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NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 1,000,000
Max Cl/Cd: 115.33 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m13-il-1000000.txt
Download as CSV file: xf-m13-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.500   0.3289   0.00778   0.00142  -0.0464   0.6098   0.0324
   0.750   0.3571   0.00770   0.00130  -0.0465   0.6004   0.0331
   1.000   0.3854   0.00763   0.00122  -0.0466   0.5908   0.0335
   1.250   0.4136   0.00761   0.00116  -0.0466   0.5811   0.0339
   1.750   0.4701   0.00759   0.00109  -0.0468   0.5585   0.0358
   2.000   0.4983   0.00760   0.00108  -0.0469   0.5466   0.0364
   2.250   0.5264   0.00762   0.00108  -0.0470   0.5353   0.0380
   2.500   0.5545   0.00763   0.00113  -0.0471   0.5243   0.0623
   2.750   0.5826   0.00769   0.00121  -0.0472   0.5132   0.0740
   3.000   0.6107   0.00772   0.00126  -0.0474   0.5019   0.0832
   3.250   0.6388   0.00777   0.00132  -0.0475   0.4914   0.0883
   3.500   0.6667   0.00784   0.00140  -0.0476   0.4772   0.0973
   3.750   0.6909   0.00607   0.00162  -0.0474   0.4602   1.0000
   4.000   0.7183   0.00624   0.00172  -0.0474   0.4393   1.0000
   4.250   0.7455   0.00647   0.00183  -0.0475   0.4118   1.0000
   4.500   0.7727   0.00670   0.00196  -0.0475   0.3837   1.0000
   4.750   0.7992   0.00708   0.00215  -0.0475   0.3347   1.0000
   5.000   0.8233   0.00802   0.00257  -0.0474   0.2263   1.0000
   5.250   0.8430   0.00996   0.00357  -0.0470   0.0353   1.0000
   5.500   0.8689   0.01049   0.00402  -0.0468   0.0192   1.0000
   5.750   0.8953   0.01085   0.00444  -0.0466   0.0166   1.0000
   6.000   0.9215   0.01122   0.00484  -0.0465   0.0148   1.0000
   6.250   0.9472   0.01170   0.00536  -0.0463   0.0132   1.0000
   6.500   0.9699   0.01279   0.00661  -0.0456   0.0116   1.0000
   6.750   0.9956   0.01317   0.00703  -0.0454   0.0112   1.0000
   7.000   1.0207   0.01363   0.00754  -0.0451   0.0105   1.0000
   7.250   1.0453   0.01413   0.00808  -0.0448   0.0096   1.0000
   7.500   1.0690   0.01478   0.00877  -0.0443   0.0090   1.0000
   7.750   1.0917   0.01553   0.00957  -0.0438   0.0085   1.0000
   8.000   1.1118   0.01667   0.01079  -0.0429   0.0081   1.0000
   8.250   1.1231   0.01941   0.01369  -0.0406   0.0076   1.0000
   8.500   1.1468   0.01987   0.01423  -0.0402   0.0074   1.0000
   8.750   1.1686   0.02070   0.01516  -0.0394   0.0071   1.0000
   9.000   1.1884   0.02196   0.01653  -0.0384   0.0067   1.0000
   9.250   1.2068   0.02364   0.01836  -0.0372   0.0065   1.0000
   9.500   1.2250   0.02528   0.02014  -0.0360   0.0062   1.0000
   9.750   1.2431   0.02669   0.02169  -0.0349   0.0060   1.0000
  10.000   1.2600   0.02815   0.02330  -0.0338   0.0057   1.0000
  10.250   1.2722   0.03084   0.02624  -0.0321   0.0057   1.0000
  10.500   1.2761   0.03543   0.03124  -0.0294   0.0058   1.0000
  10.750   1.2657   0.04209   0.03844  -0.0255   0.0062   1.0000
  11.000   1.2541   0.04686   0.04355  -0.0222   0.0064   1.0000
  11.250   1.2358   0.05060   0.04752  -0.0183   0.0066   1.0000
  11.500   1.2174   0.05441   0.05154  -0.0161   0.0067   1.0000
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