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

MA409 (smoothed) (ma409sm-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MA409 (smoothed) (ma409sm-il)
Reynolds number: 50,000
Max Cl/Cd: 43.47 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409sm-il-50000.txt
Download as CSV file: xf-ma409sm-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MA409 (smoothed)                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.4301   0.10104   0.09431  -0.0184   1.0000   0.2006
  -7.250  -0.4169   0.09655   0.08983  -0.0167   1.0000   0.2106
  -7.000  -0.4351   0.09654   0.09002  -0.0197   1.0000   0.2154
  -6.750  -0.4212   0.09189   0.08538  -0.0164   1.0000   0.2273
  -6.250  -0.4225   0.08677   0.08044  -0.0182   1.0000   0.2443
  -6.000  -0.4148   0.08341   0.07711  -0.0164   1.0000   0.2556
  -5.750  -0.4084   0.07994   0.07370  -0.0150   1.0000   0.2646
  -5.500  -0.4049   0.07702   0.07085  -0.0151   1.0000   0.2764
  -5.250  -0.4000   0.07408   0.06791  -0.0151   1.0000   0.2899
  -5.000  -0.3937   0.07108   0.06495  -0.0147   1.0000   0.3047
  -4.750  -0.3864   0.06812   0.06203  -0.0142   1.0000   0.3218
  -4.500  -0.3792   0.06520   0.05915  -0.0144   1.0000   0.3440
  -4.250  -0.2771   0.04648   0.03902  -0.0584   1.0000   0.1472
  -4.000  -0.2461   0.04167   0.03379  -0.0619   1.0000   0.1437
  -3.750  -0.2121   0.03679   0.02840  -0.0654   1.0000   0.1397
  -3.500  -0.1748   0.03231   0.02316  -0.0687   1.0000   0.1398
  -3.250  -0.1435   0.02979   0.02017  -0.0701   1.0000   0.1558
  -3.000  -0.1101   0.02715   0.01702  -0.0712   1.0000   0.1705
  -2.750  -0.0797   0.02548   0.01503  -0.0718   1.0000   0.2017
  -2.500  -0.0513   0.02393   0.01343  -0.0718   1.0000   0.2372
  -2.250  -0.0231   0.02279   0.01229  -0.0718   1.0000   0.2895
  -2.000   0.0042   0.02189   0.01147  -0.0715   1.0000   0.3504
  -1.750   0.0326   0.02102   0.01067  -0.0713   1.0000   0.4192
  -1.500   0.0612   0.02001   0.00988  -0.0712   1.0000   0.5036
  -1.250   0.0875   0.01856   0.00910  -0.0702   1.0000   0.6403
  -1.000   0.0996   0.01728   0.00830  -0.0666   1.0000   1.0000
  -0.750   0.1281   0.01753   0.00800  -0.0673   1.0000   1.0000
  -0.500   0.1538   0.01781   0.00792  -0.0674   1.0000   1.0000
  -0.250   0.1786   0.01812   0.00797  -0.0674   1.0000   1.0000
   0.000   0.2029   0.01847   0.00807  -0.0674   1.0000   1.0000
   0.250   0.2268   0.01885   0.00828  -0.0673   1.0000   1.0000
   0.500   0.2502   0.01927   0.00857  -0.0673   1.0000   1.0000
   0.750   0.2731   0.01974   0.00893  -0.0672   1.0000   1.0000
   1.000   0.2956   0.02025   0.00937  -0.0672   1.0000   1.0000
   1.250   0.3176   0.02082   0.00989  -0.0671   1.0000   1.0000
   1.500   0.3391   0.02146   0.01050  -0.0671   1.0000   1.0000
   1.750   0.3600   0.02216   0.01120  -0.0672   1.0000   1.0000
   2.000   0.3802   0.02295   0.01200  -0.0672   1.0000   1.0000
   2.250   0.3997   0.02384   0.01294  -0.0674   1.0000   1.0000
   2.500   0.4184   0.02483   0.01398  -0.0676   1.0000   1.0000
   2.750   0.4362   0.02594   0.01516  -0.0679   1.0000   1.0000
   3.000   0.4767   0.02730   0.01665  -0.0726   0.9878   1.0000
   3.250   0.5383   0.02855   0.01814  -0.0806   0.9628   1.0000
   3.500   0.5958   0.02944   0.01928  -0.0871   0.9365   1.0000
   3.750   0.6502   0.02998   0.02011  -0.0923   0.9091   1.0000
   4.000   0.7026   0.03018   0.02069  -0.0964   0.8808   1.0000
   4.250   0.7551   0.02995   0.02084  -0.0997   0.8519   1.0000
   4.500   0.8103   0.02913   0.02048  -0.1022   0.8230   1.0000
   4.750   0.8634   0.02778   0.01967  -0.1030   0.7933   1.0000
   5.000   0.9030   0.02664   0.01896  -0.1015   0.7568   1.0000
   5.250   0.9436   0.02346   0.01603  -0.0956   0.6943   1.0000
   5.500   0.9604   0.02212   0.01452  -0.0883   0.6057   1.0000
   5.750   0.9684   0.02228   0.01423  -0.0814   0.4827   1.0000
   6.000   0.9550   0.02587   0.01556  -0.0734   0.2241   1.0000
   6.250   0.9712   0.02930   0.01825  -0.0709   0.1507   1.0000
   6.500   0.9960   0.03177   0.02051  -0.0697   0.1226   1.0000
   6.750   1.0284   0.03456   0.02329  -0.0692   0.1090   1.0000
   7.000   1.0612   0.03790   0.02676  -0.0689   0.1023   1.0000
   7.250   1.0867   0.04077   0.02990  -0.0679   0.0948   1.0000
   7.500   1.1094   0.04446   0.03390  -0.0668   0.0904   1.0000
   7.750   1.1293   0.04839   0.03840  -0.0651   0.0902   1.0000
   8.000   1.1458   0.05276   0.04326  -0.0633   0.0908   1.0000
   8.250   1.1615   0.05782   0.04865  -0.0618   0.0919   1.0000
   8.500   1.1557   0.06192   0.05395  -0.0577   0.0965   1.0000
   8.750   1.1524   0.06736   0.05993  -0.0554   0.1003   1.0000
   9.000   1.1585   0.07307   0.06590  -0.0541   0.1039   1.0000
   9.250   1.1313   0.07833   0.07177  -0.0517   0.1091   1.0000
   9.500   1.1090   0.08416   0.07788  -0.0508   0.1132   1.0000
   9.750   1.1029   0.09001   0.08385  -0.0503   0.1195   1.0000
  10.000   1.0615   0.09569   0.08967  -0.0515   0.1209   1.0000
  10.250   1.0209   0.10441   0.09842  -0.0580   0.1236   1.0000
  10.500   0.8776   0.10133   0.09551  -0.0442   0.1140   1.0000
  10.750   0.9458   0.14315   0.13696  -0.0945   0.2804   1.0000
<< Back to MA409 (smoothed) (ma409sm-il)

Polar data table (+)

Polar graphs


<< Back to MA409 (smoothed) (ma409sm-il)