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

MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 50,000
Max Cl/Cd: 42.52 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ma409-il-50000-n5.txt
Download as CSV file: xf-ma409-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MA409 (original)                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4199   0.09650   0.08988  -0.0277   1.0000   0.0421
  -8.750  -0.4240   0.09330   0.08679  -0.0281   1.0000   0.0411
  -8.250  -0.4323   0.08463   0.07836  -0.0344   1.0000   0.0384
  -8.000  -0.4294   0.08093   0.07471  -0.0357   1.0000   0.0382
  -7.750  -0.4255   0.07697   0.07080  -0.0376   1.0000   0.0380
  -7.500  -0.4201   0.07266   0.06653  -0.0401   1.0000   0.0378
  -7.250  -0.4125   0.06791   0.06178  -0.0433   1.0000   0.0377
  -7.000  -0.4018   0.06254   0.05637  -0.0473   1.0000   0.0376
  -6.750  -0.3872   0.05645   0.05014  -0.0521   1.0000   0.0375
  -6.500  -0.3677   0.04951   0.04290  -0.0578   1.0000   0.0379
  -6.250  -0.3471   0.04507   0.03820  -0.0610   1.0000   0.0401
  -6.000  -0.3239   0.04137   0.03415  -0.0635   1.0000   0.0439
  -5.750  -0.2916   0.03507   0.02694  -0.0682   1.0000   0.0473
  -5.500  -0.2633   0.03161   0.02289  -0.0698   1.0000   0.0519
  -5.250  -0.2343   0.02909   0.01978  -0.0710   1.0000   0.0648
  -5.000  -0.2055   0.02692   0.01717  -0.0717   1.0000   0.0845
  -4.750  -0.1819   0.02685   0.01681  -0.0718   1.0000   0.1187
  -4.500  -0.1560   0.02581   0.01545  -0.0718   1.0000   0.1340
  -4.250  -0.1300   0.02479   0.01416  -0.0717   1.0000   0.1418
  -4.000  -0.1038   0.02398   0.01299  -0.0716   1.0000   0.1507
  -3.750  -0.0775   0.02324   0.01199  -0.0714   1.0000   0.1571
  -3.500  -0.0508   0.02260   0.01109  -0.0712   1.0000   0.1639
  -3.250  -0.0238   0.02213   0.01041  -0.0712   1.0000   0.1751
  -3.000   0.0029   0.02171   0.00989  -0.0713   1.0000   0.1909
  -2.750   0.0300   0.02130   0.00948  -0.0715   1.0000   0.2101
  -2.500   0.0574   0.02090   0.00915  -0.0717   1.0000   0.2427
  -2.250   0.0849   0.02053   0.00895  -0.0721   1.0000   0.3078
  -2.000   0.1119   0.02014   0.00890  -0.0724   1.0000   0.4051
  -1.750   0.1382   0.01964   0.00884  -0.0724   1.0000   0.5232
  -1.500   0.1522   0.01856   0.00872  -0.0692   0.9956   1.0000
  -1.250   0.1889   0.01888   0.00868  -0.0714   0.9893   1.0000
  -1.000   0.2248   0.01922   0.00874  -0.0735   0.9832   1.0000
  -0.750   0.2613   0.01957   0.00888  -0.0757   0.9776   1.0000
  -0.500   0.2942   0.01988   0.00904  -0.0771   0.9703   1.0000
  -0.250   0.3300   0.02019   0.00924  -0.0791   0.9623   1.0000
   0.000   0.3693   0.02048   0.00947  -0.0817   0.9536   1.0000
   0.250   0.4050   0.02069   0.00964  -0.0834   0.9416   1.0000
   0.500   0.4418   0.02084   0.00980  -0.0852   0.9288   1.0000
   0.750   0.4792   0.02092   0.00994  -0.0870   0.9147   1.0000
   1.000   0.5133   0.02102   0.01012  -0.0882   0.9015   1.0000
   1.250   0.5462   0.02114   0.01034  -0.0891   0.8888   1.0000
   1.500   0.5793   0.02123   0.01060  -0.0900   0.8756   1.0000
   1.750   0.6126   0.02127   0.01081  -0.0908   0.8614   1.0000
   2.000   0.6454   0.02126   0.01099  -0.0914   0.8464   1.0000
   2.250   0.6791   0.02116   0.01117  -0.0919   0.8303   1.0000
   2.500   0.7110   0.02102   0.01127  -0.0918   0.8097   1.0000
   2.750   0.7494   0.02062   0.01112  -0.0924   0.7879   1.0000
   3.000   0.7873   0.02028   0.01105  -0.0926   0.7606   1.0000
   3.250   0.8215   0.02020   0.01118  -0.0922   0.7288   1.0000
   3.500   0.8422   0.02056   0.01184  -0.0904   0.6887   1.0000
   3.750   0.8597   0.02022   0.01131  -0.0862   0.5928   1.0000
   4.000   0.8842   0.02096   0.01160  -0.0840   0.5169   1.0000
   4.250   0.8963   0.02193   0.01234  -0.0807   0.4056   1.0000
   4.500   0.9022   0.02397   0.01308  -0.0772   0.2152   1.0000
   4.750   0.9147   0.02696   0.01484  -0.0756   0.0658   1.0000
   5.000   0.9341   0.02887   0.01661  -0.0743   0.0454   1.0000
   5.250   0.9531   0.03053   0.01844  -0.0728   0.0395   1.0000
   5.500   0.9707   0.03231   0.02044  -0.0711   0.0367   1.0000
   5.750   0.9890   0.03407   0.02247  -0.0694   0.0338   1.0000
   6.000   1.0058   0.03609   0.02460  -0.0678   0.0308   1.0000
   6.250   1.0249   0.03828   0.02714  -0.0662   0.0291   1.0000
   6.500   1.0480   0.04071   0.02990  -0.0649   0.0283   1.0000
   6.750   1.0717   0.04357   0.03312  -0.0637   0.0277   1.0000
   7.000   1.0927   0.04680   0.03679  -0.0624   0.0274   1.0000
   7.250   1.1092   0.05030   0.04077  -0.0607   0.0271   1.0000
   7.500   1.1205   0.05395   0.04492  -0.0588   0.0267   1.0000
   7.750   1.1267   0.05771   0.04917  -0.0566   0.0262   1.0000
   8.000   1.1280   0.06158   0.05349  -0.0542   0.0258   1.0000
   8.250   1.1246   0.06552   0.05783  -0.0519   0.0256   1.0000
   8.500   1.1167   0.06951   0.06218  -0.0496   0.0255   1.0000
   8.750   1.1034   0.07336   0.06632  -0.0472   0.0256   1.0000
   9.000   1.0870   0.07747   0.07068  -0.0455   0.0257   1.0000
   9.250   1.0688   0.08206   0.07548  -0.0450   0.0259   1.0000
   9.500   1.0497   0.08733   0.08093  -0.0461   0.0262   1.0000
   9.750   1.0307   0.09342   0.08718  -0.0488   0.0266   1.0000
<< Back to MA409 (original) (modified line 7) (ma409-il)

Polar data table (+)

Polar graphs


<< Back to MA409 (original) (modified line 7) (ma409-il)