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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 100,000
Max Cl/Cd: 59.56 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ma409-il-100000-n5.txt
Download as CSV file: xf-ma409-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4344   0.08634   0.08185  -0.0290   1.0000   0.0203
  -8.500  -0.4380   0.08320   0.07879  -0.0291   1.0000   0.0201
  -8.250  -0.4399   0.07975   0.07541  -0.0302   1.0000   0.0200
  -8.000  -0.4406   0.07603   0.07175  -0.0318   1.0000   0.0198
  -7.750  -0.4395   0.07194   0.06771  -0.0341   1.0000   0.0197
  -7.500  -0.4360   0.06735   0.06316  -0.0372   1.0000   0.0195
  -7.250  -0.4291   0.06205   0.05786  -0.0414   1.0000   0.0193
  -7.000  -0.3968   0.05226   0.04789  -0.0542   0.9953   0.0191
  -6.750  -0.3546   0.03769   0.03260  -0.0704   0.9900   0.0187
  -6.500  -0.3149   0.02886   0.02258  -0.0786   0.9871   0.0189
  -6.250  -0.2806   0.02481   0.01768  -0.0815   0.9839   0.0201
  -6.000  -0.2473   0.02216   0.01454  -0.0835   0.9809   0.0230
  -5.750  -0.2134   0.02044   0.01246  -0.0851   0.9784   0.0259
  -5.500  -0.1806   0.01906   0.01076  -0.0862   0.9754   0.0311
  -5.250  -0.1488   0.01840   0.01005  -0.0871   0.9714   0.0499
  -5.000  -0.1154   0.01838   0.00973  -0.0883   0.9679   0.0779
  -4.750  -0.0840   0.01814   0.00929  -0.0893   0.9641   0.0922
  -4.500  -0.0543   0.01805   0.00904  -0.0900   0.9594   0.1069
  -4.250  -0.0223   0.01753   0.00839  -0.0910   0.9563   0.1116
  -4.000   0.0110   0.01712   0.00777  -0.0921   0.9538   0.1180
  -3.750   0.0393   0.01668   0.00732  -0.0924   0.9486   0.1236
  -3.500   0.0709   0.01630   0.00681  -0.0932   0.9443   0.1308
  -3.250   0.1048   0.01589   0.00638  -0.0946   0.9409   0.1409
  -3.000   0.1352   0.01553   0.00602  -0.0952   0.9356   0.1584
  -2.750   0.1665   0.01510   0.00572  -0.0960   0.9302   0.1927
  -2.500   0.2002   0.01467   0.00553  -0.0973   0.9261   0.2640
  -2.250   0.2278   0.01442   0.00542  -0.0973   0.9184   0.3303
  -2.000   0.2602   0.01410   0.00528  -0.0982   0.9134   0.4021
  -1.750   0.2881   0.01382   0.00521  -0.0983   0.9067   0.4809
  -1.500   0.3143   0.01308   0.00518  -0.0977   0.9013   0.6672
  -1.250   0.3417   0.01250   0.00499  -0.0968   0.8947   1.0000
  -1.000   0.3709   0.01254   0.00491  -0.0970   0.8874   1.0000
  -0.750   0.4003   0.01257   0.00486  -0.0973   0.8802   1.0000
  -0.500   0.4297   0.01260   0.00482  -0.0975   0.8726   1.0000
  -0.250   0.4578   0.01263   0.00480  -0.0974   0.8629   1.0000
   0.000   0.4876   0.01263   0.00477  -0.0976   0.8537   1.0000
   0.250   0.5170   0.01264   0.00479  -0.0978   0.8445   1.0000
   0.500   0.5453   0.01270   0.00486  -0.0977   0.8352   1.0000
   0.750   0.5755   0.01271   0.00490  -0.0980   0.8262   1.0000
   1.250   0.6377   0.01264   0.00491  -0.0985   0.8003   1.0000
   1.500   0.6699   0.01260   0.00487  -0.0988   0.7812   1.0000
   1.750   0.7018   0.01265   0.00494  -0.0990   0.7623   1.0000
   2.000   0.7320   0.01283   0.00513  -0.0990   0.7440   1.0000
   2.250   0.7587   0.01307   0.00545  -0.0985   0.7229   1.0000
   2.500   0.7822   0.01330   0.00584  -0.0975   0.6938   1.0000
   2.750   0.7999   0.01343   0.00566  -0.0945   0.5865   1.0000
   3.000   0.8198   0.01433   0.00611  -0.0926   0.5142   1.0000
   3.250   0.8394   0.01498   0.00652  -0.0910   0.4317   1.0000
   3.500   0.8545   0.01614   0.00697  -0.0887   0.3095   1.0000
   3.750   0.8697   0.01788   0.00775  -0.0870   0.1445   1.0000
   4.000   0.8849   0.01994   0.00897  -0.0852   0.0316   1.0000
   4.250   0.9066   0.02095   0.01014  -0.0840   0.0261   1.0000
   4.500   0.9264   0.02224   0.01160  -0.0825   0.0216   1.0000
   4.750   0.9467   0.02340   0.01298  -0.0811   0.0202   1.0000
   5.000   0.9658   0.02468   0.01447  -0.0794   0.0194   1.0000
   5.250   0.9840   0.02613   0.01610  -0.0777   0.0187   1.0000
   5.500   1.0027   0.02770   0.01782  -0.0761   0.0175   1.0000
   5.750   1.0223   0.02940   0.01965  -0.0747   0.0159   1.0000
   6.000   1.0429   0.03138   0.02187  -0.0734   0.0150   1.0000
   6.250   1.0646   0.03365   0.02435  -0.0722   0.0147   1.0000
   6.500   1.0859   0.03618   0.02716  -0.0710   0.0145   1.0000
   6.750   1.1056   0.03895   0.03028  -0.0695   0.0144   1.0000
   7.000   1.1225   0.04195   0.03368  -0.0677   0.0145   1.0000
   7.250   1.1359   0.04518   0.03737  -0.0656   0.0146   1.0000
   7.500   1.1445   0.04872   0.04143  -0.0630   0.0149   1.0000
   7.750   1.1489   0.05237   0.04555  -0.0603   0.0150   1.0000
   8.000   1.1492   0.05602   0.04963  -0.0575   0.0149   1.0000
   8.250   1.1453   0.05969   0.05365  -0.0547   0.0149   1.0000
   8.500   1.1370   0.06332   0.05760  -0.0519   0.0148   1.0000
   8.750   1.1225   0.06690   0.06146  -0.0487   0.0148   1.0000
   9.000   1.1054   0.07075   0.06557  -0.0464   0.0149   1.0000
   9.250   1.0859   0.07526   0.07032  -0.0454   0.0151   1.0000
   9.500   1.0645   0.08076   0.07604  -0.0463   0.0155   1.0000
   9.750   1.0426   0.08737   0.08285  -0.0493   0.0159   1.0000
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