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NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M9 AIRFOIL (m9-il)
Reynolds number: 50,000
Max Cl/Cd: 30.57 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m9-il-50000-n5.txt
Download as CSV file: xf-m9-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M9 AIRFOIL                                 
  
 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  -0.2251   0.11364   0.10797  -0.0107   1.0000   0.0897
  -8.750  -0.2172   0.11119   0.10562  -0.0131   1.0000   0.0922
  -8.500  -0.2124   0.10961   0.10417  -0.0163   1.0000   0.0941
  -8.250  -0.2022   0.10837   0.10295  -0.0236   0.9265   0.0952
  -8.000  -0.1938   0.10704   0.10156  -0.0296   0.8931   0.0956
  -7.750  -0.1603   0.09851   0.09294  -0.0257   0.8776   0.1006
  -7.500  -0.1507   0.09601   0.09037  -0.0268   0.8563   0.1036
  -7.250  -0.1450   0.09397   0.08830  -0.0277   0.8371   0.1059
  -7.000  -0.1424   0.09266   0.08695  -0.0299   0.8201   0.1087
  -6.750  -0.1366   0.09147   0.08570  -0.0335   0.8051   0.1100
  -6.500  -0.1268   0.08948   0.08366  -0.0372   0.7908   0.1104
  -6.250  -0.1141   0.08721   0.08132  -0.0409   0.7779   0.1107
  -6.000  -0.1021   0.08192   0.07603  -0.0345   0.7673   0.1151
  -5.500  -0.0712   0.07371   0.06755  -0.0448   0.7456   0.0806
  -5.250  -0.0561   0.07063   0.06438  -0.0455   0.7359   0.0790
  -5.000  -0.0356   0.06739   0.06104  -0.0496   0.7247   0.0801
  -4.750  -0.0128   0.06386   0.05733  -0.0539   0.7155   0.0811
  -4.500   0.0092   0.06051   0.05385  -0.0565   0.7053   0.0810
  -4.250   0.0305   0.05776   0.05101  -0.0581   0.6956   0.0833
  -4.000   0.0507   0.05600   0.04915  -0.0582   0.6864   0.0874
  -3.750   0.0771   0.05296   0.04593  -0.0611   0.6768   0.0887
  -3.500   0.1051   0.04982   0.04252  -0.0640   0.6689   0.0916
  -3.250   0.1381   0.04565   0.03796  -0.0686   0.6598   0.0948
  -3.000   0.1642   0.04299   0.03504  -0.0697   0.6523   0.0969
  -2.750   0.1877   0.04223   0.03422  -0.0695   0.6430   0.1018
  -2.500   0.2172   0.03946   0.03106  -0.0712   0.6360   0.1061
  -2.250   0.2500   0.03517   0.02609  -0.0740   0.6286   0.1114
  -2.000   0.2735   0.03527   0.02623  -0.0730   0.6206   0.1163
  -1.750   0.3022   0.03327   0.02381  -0.0737   0.6140   0.1221
  -1.500   0.3305   0.03184   0.02204  -0.0742   0.6063   0.1302
  -1.250   0.3570   0.03144   0.02147  -0.0735   0.6008   0.1376
  -1.000   0.3855   0.03067   0.02048  -0.0740   0.5927   0.1463
  -0.750   0.4110   0.03095   0.02071  -0.0732   0.5864   0.1556
  -0.500   0.4374   0.03094   0.02061  -0.0726   0.5803   0.1641
  -0.250   0.4634   0.03117   0.02078  -0.0723   0.5732   0.1759
   0.000   0.4898   0.03129   0.02077  -0.0715   0.5681   0.1896
   0.500   0.5436   0.03166   0.02075  -0.0710   0.5550   0.2288
   0.750   0.5711   0.03156   0.02041  -0.0704   0.5504   0.2516
   1.000   0.5981   0.03173   0.02042  -0.0707   0.5434   0.2744
   1.250   0.6258   0.03167   0.02017  -0.0706   0.5367   0.2968
   1.500   0.6545   0.03142   0.01966  -0.0702   0.5306   0.3162
   1.750   0.6808   0.03155   0.01969  -0.0702   0.5207   0.3306
   2.000   0.7095   0.03112   0.01907  -0.0695   0.5137   0.3498
   2.250   0.7353   0.03114   0.01910  -0.0694   0.5030   0.3698
   2.500   0.7632   0.03089   0.01870  -0.0689   0.4955   0.3840
   2.750   0.7886   0.03119   0.01900  -0.0687   0.4864   0.3888
   3.000   0.8170   0.03102   0.01865  -0.0680   0.4807   0.3899
   3.250   0.8397   0.03171   0.01941  -0.0679   0.4711   0.3917
   3.500   0.8668   0.03166   0.01925  -0.0671   0.4648   0.3935
   3.750   0.8891   0.03228   0.01992  -0.0667   0.4556   0.3944
   4.000   0.9149   0.03233   0.01990  -0.0659   0.4486   0.3946
   4.250   0.9367   0.03291   0.02053  -0.0653   0.4398   0.3947
   4.500   0.9612   0.03306   0.02065  -0.0644   0.4321   0.3949
   4.750   0.9825   0.03361   0.02125  -0.0637   0.4231   0.3951
   5.000   1.0060   0.03383   0.02147  -0.0628   0.4151   0.3955
   5.250   1.0261   0.03448   0.02218  -0.0620   0.4061   0.3964
   5.500   1.0489   0.03473   0.02244  -0.0611   0.3980   0.3978
   5.750   1.0673   0.03549   0.02333  -0.0602   0.3887   0.3991
   6.000   1.0900   0.03572   0.02355  -0.0592   0.3807   0.4005
   6.250   1.1060   0.03672   0.02467  -0.0583   0.3709   0.4016
   6.500   1.1289   0.03693   0.02486  -0.0573   0.3636   0.4027
   6.750   1.1421   0.03824   0.02632  -0.0564   0.3542   0.4035
   7.000   1.1634   0.03868   0.02678  -0.0554   0.3475   0.4045
   7.250   1.1762   0.04006   0.02827  -0.0545   0.3399   0.4055
   7.500   1.1912   0.04114   0.02943  -0.0535   0.3328   0.4068
   7.750   1.2170   0.04117   0.02943  -0.0526   0.3282   0.4087
   8.000   1.2118   0.04420   0.03272  -0.0514   0.3196   0.4095
   8.250   1.2320   0.04472   0.03326  -0.0504   0.3147   0.4118
   8.500   1.2372   0.04668   0.03534  -0.0493   0.3091   0.4136
   8.750   1.2194   0.05058   0.03941  -0.0480   0.3026   0.4141
   9.000   1.2355   0.05151   0.04042  -0.0471   0.2989   0.4173
   9.250   1.2657   0.05124   0.04020  -0.0462   0.2964   0.4241
   9.500   1.1499   0.06823   0.05738  -0.0504   0.2845   0.4142
   9.750   1.1690   0.06885   0.05808  -0.0495   0.2823   0.4178
  10.250   1.0925   0.08876   0.07811  -0.0563   0.2671   0.4149
  10.500   1.1126   0.08906   0.07850  -0.0552   0.2656   0.4192
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