Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA M2 (nacam2-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA M2 (nacam2-il)
Reynolds number: 200,000
Max Cl/Cd: 36.07 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-nacam2-il-200000-n5.txt
Download as CSV file: xf-nacam2-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M2                                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.7540   0.09741   0.09393   0.0096   1.0000   0.0157
 -10.500  -0.7733   0.08735   0.08393   0.0034   1.0000   0.0155
 -10.250  -0.9120   0.05151   0.04741  -0.0180   1.0000   0.0135
 -10.000  -0.9213   0.04745   0.04308  -0.0170   1.0000   0.0137
  -9.750  -0.9162   0.04519   0.04065  -0.0161   1.0000   0.0141
  -9.500  -0.9085   0.04293   0.03818  -0.0152   1.0000   0.0145
  -9.250  -0.9018   0.03992   0.03485  -0.0141   1.0000   0.0150
  -9.000  -0.8930   0.03676   0.03131  -0.0128   1.0000   0.0156
  -8.750  -0.8817   0.03364   0.02775  -0.0114   1.0000   0.0165
  -8.500  -0.8672   0.03080   0.02445  -0.0101   1.0000   0.0177
  -8.250  -0.8487   0.02872   0.02191  -0.0090   1.0000   0.0193
  -8.000  -0.8303   0.02664   0.01956  -0.0081   1.0000   0.0205
  -7.750  -0.8088   0.02552   0.01833  -0.0075   1.0000   0.0217
  -7.500  -0.7865   0.02440   0.01702  -0.0069   1.0000   0.0233
  -7.250  -0.7634   0.02341   0.01581  -0.0062   1.0000   0.0258
  -7.000  -0.7396   0.02246   0.01460  -0.0056   1.0000   0.0278
  -6.750  -0.7169   0.02095   0.01290  -0.0049   1.0000   0.0295
  -6.500  -0.6939   0.01993   0.01185  -0.0044   1.0000   0.0320
  -6.250  -0.6698   0.01914   0.01098  -0.0038   1.0000   0.0343
  -6.000  -0.6456   0.01834   0.01004  -0.0032   1.0000   0.0365
  -5.750  -0.6210   0.01773   0.00930  -0.0027   1.0000   0.0385
  -5.500  -0.5974   0.01685   0.00835  -0.0020   1.0000   0.0398
  -5.250  -0.5746   0.01593   0.00742  -0.0012   1.0000   0.0421
  -5.000  -0.5507   0.01534   0.00679  -0.0006   1.0000   0.0446
  -4.750  -0.5267   0.01479   0.00618   0.0000   1.0000   0.0465
  -4.500  -0.5026   0.01429   0.00562   0.0007   1.0000   0.0483
  -4.250  -0.4783   0.01386   0.00512   0.0014   1.0000   0.0501
  -4.000  -0.4542   0.01342   0.00463   0.0021   1.0000   0.0528
  -3.750  -0.4301   0.01302   0.00424   0.0027   1.0000   0.0573
  -3.500  -0.4060   0.01267   0.00389   0.0034   1.0000   0.0650
  -3.250  -0.3825   0.01219   0.00357   0.0041   1.0000   0.0921
  -3.000  -0.3596   0.01157   0.00325   0.0047   1.0000   0.1578
  -2.750  -0.3369   0.01096   0.00301   0.0053   1.0000   0.2408
  -2.500  -0.3141   0.01040   0.00283   0.0060   1.0000   0.3295
  -2.000  -0.2693   0.00934   0.00262   0.0076   1.0000   0.5325
  -1.750  -0.2472   0.00891   0.00263   0.0088   1.0000   0.6317
  -1.500  -0.2250   0.00862   0.00269   0.0102   1.0000   0.7153
  -1.250  -0.1986   0.00842   0.00281   0.0109   0.9974   0.7970
  -1.000  -0.1663   0.00838   0.00299   0.0108   0.9922   0.8749
  -0.750  -0.1281   0.00845   0.00311   0.0091   0.9873   0.9183
  -0.500  -0.0872   0.00849   0.00313   0.0065   0.9815   0.9366
