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

NACA M27 AIRFOIL (m27-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M27 AIRFOIL (m27-il)
Reynolds number: 200,000
Max Cl/Cd: 62.75 at α=10.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m27-il-200000.txt
Download as CSV file: xf-m27-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M27 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2072   0.12108   0.11693  -0.0093   0.6740   0.0294
  -9.500  -0.2001   0.11838   0.11420  -0.0112   0.6701   0.0301
  -9.250  -0.1977   0.11721   0.11299  -0.0151   0.6669   0.0307
  -8.750  -0.1839   0.11045   0.10626  -0.0186   0.6588   0.0311
  -8.500  -0.1732   0.10648   0.10226  -0.0183   0.6547   0.0314
  -8.250  -0.1649   0.10342   0.09916  -0.0191   0.6513   0.0319
  -8.000  -0.1570   0.10059   0.09629  -0.0204   0.6480   0.0324
  -7.750  -0.1486   0.09772   0.09346  -0.0223   0.6434   0.0330
  -7.500  -0.1410   0.09489   0.09063  -0.0244   0.6394   0.0337
  -7.250  -0.1356   0.09229   0.08799  -0.0262   0.6360   0.0345
  -7.000  -0.1281   0.08983   0.08545  -0.0284   0.6332   0.0354
  -6.750  -0.1120   0.08939   0.08494  -0.0374   0.6296   0.0366
  -6.500  -0.0975   0.08613   0.08167  -0.0407   0.6257   0.0369
  -6.250  -0.0879   0.08151   0.07706  -0.0389   0.6220   0.0372
  -6.000  -0.0760   0.07821   0.07370  -0.0389   0.6188   0.0377
  -5.750  -0.0620   0.07545   0.07085  -0.0399   0.6159   0.0384
  -5.500  -0.0449   0.07276   0.06817  -0.0418   0.6115   0.0394
  -5.250  -0.0265   0.07015   0.06551  -0.0437   0.6075   0.0406
  -5.000  -0.0053   0.06770   0.06295  -0.0460   0.6042   0.0423
  -4.750   0.0325   0.06680   0.06171  -0.0521   0.6015   0.0440
  -4.500   0.0432   0.06280   0.05774  -0.0513   0.5984   0.0445
  -4.250   0.0595   0.05990   0.05487  -0.0516   0.5942   0.0452
  -4.000   0.0791   0.05750   0.05241  -0.0522   0.5904   0.0463
  -3.750   0.1010   0.05529   0.05008  -0.0530   0.5872   0.0478
  -3.500   0.1260   0.05333   0.04792  -0.0541   0.5845   0.0501
  -3.250   0.1693   0.05311   0.04735  -0.0573   0.5807   0.0525
  -3.000   0.1823   0.04938   0.04371  -0.0570   0.5770   0.0533
  -2.750   0.2006   0.04704   0.04134  -0.0568   0.5736   0.0546
  -2.500   0.2234   0.04522   0.03940  -0.0570   0.5706   0.0566
  -2.250   0.2496   0.04367   0.03764  -0.0573   0.5680   0.0594
  -2.000   0.2904   0.04371   0.03732  -0.0585   0.5639   0.0628
  -1.750   0.3074   0.04057   0.03424  -0.0585   0.5602   0.0638
  -1.500   0.3286   0.03868   0.03230  -0.0584   0.5570   0.0653
  -1.250   0.3530   0.03722   0.03071  -0.0583   0.5543   0.0679
  -1.000   0.3940   0.03870   0.03162  -0.0582   0.5518   0.0745
  -0.750   0.4112   0.03499   0.02807  -0.0585   0.5480   0.0761
  -0.500   0.4335   0.03346   0.02655  -0.0585   0.5441   0.0784
  -0.250   0.4597   0.03236   0.02532  -0.0584   0.5409   0.0823
   0.000   0.4928   0.03195   0.02450  -0.0581   0.5383   0.0890
   0.250   0.5159   0.03020   0.02269  -0.0580   0.5360   0.0911
   0.500   0.5411   0.02934   0.02185  -0.0581   0.5323   0.0949
   0.750   0.5713   0.02897   0.02123  -0.0580   0.5285   0.1040
   1.000   0.5960   0.02764   0.01990  -0.0580   0.5253   0.1072
   1.250   0.6255   0.02727   0.01924  -0.0577   0.5226   0.1191
   1.500   0.6510   0.02614   0.01806  -0.0578   0.5203   0.1255
   1.750   0.6772   0.02569   0.01759  -0.0578   0.5169   0.1390
   2.000   0.7036   0.02525   0.01712  -0.0579   0.5132   0.1534
   2.250   0.7304   0.02469   0.01650  -0.0579   0.5100   0.1702
