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=200,000 Ncrit=9


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 200,000
Max Cl/Cd: 81.47 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409-il-200000.txt
Download as CSV file: xf-ma409-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4291   0.09960   0.09627  -0.0290   1.0000   0.0471
  -9.250  -0.4354   0.09726   0.09402  -0.0299   1.0000   0.0473
  -9.000  -0.4427   0.09482   0.09166  -0.0303   1.0000   0.0474
  -8.750  -0.4435   0.09161   0.08849  -0.0321   1.0000   0.0475
  -8.500  -0.4432   0.08808   0.08501  -0.0342   1.0000   0.0476
  -8.250  -0.4486   0.08261   0.07960  -0.0338   1.0000   0.0482
  -8.000  -0.4476   0.07939   0.07642  -0.0316   1.0000   0.0488
  -7.750  -0.4455   0.07656   0.07362  -0.0305   1.0000   0.0494
  -7.500  -0.4436   0.07368   0.07076  -0.0303   1.0000   0.0500
  -7.250  -0.4423   0.07072   0.06783  -0.0307   1.0000   0.0507
  -7.000  -0.4117   0.06461   0.06165  -0.0398   0.9961   0.0517
  -6.750  -0.3531   0.03710   0.03330  -0.0715   0.9892   0.0295
  -6.500  -0.3137   0.02704   0.02217  -0.0798   0.9862   0.0302
  -6.250  -0.2771   0.02138   0.01551  -0.0833   0.9847   0.0303
  -6.000  -0.2413   0.01873   0.01231  -0.0853   0.9833   0.0320
  -5.750  -0.2070   0.01701   0.01020  -0.0865   0.9811   0.0352
  -5.500  -0.1744   0.01610   0.00916  -0.0877   0.9772   0.0424
  -5.250  -0.1391   0.01550   0.00840  -0.0890   0.9740   0.0616
  -5.000  -0.1032   0.01553   0.00830  -0.0909   0.9712   0.0843
  -4.750  -0.0685   0.01555   0.00822  -0.0926   0.9679   0.0987
  -4.500  -0.0374   0.01508   0.00767  -0.0934   0.9627   0.1066
  -4.250  -0.0024   0.01488   0.00730  -0.0949   0.9593   0.1149
  -4.000   0.0336   0.01413   0.00657  -0.0967   0.9572   0.1196
  -3.750   0.0673   0.01368   0.00608  -0.0978   0.9535   0.1249
  -3.500   0.0985   0.01322   0.00559  -0.0985   0.9486   0.1320
  -3.250   0.1331   0.01287   0.00522  -0.0999   0.9456   0.1420
  -3.000   0.1681   0.01238   0.00481  -0.1014   0.9434   0.1565
  -2.750   0.1991   0.01185   0.00450  -0.1022   0.9395   0.1891
  -2.500   0.2285   0.01126   0.00442  -0.1029   0.9351   0.3186
  -2.250   0.2602   0.01088   0.00437  -0.1038   0.9317   0.4201
  -2.000   0.2927   0.01043   0.00424  -0.1047   0.9289   0.5095
  -1.750   0.3143   0.00972   0.00429  -0.1032   0.9232   0.7152
  -1.500   0.3462   0.00923   0.00412  -0.1033   0.9194   1.0000
  -1.250   0.3791   0.00917   0.00395  -0.1041   0.9155   1.0000
  -1.000   0.4062   0.00918   0.00390  -0.1038   0.9080   1.0000
  -0.750   0.4373   0.00912   0.00377  -0.1043   0.9027   1.0000
  -0.500   0.4650   0.00913   0.00376  -0.1041   0.8957   1.0000
  -0.250   0.4945   0.00911   0.00370  -0.1042   0.8897   1.0000
   0.000   0.5222   0.00909   0.00367  -0.1040   0.8814   1.0000
   0.250   0.5515   0.00900   0.00356  -0.1039   0.8719   1.0000
   0.500   0.5797   0.00892   0.00349  -0.1035   0.8602   1.0000
   0.750   0.6071   0.00889   0.00347  -0.1031   0.8478   1.0000
   1.000   0.6351   0.00886   0.00343  -0.1028   0.8342   1.0000
   1.250   0.6630   0.00883   0.00338  -0.1023   0.8162   1.0000
   1.500   0.6912   0.00885   0.00334  -0.1018   0.7952   1.0000
   1.750   0.7188   0.00901   0.00346  -0.1013   0.7751   1.0000
   2.000   0.7451   0.00933   0.00365  -0.1005   0.7488   1.0000
   2.250   0.7672   0.00967   0.00388  -0.0988   0.7053   1.0000
   2.500   0.7862   0.00965   0.00367  -0.0965   0.5986   1.0000
   2.750   0.8041   0.01056   0.00405  -0.0943   0.5481   1.0000
   3.000   0.8261   0.01098   0.00438  -0.0931   0.4777   1.0000
   3.250   0.8468   0.01155   0.00466  -0.0917   0.4012   1.0000
   3.500   0.8635   0.01264   0.00515  -0.0898   0.2908   1.0000
   3.750   0.8752   0.01477   0.00604  -0.0876   0.0842   1.0000
   4.000   0.8949   0.01606   0.00699  -0.0862   0.0359   1.0000
   4.250   0.9180   0.01684   0.00794  -0.0851   0.0318   1.0000
   4.500   0.9397   0.01781   0.00904  -0.0838   0.0294   1.0000
   4.750   0.9591   0.01910   0.01043  -0.0822   0.0280   1.0000
   5.000   0.9760   0.02099   0.01238  -0.0802   0.0268   1.0000
   5.250   0.9987   0.02197   0.01348  -0.0791   0.0252   1.0000
   5.500   1.0206   0.02366   0.01526  -0.0778   0.0248   1.0000
   5.750   1.0442   0.02571   0.01744  -0.0766   0.0248   1.0000
   6.000   1.0687   0.02826   0.02020  -0.0756   0.0253   1.0000
   6.250   1.0927   0.03304   0.02530  -0.0747   0.0269   1.0000
   6.500   1.1187   0.03309   0.02566  -0.0730   0.0293   1.0000
  13.500   0.9800   0.19058   0.18733  -0.1008   0.0369   1.0000
  13.750   0.9794   0.19630   0.19301  -0.1072   0.0356   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)