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NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 500,000
Max Cl/Cd: 88.64 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m16-il-500000-n5.txt
Download as CSV file: xf-m16-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M16 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5102   0.09806   0.09578   0.0225   0.8176   0.0053
  -7.500  -0.5073   0.09448   0.09213   0.0201   0.7983   0.0053
  -7.250  -0.4995   0.09043   0.08801   0.0164   0.7818   0.0054
  -7.000  -0.4886   0.08650   0.08401   0.0128   0.7669   0.0057
  -6.750  -0.4758   0.08217   0.07961   0.0087   0.7536   0.0062
  -6.250  -0.4443   0.07246   0.06972  -0.0001   0.7304   0.0065
  -6.000  -0.4253   0.06734   0.06451  -0.0044   0.7203   0.0065
  -5.750  -0.4041   0.06177   0.05880  -0.0087   0.7111   0.0067
  -5.500  -0.3803   0.05536   0.05223  -0.0131   0.7025   0.0070
  -5.250  -0.3558   0.05097   0.04766  -0.0160   0.6941   0.0075
  -5.000  -0.3307   0.04821   0.04476  -0.0178   0.6852   0.0081
  -4.750  -0.3039   0.04409   0.04045  -0.0200   0.6777   0.0093
  -4.500  -0.2742   0.03599   0.03194  -0.0225   0.6717   0.0106
  -4.250  -0.2472   0.03342   0.02918  -0.0234   0.6642   0.0115
  -4.000  -0.2164   0.02110   0.01587  -0.0240   0.6601   0.0171
  -3.750  -0.1893   0.02089   0.01554  -0.0242   0.6528   0.0180
  -3.500  -0.1621   0.02184   0.01650  -0.0245   0.6451   0.0186
  -3.250  -0.1352   0.02242   0.01706  -0.0248   0.6383   0.0198
  -3.000  -0.1076   0.02126   0.01570  -0.0249   0.6317   0.0217
  -2.750  -0.0793   0.01930   0.01341  -0.0249   0.6259   0.0240
  -2.500  -0.0509   0.01752   0.01131  -0.0248   0.6199   0.0256
  -2.250  -0.0226   0.01634   0.00984  -0.0248   0.6140   0.0266
  -2.000   0.0056   0.01549   0.00877  -0.0248   0.6081   0.0273
  -1.750   0.0339   0.01472   0.00780  -0.0248   0.6021   0.0277
  -1.500   0.0622   0.01435   0.00729  -0.0249   0.5964   0.0281
  -1.250   0.0901   0.01265   0.00532  -0.0250   0.5906   0.0304
  -1.000   0.1182   0.01222   0.00481  -0.0251   0.5848   0.0317
  -0.750   0.1463   0.01185   0.00438  -0.0252   0.5788   0.0326
  -0.500   0.1745   0.01151   0.00397  -0.0252   0.5726   0.0334
  -0.250   0.2026   0.01119   0.00358  -0.0253   0.5667   0.0340
   0.000   0.2307   0.01088   0.00324  -0.0254   0.5604   0.0344
   0.250   0.2587   0.01068   0.00299  -0.0255   0.5543   0.0353
   0.500   0.2869   0.01044   0.00273  -0.0256   0.5476   0.0360
   1.000   0.3431   0.01000   0.00224  -0.0257   0.5345   0.0354
   1.250   0.3713   0.00985   0.00205  -0.0258   0.5277   0.0351
   1.500   0.3996   0.00972   0.00191  -0.0260   0.5207   0.0349
   1.750   0.4278   0.00964   0.00181  -0.0261   0.5134   0.0348
   2.000   0.4562   0.00957   0.00175  -0.0263   0.5061   0.0348
   2.500   0.5129   0.00953   0.00168  -0.0266   0.4908   0.0351
   2.750   0.5412   0.00954   0.00168  -0.0268   0.4834   0.0358
   3.000   0.5695   0.00955   0.00172  -0.0270   0.4752   0.0370
   3.250   0.5978   0.00959   0.00176  -0.0272   0.4677   0.0399
   3.500   0.6248   0.00759   0.00193  -0.0274   0.4599   1.0000
   3.750   0.6525   0.00770   0.00204  -0.0274   0.4522   1.0000
   4.000   0.6801   0.00784   0.00215  -0.0275   0.4376   1.0000
   4.250   0.7076   0.00806   0.00227  -0.0277   0.4139   1.0000
   4.500   0.7350   0.00830   0.00243  -0.0278   0.3857   1.0000
   4.750   0.7623   0.00860   0.00264  -0.0280   0.3533   1.0000
   5.000   0.7882   0.00948   0.00305  -0.0284   0.2561   1.0000
   5.250   0.8098   0.01202   0.00444  -0.0293   0.0210   1.0000
   5.500   0.8365   0.01249   0.00495  -0.0293   0.0137   1.0000
   5.750   0.8632   0.01288   0.00544  -0.0293   0.0122   1.0000
   6.000   0.8896   0.01335   0.00599  -0.0294   0.0106   1.0000
   6.250   0.9154   0.01389   0.00657  -0.0294   0.0088   1.0000
   6.500   0.9398   0.01486   0.00767  -0.0293   0.0077   1.0000
   6.750   0.9646   0.01556   0.00848  -0.0292   0.0072   1.0000
   7.000   0.9883   0.01643   0.00946  -0.0290   0.0067   1.0000
   7.250   1.0114   0.01733   0.01045  -0.0287   0.0061   1.0000
   7.500   1.0350   0.01797   0.01112  -0.0286   0.0055   1.0000
   7.750   1.0554   0.01914   0.01234  -0.0281   0.0050   1.0000
   8.000   1.0751   0.02040   0.01371  -0.0274   0.0047   1.0000
   8.250   1.0943   0.02168   0.01516  -0.0266   0.0045   1.0000
   8.500   1.1122   0.02318   0.01679  -0.0256   0.0043   1.0000
   8.750   1.1295   0.02486   0.01862  -0.0244   0.0040   1.0000
   9.000   1.1465   0.02673   0.02065  -0.0232   0.0038   1.0000
   9.250   1.1631   0.02884   0.02295  -0.0219   0.0037   1.0000
   9.500   1.1786   0.03134   0.02568  -0.0206   0.0036   1.0000
   9.750   1.1919   0.03421   0.02882  -0.0192   0.0036   1.0000
  10.000   1.2037   0.03645   0.03127  -0.0180   0.0034   1.0000
  10.250   1.2135   0.03830   0.03330  -0.0168   0.0033   1.0000
  10.500   1.2201   0.04002   0.03518  -0.0156   0.0032   1.0000
  10.750   1.2191   0.04213   0.03746  -0.0136   0.0031   1.0000
  11.000   1.2092   0.04542   0.04098  -0.0118   0.0030   1.0000
  11.250   1.1997   0.04951   0.04536  -0.0107   0.0029   1.0000
  11.500   1.1890   0.05394   0.05006  -0.0103   0.0029   1.0000
  11.750   1.1762   0.05862   0.05497  -0.0108   0.0028   1.0000
  12.000   1.1618   0.06373   0.06028  -0.0119   0.0028   1.0000
  12.250   1.1462   0.06924   0.06599  -0.0138   0.0028   1.0000
  12.500   1.1301   0.07503   0.07195  -0.0162   0.0028   1.0000
  12.750   1.1133   0.08127   0.07834  -0.0192   0.0028   1.0000
  13.000   1.0957   0.08802   0.08524  -0.0227   0.0028   1.0000
  13.250   1.0782   0.09517   0.09253  -0.0267   0.0029   1.0000
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