  -0.250  -0.0441   0.00850   0.00314   0.0033   0.9742   0.9504
   0.000   0.0000   0.00851   0.00313   0.0000   0.9631   0.9631
   0.250   0.0441   0.00850   0.00314  -0.0033   0.9504   0.9742
   0.500   0.0871   0.00848   0.00313  -0.0065   0.9367   0.9814
   0.750   0.1281   0.00845   0.00311  -0.0091   0.9183   0.9873
   1.000   0.1663   0.00838   0.00299  -0.0108   0.8747   0.9922
   1.250   0.1986   0.00842   0.00281  -0.0109   0.7975   0.9974
   1.500   0.2249   0.00862   0.00269  -0.0101   0.7155   1.0000
   1.750   0.2471   0.00891   0.00263  -0.0088   0.6315   1.0000
   2.000   0.2692   0.00934   0.00262  -0.0076   0.5325   1.0000
   2.250   0.2918   0.00979   0.00269  -0.0067   0.4392   1.0000
   2.500   0.3140   0.01040   0.00283  -0.0060   0.3296   1.0000
   2.750   0.3368   0.01095   0.00301  -0.0053   0.2410   1.0000
   3.000   0.3596   0.01156   0.00325  -0.0047   0.1582   1.0000
   3.250   0.3824   0.01219   0.00357  -0.0041   0.0923   1.0000
   3.500   0.4059   0.01267   0.00389  -0.0034   0.0651   1.0000
   3.750   0.4301   0.01302   0.00424  -0.0027   0.0573   1.0000
   4.000   0.4542   0.01342   0.00463  -0.0021   0.0528   1.0000
   4.250   0.4783   0.01386   0.00512  -0.0014   0.0501   1.0000
   4.500   0.5025   0.01429   0.00562  -0.0007   0.0483   1.0000
   4.750   0.5266   0.01478   0.00618   0.0000   0.0465   1.0000
   5.000   0.5507   0.01534   0.00679   0.0006   0.0446   1.0000
   5.250   0.5746   0.01593   0.00742   0.0012   0.0421   1.0000
   5.500   0.5974   0.01685   0.00835   0.0020   0.0398   1.0000
   5.750   0.6210   0.01773   0.00930   0.0027   0.0385   1.0000
   6.000   0.6456   0.01835   0.01004   0.0032   0.0365   1.0000
   6.250   0.6698   0.01914   0.01098   0.0038   0.0343   1.0000
   6.500   0.6938   0.01993   0.01186   0.0044   0.0320   1.0000
   6.750   0.7169   0.02094   0.01290   0.0049   0.0295   1.0000
   7.000   0.7396   0.02246   0.01460   0.0056   0.0278   1.0000
   7.250   0.7634   0.02341   0.01581   0.0062   0.0258   1.0000
   7.500   0.7865   0.02441   0.01703   0.0069   0.0233   1.0000
   7.750   0.8088   0.02552   0.01833   0.0075   0.0217   1.0000
   8.000   0.8303   0.02664   0.01956   0.0081   0.0205   1.0000
   8.250   0.8487   0.02873   0.02191   0.0090   0.0193   1.0000
   8.500   0.8673   0.03080   0.02446   0.0101   0.0177   1.0000
   8.750   0.8818   0.03364   0.02776   0.0114   0.0165   1.0000
   9.000   0.8931   0.03677   0.03132   0.0128   0.0156   1.0000
   9.250   0.9018   0.03992   0.03485   0.0140   0.0150   1.0000
   9.500   0.9088   0.04290   0.03814   0.0152   0.0145   1.0000
   9.750   0.9170   0.04508   0.04053   0.0161   0.0141   1.0000
  10.000   0.9222   0.04733   0.04295   0.0170   0.0137   1.0000
  10.250   0.9123   0.05150   0.04739   0.0180   0.0135   1.0000
  10.500   0.7737   0.08749   0.08407  -0.0036   0.0155   1.0000
  10.750   0.7545   0.09753   0.09405  -0.0098   0.0157   1.0000
<< Back to NACA M2 (nacam2-il)

Polar data table (+)

Polar graphs


<< Back to NACA M2 (nacam2-il)