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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 50,000
Max Cl/Cd: 34.37 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m10-il-50000.txt
Download as CSV file: xf-m10-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M10 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5906   0.11194   0.10526   0.0130   1.0000   0.1764
  -8.250  -0.5752   0.10664   0.09995   0.0147   1.0000   0.1864
  -8.000  -0.5922   0.10530   0.09875   0.0089   1.0000   0.1910
  -7.750  -0.5761   0.10008   0.09353   0.0114   1.0000   0.2033
  -7.500  -0.5685   0.09586   0.08935   0.0116   1.0000   0.2131
  -7.250  -0.5667   0.09218   0.08573   0.0095   1.0000   0.2239
  -7.000  -0.5625   0.08848   0.08206   0.0080   1.0000   0.2374
  -6.750  -0.5547   0.08469   0.07827   0.0081   1.0000   0.2541
  -6.500  -0.5548   0.08178   0.07538   0.0046   1.0000   0.2757
  -6.250  -0.5468   0.07813   0.07177   0.0064   1.0000   0.3038
  -6.000  -0.5354   0.07423   0.06794   0.0113   1.0000   0.3379
  -5.750  -0.5297   0.07107   0.06483   0.0142   1.0000   0.3784
  -5.500  -0.5189   0.06790   0.06172   0.0200   1.0000   0.4261
  -5.250  -0.2061   0.04887   0.04162   0.0139   1.0000   1.0000
  -5.000  -0.1935   0.04624   0.03900   0.0123   1.0000   1.0000
  -4.750  -0.2213   0.04653   0.03943   0.0204   1.0000   0.9751
  -4.500  -0.2886   0.04879   0.04199   0.0347   1.0000   0.9020
  -4.250  -0.3561   0.04955   0.04306   0.0428   1.0000   0.8043
  -4.000  -0.3384   0.03848   0.02953  -0.0265   1.0000   0.1922
  -3.750  -0.3083   0.03480   0.02535  -0.0269   1.0000   0.1709
  -3.500  -0.2776   0.03222   0.02195  -0.0271   1.0000   0.1625
  -3.250  -0.2502   0.02998   0.01924  -0.0267   1.0000   0.1630
  -3.000  -0.2232   0.02767   0.01668  -0.0263   1.0000   0.1611
  -2.750  -0.1957   0.02580   0.01447  -0.0257   1.0000   0.1609
  -2.500  -0.1683   0.02427   0.01254  -0.0250   1.0000   0.1627
  -2.250  -0.1424   0.02278   0.01108  -0.0244   1.0000   0.1721
  -2.000  -0.1158   0.02157   0.00980  -0.0237   1.0000   0.1847
  -1.750  -0.0876   0.02041   0.00857  -0.0231   1.0000   0.1969
  -1.500  -0.0618   0.01934   0.00751  -0.0222   1.0000   0.2193
  -1.250  -0.0113   0.01489   0.00569  -0.0243   1.0000   1.0000
  -1.000   0.0095   0.01495   0.00525  -0.0229   1.0000   1.0000
  -0.750   0.0294   0.01503   0.00506  -0.0216   1.0000   1.0000
  -0.500   0.0487   0.01515   0.00497  -0.0204   1.0000   1.0000
  -0.250   0.0676   0.01532   0.00498  -0.0192   1.0000   1.0000
   0.000   0.0863   0.01554   0.00504  -0.0181   1.0000   1.0000
   0.250   0.1049   0.01581   0.00521  -0.0172   1.0000   1.0000
   0.500   0.1233   0.01615   0.00545  -0.0163   1.0000   1.0000
   0.750   0.1417   0.01654   0.00577  -0.0157   1.0000   1.0000
   1.000   0.1600   0.01700   0.00618  -0.0152   1.0000   1.0000
   1.250   0.1782   0.01752   0.00668  -0.0149   1.0000   1.0000
   1.500   0.1964   0.01812   0.00725  -0.0147   1.0000   1.0000
   1.750   0.2144   0.01878   0.00792  -0.0147   1.0000   1.0000
   2.000   0.2322   0.01953   0.00869  -0.0149   1.0000   1.0000
   2.250   0.2498   0.02035   0.00955  -0.0152   1.0000   1.0000
   2.500   0.3116   0.02132   0.01068  -0.0236   0.9799   1.0000
   2.750   0.3756   0.02215   0.01174  -0.0320   0.9569   1.0000
   3.000   0.4331   0.02282   0.01272  -0.0387   0.9318   1.0000
   3.250   0.4925   0.02332   0.01357  -0.0451   0.9063   1.0000
   3.500   0.5470   0.02365   0.01428  -0.0497   0.8799   1.0000
   3.750   0.5927   0.02386   0.01492  -0.0519   0.8517   1.0000
   4.000   0.6318   0.02366   0.01509  -0.0512   0.8177   1.0000
   4.250   0.6527   0.02156   0.01308  -0.0412   0.7487   1.0000
   4.500   0.6651   0.02041   0.01189  -0.0327   0.6842   1.0000
   4.750   0.6781   0.01973   0.01128  -0.0259   0.6056   1.0000
   5.000   0.6749   0.02133   0.01072  -0.0170   0.2556   1.0000
   5.250   0.6928   0.02430   0.01281  -0.0154   0.1765   1.0000
   5.500   0.7152   0.02633   0.01463  -0.0140   0.1474   1.0000
   5.750   0.7417   0.02843   0.01667  -0.0127   0.1330   1.0000
   6.000   0.7693   0.03072   0.01904  -0.0117   0.1218   1.0000
   6.250   0.7960   0.03335   0.02186  -0.0109   0.1134   1.0000
   6.500   0.8229   0.03631   0.02529  -0.0099   0.1110   1.0000
   6.750   0.8472   0.03959   0.02909  -0.0089   0.1093   1.0000
   7.000   0.8689   0.04290   0.03270  -0.0081   0.1053   1.0000
   7.250   0.8879   0.04681   0.03716  -0.0071   0.1051   1.0000
   7.500   0.9009   0.05181   0.04304  -0.0061   0.1105   1.0000
   7.750   0.9154   0.05701   0.04856  -0.0056   0.1150   1.0000
   8.000   0.9150   0.06330   0.05573  -0.0059   0.1255   1.0000
   8.250   0.9085   0.07045   0.06343  -0.0076   0.1377   1.0000
   8.500   0.8147   0.06590   0.05945  -0.0018   0.1362   1.0000
   8.750   0.8012   0.07261   0.06626  -0.0028   0.1452   1.0000
   9.000   0.7562   0.07989   0.07354  -0.0069   0.1473   1.0000
   9.250   0.7197   0.09094   0.08452  -0.0148   0.1675   1.0000
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