   2.500   0.7721   0.02209   0.01309  -0.0569   0.5077   0.0703
   2.750   0.8015   0.02154   0.01236  -0.0569   0.5053   0.0708
   3.000   0.8287   0.02155   0.01227  -0.0569   0.5024   0.0744
   3.250   0.8535   0.02128   0.01214  -0.0568   0.4987   0.0774
   3.500   0.8788   0.02119   0.01210  -0.0566   0.4955   0.0819
   3.750   0.9044   0.02099   0.01188  -0.0562   0.4928   0.0896
   4.000   0.9296   0.02076   0.01168  -0.0557   0.4904   0.1129
   4.250   0.9549   0.02062   0.01153  -0.0552   0.4883   0.1700
   4.500   0.9764   0.02085   0.01213  -0.0548   0.4843   0.2782
   4.750   1.1285   0.02048   0.01289  -0.0816   0.4788   1.0000
   5.000   1.1529   0.02066   0.01300  -0.0812   0.4763   1.0000
   5.250   1.1781   0.02079   0.01304  -0.0809   0.4741   1.0000
   5.500   1.2019   0.02114   0.01335  -0.0805   0.4715   1.0000
   5.750   1.2217   0.02175   0.01409  -0.0799   0.4676   1.0000
   6.000   1.2434   0.02217   0.01455  -0.0794   0.4645   1.0000
   6.250   1.2665   0.02245   0.01483  -0.0789   0.4619   1.0000
   6.500   1.2907   0.02264   0.01498  -0.0785   0.4597   1.0000
   6.750   1.3162   0.02282   0.01508  -0.0783   0.4577   1.0000
   7.000   1.3341   0.02354   0.01594  -0.0774   0.4543   1.0000
   7.250   1.3512   0.02422   0.01674  -0.0764   0.4506   1.0000
   7.500   1.3716   0.02464   0.01721  -0.0756   0.4477   1.0000
   7.750   1.3942   0.02491   0.01750  -0.0751   0.4454   1.0000
   8.000   1.4204   0.02489   0.01743  -0.0749   0.4431   1.0000
   8.250   1.4394   0.02529   0.01789  -0.0739   0.4392   1.0000
   8.500   1.4534   0.02561   0.01834  -0.0723   0.4335   1.0000
   8.750   1.4826   0.02483   0.01744  -0.0722   0.4286   1.0000
   9.000   1.5046   0.02483   0.01746  -0.0714   0.4242   1.0000
   9.250   1.5146   0.02544   0.01824  -0.0693   0.4193   1.0000
   9.500   1.5372   0.02519   0.01798  -0.0686   0.4145   1.0000
   9.750   1.5600   0.02492   0.01766  -0.0678   0.4091   1.0000
  10.000   1.5659   0.02551   0.01845  -0.0651   0.4037   1.0000
  10.250   1.5864   0.02528   0.01819  -0.0640   0.3983   1.0000
  10.500   1.5959   0.02558   0.01856  -0.0616   0.3925   1.0000
  10.750   1.5970   0.02608   0.01917  -0.0581   0.3862   1.0000
  11.000   1.6021   0.02615   0.01920  -0.0549   0.3787   1.0000
  11.250   1.5850   0.02752   0.02071  -0.0497   0.3735   1.0000
  11.500   1.5927   0.02791   0.02107  -0.0476   0.3660   1.0000
  11.750   1.5790   0.03007   0.02342  -0.0446   0.3598   1.0000
  12.000   1.5796   0.03135   0.02466  -0.0428   0.3501   1.0000
  12.250   1.5652   0.03427   0.02776  -0.0409   0.3415   1.0000
  12.500   1.5578   0.03677   0.03031  -0.0394   0.3312   1.0000
  12.750   1.5466   0.03978   0.03334  -0.0381   0.3191   1.0000
  13.000   1.5318   0.04332   0.03689  -0.0369   0.3060   1.0000
  13.250   1.5150   0.04729   0.04092  -0.0358   0.2927   1.0000
  13.500   1.4966   0.05168   0.04531  -0.0350   0.2766   1.0000
  13.750   1.4767   0.05646   0.05007  -0.0344   0.2579   1.0000
  14.000   1.4554   0.06149   0.05503  -0.0338   0.2373   1.0000
  14.250   1.4320   0.06695   0.06040  -0.0334   0.2148   1.0000
  14.500   1.4058   0.07289   0.06623  -0.0332   0.1921   1.0000
  14.750   1.3785   0.07917   0.07240  -0.0331   0.1695   1.0000
  15.000   1.3497   0.08582   0.07893  -0.0333   0.1462   1.0000
  15.250   1.3220   0.09253   0.08552  -0.0336   0.1213   1.0000
  15.500   1.2950   0.09929   0.09215  -0.0342   0.0931   1.0000
<< Back to NACA M27 AIRFOIL (m27-il)

Polar data table (+)

Polar graphs


<< Back to NACA M27 AIRFOIL (m27-il)