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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 200,000
Max Cl/Cd: 72.82 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ma409-il-200000-n5.txt
Download as CSV file: xf-ma409-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.3269   0.10638   0.10314  -0.0282   1.0000   0.0151
 -11.000  -0.3283   0.10228   0.09907  -0.0288   1.0000   0.0139
 -10.750  -0.4266   0.11235   0.10886  -0.0217   1.0000   0.0162
 -10.500  -0.4248   0.10828   0.10483  -0.0230   1.0000   0.0145
 -10.250  -0.4271   0.10306   0.09965  -0.0251   1.0000   0.0126
 -10.000  -0.4266   0.09841   0.09506  -0.0268   1.0000   0.0115
  -9.750  -0.4229   0.09534   0.09203  -0.0276   1.0000   0.0111
  -9.250  -0.4221   0.08931   0.08605  -0.0280   1.0000   0.0106
  -9.000  -0.4275   0.08614   0.08294  -0.0278   1.0000   0.0104
  -8.750  -0.4237   0.08175   0.07860  -0.0310   0.9951   0.0103
  -8.500  -0.4132   0.07634   0.07321  -0.0370   0.9877   0.0101
  -8.250  -0.3988   0.07012   0.06700  -0.0450   0.9798   0.0100
  -8.000  -0.3813   0.06292   0.05979  -0.0549   0.9711   0.0099
  -7.750  -0.3584   0.05356   0.05036  -0.0678   0.9626   0.0097
  -7.500  -0.3249   0.03363   0.02977  -0.0900   0.9527   0.0092
  -7.250  -0.2983   0.02617   0.02143  -0.0956   0.9454   0.0092
  -7.000  -0.2685   0.02300   0.01766  -0.0978   0.9412   0.0101
  -6.750  -0.2399   0.02105   0.01528  -0.0988   0.9366   0.0110
  -6.500  -0.2119   0.01803   0.01167  -0.0997   0.9314   0.0116
  -6.250  -0.1817   0.01618   0.00943  -0.1006   0.9274   0.0125
  -6.000  -0.1526   0.01498   0.00791  -0.1009   0.9223   0.0139
  -5.750  -0.1233   0.01408   0.00677  -0.1013   0.9168   0.0169
  -5.500  -0.0921   0.01332   0.00578  -0.1018   0.9122   0.0203
  -5.250  -0.0640   0.01264   0.00524  -0.1019   0.9052   0.0447
  -5.000  -0.0335   0.01257   0.00507  -0.1023   0.8989   0.0651
  -4.750  -0.0048   0.01250   0.00484  -0.1024   0.8906   0.0745
  -4.500   0.0249   0.01224   0.00452  -0.1028   0.8823   0.0828
  -4.250   0.0542   0.01207   0.00421  -0.1030   0.8719   0.0898
  -4.000   0.0832   0.01168   0.00376  -0.1033   0.8613   0.0933
  -3.750   0.1124   0.01141   0.00342  -0.1035   0.8516   0.0982
  -3.250   0.1710   0.01094   0.00284  -0.1039   0.8310   0.1132
  -3.000   0.1999   0.01072   0.00258  -0.1041   0.8210   0.1265
  -2.750   0.2286   0.01051   0.00237  -0.1042   0.8131   0.1496
  -2.500   0.2567   0.01025   0.00222  -0.1043   0.8065   0.1899
  -2.250   0.2847   0.01000   0.00214  -0.1044   0.7999   0.2620
  -2.000   0.3125   0.00985   0.00208  -0.1044   0.7934   0.3136
  -1.500   0.3679   0.00961   0.00203  -0.1043   0.7834   0.4025
  -1.250   0.3952   0.00948   0.00202  -0.1042   0.7781   0.4600
  -1.000   0.4211   0.00912   0.00206  -0.1038   0.7730   0.5797
  -0.750   0.4477   0.00829   0.00206  -0.1028   0.7676   1.0000
  -0.500   0.4755   0.00841   0.00208  -0.1027   0.7608   1.0000
  -0.250   0.5030   0.00853   0.00214  -0.1026   0.7541   1.0000
   0.000   0.5305   0.00870   0.00223  -0.1024   0.7459   1.0000
   0.250   0.5573   0.00886   0.00234  -0.1021   0.7350   1.0000
   0.500   0.5841   0.00901   0.00246  -0.1018   0.7245   1.0000
   0.750   0.6108   0.00916   0.00260  -0.1015   0.7138   1.0000
   1.000   0.6373   0.00926   0.00275  -0.1012   0.7014   1.0000
   1.250   0.6634   0.00924   0.00288  -0.1009   0.6797   1.0000
   1.500   0.6845   0.00940   0.00270  -0.0991   0.5877   1.0000
   1.750   0.7068   0.00983   0.00293  -0.0980   0.5652   1.0000
   2.000   0.7316   0.01020   0.00325  -0.0974   0.5476   1.0000
   2.250   0.7563   0.01050   0.00355  -0.0967   0.5028   1.0000
   2.500   0.7774   0.01105   0.00377  -0.0955   0.4274   1.0000
   2.750   0.7958   0.01195   0.00415  -0.0940   0.3252   1.0000
   3.000   0.8137   0.01311   0.00465  -0.0926   0.1961   1.0000
   3.250   0.8290   0.01486   0.00552  -0.0909   0.0383   1.0000
   3.500   0.8514   0.01570   0.00628  -0.0899   0.0194   1.0000
   3.750   0.8750   0.01636   0.00709  -0.0890   0.0157   1.0000
   4.000   0.8977   0.01719   0.00812  -0.0880   0.0139   1.0000
   4.250   0.9178   0.01841   0.00956  -0.0865   0.0126   1.0000
   4.500   0.9401   0.01920   0.01043  -0.0856   0.0113   1.0000
   4.750   0.9607   0.02026   0.01160  -0.0843   0.0104   1.0000
   5.000   0.9805   0.02154   0.01297  -0.0828   0.0099   1.0000
   5.250   1.0006   0.02296   0.01449  -0.0814   0.0095   1.0000
   5.500   1.0215   0.02455   0.01619  -0.0801   0.0092   1.0000
   5.750   1.0433   0.02634   0.01811  -0.0788   0.0090   1.0000
   6.000   1.0649   0.02810   0.02009  -0.0778   0.0085   1.0000
   6.250   1.0847   0.03003   0.02212  -0.0768   0.0076   1.0000
   6.500   1.1021   0.03390   0.02624  -0.0755   0.0070   1.0000
   6.750   1.1193   0.03599   0.02871  -0.0737   0.0068   1.0000
   7.000   1.1339   0.03862   0.03171  -0.0717   0.0068   1.0000
   7.250   1.1453   0.04156   0.03504  -0.0695   0.0067   1.0000
   7.500   1.1532   0.04480   0.03866  -0.0669   0.0068   1.0000
   7.750   1.1572   0.04830   0.04253  -0.0642   0.0068   1.0000
   8.000   1.1569   0.05213   0.04669  -0.0614   0.0069   1.0000
   9.500   1.1035   0.07120   0.06720  -0.0463   0.0075   1.0000
   9.750   1.0855   0.07602   0.07226  -0.0469   0.0076   1.0000
  10.000   1.0645   0.08234   0.07881  -0.0497   0.0077   1.0000
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