0.00286  -0.0055   0.1823   1.0000
   2.250   0.2798   0.01141   0.00302  -0.0050   0.1521   1.0000
   2.500   0.3040   0.01167   0.00320  -0.0045   0.1344   1.0000
   2.750   0.3283   0.01186   0.00338  -0.0039   0.1246   1.0000
   3.000   0.3526   0.01209   0.00359  -0.0033   0.1151   1.0000
   3.250   0.3767   0.01235   0.00381  -0.0027   0.1038   1.0000
   3.500   0.4010   0.01257   0.00403  -0.0021   0.0896   1.0000
   3.750   0.4254   0.01281   0.00425  -0.0015   0.0732   1.0000
   4.000   0.4496   0.01314   0.00453  -0.0009   0.0609   1.0000
   4.250   0.4739   0.01352   0.00490  -0.0003   0.0545   1.0000
   4.500   0.4981   0.01399   0.00535   0.0003   0.0506   1.0000
   4.750   0.5225   0.01445   0.00588   0.0009   0.0484   1.0000
   5.000   0.5469   0.01497   0.00647   0.0015   0.0463   1.0000
   5.250   0.5713   0.01552   0.00709   0.0020   0.0438   1.0000
   5.500   0.5953   0.01621   0.00778   0.0025   0.0411   1.0000
   5.750   0.6186   0.01721   0.00882   0.0032   0.0393   1.0000
   6.000   0.6434   0.01786   0.00961   0.0037   0.0377   1.0000
   6.250   0.6680   0.01859   0.01047   0.0042   0.0354   1.0000
   6.500   0.6923   0.01944   0.01146   0.0047   0.0332   1.0000
   6.750   0.7162   0.02029   0.01236   0.0051   0.0307   1.0000
   7.000   0.7385   0.02200   0.01422   0.0057   0.0284   1.0000
   7.250   0.7628   0.02295   0.01543   0.0062   0.0267   1.0000
   7.500   0.7863   0.02402   0.01675   0.0068   0.0242   1.0000
   7.750   0.8091   0.02514   0.01808   0.0073   0.0223   1.0000
   8.000   0.8311   0.02630   0.01937   0.0078   0.0210   1.0000
   8.250   0.8493   0.02854   0.02183   0.0085   0.0197   1.0000
   8.500   0.8689   0.03046   0.02421   0.0094   0.0183   1.0000
   8.750   0.8843   0.03332   0.02756   0.0105   0.0169   1.0000
   9.000   0.8948   0.03694   0.03167   0.0118   0.0160   1.0000
   9.250   0.9013   0.04093   0.03612   0.0130   0.0153   1.0000
   9.500   0.9020   0.04552   0.04114   0.0141   0.0149   1.0000
   9.750   0.8942   0.05111   0.04712   0.0149   0.0146   1.0000
  10.000   0.8377   0.06531   0.06183   0.0135   0.0149   1.0000
  10.250   0.7796   0.08434   0.08098  -0.0023   0.0158   1.0000
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