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

NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 500,000
Max Cl/Cd: 98.09 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m13-il-500000.txt
Download as CSV file: xf-m13-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M13 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4796   0.10991   0.10775   0.0071   1.0000   0.0151
  -8.500  -0.4756   0.10656   0.10442   0.0044   1.0000   0.0151
  -8.250  -0.4713   0.10309   0.10097   0.0017   1.0000   0.0152
  -8.000  -0.4665   0.09950   0.09741  -0.0009   1.0000   0.0152
  -7.750  -0.4605   0.09575   0.09367  -0.0041   1.0000   0.0152
  -7.500  -0.4500   0.09151   0.08944  -0.0084   1.0000   0.0153
  -7.250  -0.4460   0.08554   0.08349  -0.0112   1.0000   0.0155
  -7.000  -0.4374   0.08225   0.08021  -0.0106   1.0000   0.0159
  -6.750  -0.4253   0.07897   0.07691  -0.0126   1.0000   0.0163
  -6.500  -0.4111   0.07548   0.07342  -0.0157   1.0000   0.0168
  -6.250  -0.3947   0.07193   0.06985  -0.0192   1.0000   0.0177
  -6.000  -0.3744   0.06811   0.06600  -0.0235   1.0000   0.0195
  -5.750  -0.3411   0.06388   0.06166  -0.0309   1.0000   0.0207
  -5.500  -0.3171   0.05973   0.05744  -0.0344   1.0000   0.0209
  -5.250  -0.2937   0.05546   0.05308  -0.0373   1.0000   0.0209
  -5.000  -0.2621   0.05041   0.04789  -0.0417   0.9876   0.0210
  -4.750  -0.2354   0.04278   0.04008  -0.0469   0.9646   0.0219
  -4.500  -0.2103   0.04071   0.03792  -0.0485   0.9379   0.0229
  -4.250  -0.1881   0.03876   0.03581  -0.0487   0.9118   0.0248
  -4.000  -0.1552   0.03647   0.03321  -0.0490   0.8901   0.0288
  -3.750  -0.1285   0.03334   0.02977  -0.0491   0.8711   0.0291
  -3.500  -0.1074   0.02696   0.02296  -0.0499   0.8547   0.0306
  -3.250  -0.0843   0.02596   0.02185  -0.0499   0.8367   0.0322
  -3.000  -0.0584   0.02444   0.02011  -0.0498   0.8202   0.0356
  -2.750  -0.0306   0.02154   0.01676  -0.0496   0.8056   0.0416
  -2.500  -0.0050   0.02083   0.01595  -0.0496   0.7905   0.0446
  -2.250   0.0225   0.01940   0.01420  -0.0495   0.7769   0.0527
  -1.500   0.1086   0.01258   0.00606  -0.0481   0.7424   0.0406
  -1.250   0.1366   0.01205   0.00540  -0.0479   0.7311   0.0399
  -1.000   0.1644   0.01132   0.00454  -0.0478   0.7206   0.0399
  -0.750   0.1920   0.01081   0.00393  -0.0476   0.7105   0.0392
  -0.500   0.2195   0.01015   0.00318  -0.0474   0.7003   0.0389
  -0.250   0.2471   0.00973   0.00269  -0.0473   0.6904   0.0391
   0.000   0.2748   0.00940   0.00228  -0.0472   0.6809   0.0398
   0.250   0.3026   0.00916   0.00198  -0.0471   0.6711   0.0409
   0.500   0.3306   0.00902   0.00180  -0.0471   0.6612   0.0420
   0.750   0.3586   0.00895   0.00169  -0.0471   0.6518   0.0433
   1.000   0.3866   0.00891   0.00162  -0.0471   0.6421   0.0464
   1.250   0.4147   0.00887   0.00161  -0.0472   0.6317   0.0500
   1.500   0.4428   0.00886   0.00162  -0.0472   0.6217   0.0633
   1.750   0.4707   0.00891   0.00164  -0.0472   0.6116   0.0730
   2.000   0.4987   0.00891   0.00165  -0.0472   0.6008   0.0828
   2.250   0.5267   0.00891   0.00167  -0.0473   0.5894   0.0931
   2.500   0.5546   0.00891   0.00172  -0.0473   0.5778   0.1102
   2.750   0.5814   0.00697   0.00186  -0.0476   0.5668   1.0000
   3.000   0.6088   0.00709   0.00191  -0.0475   0.5556   1.0000
   3.250   0.6364   0.00719   0.00199  -0.0475   0.5443   1.0000
   3.500   0.6639   0.00730   0.00208  -0.0475   0.5332   1.0000
   3.750   0.6914   0.00743   0.00220  -0.0474   0.5224   1.0000
   4.000   0.7188   0.00756   0.00231  -0.0474   0.5081   1.0000
   4.250   0.7460   0.00771   0.00241  -0.0473   0.4897   1.0000
   4.500   0.7727   0.00791   0.00252  -0.0472   0.4617   1.0000
   4.750   0.7994   0.00815   0.00268  -0.0471   0.4323   1.0000
   5.000   0.8258   0.00845   0.00287  -0.0470   0.3973   1.0000
   5.250   0.8509   0.00901   0.00313  -0.0468   0.3253   1.0000
   5.500   0.8648   0.01196   0.00455  -0.0461   0.0410   1.0000
   5.750   0.8893   0.01277   0.00537  -0.0456   0.0256   1.0000
   6.000   0.9147   0.01329   0.00598  -0.0452   0.0230   1.0000
   6.250   0.9392   0.01393   0.00667  -0.0449   0.0200   1.0000
   6.500   0.9597   0.01531   0.00819  -0.0439   0.0177   1.0000
   6.750   0.9825   0.01616   0.00913  -0.0432   0.0170   1.0000
   7.000   1.0046   0.01707   0.01013  -0.0424   0.0163   1.0000
   7.250   1.0254   0.01821   0.01138  -0.0414   0.0155   1.0000
   7.500   1.0465   0.01928   0.01252  -0.0405   0.0146   1.0000
   7.750   1.0681   0.02018   0.01345  -0.0398   0.0135   1.0000
   8.000   1.0881   0.02147   0.01480  -0.0388   0.0130   1.0000
   8.250   1.1073   0.02315   0.01655  -0.0376   0.0125   1.0000
   8.500   1.1265   0.02552   0.01905  -0.0365   0.0121   1.0000
   8.750   1.1463   0.02833   0.02204  -0.0354   0.0120   1.0000
   9.000   1.1665   0.03047   0.02439  -0.0342   0.0122   1.0000
  18.250   0.7392   0.19971   0.19766  -0.0625   0.0123   1.0000
  18.500   0.7423   0.20248   0.20044  -0.0630   0.0118   1.0000
<< Back to NACA M13 AIRFOIL (m13-il)

Polar data table (+)

Polar graphs


<< Back to NACA M13 AIRFOIL (m13-il)