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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 500,000
Max Cl/Cd: 72.85 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m10-il-500000.txt
Download as CSV file: xf-m10-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5903   0.10136   0.09916   0.0124   1.0000   0.0191
  -8.500  -0.5864   0.09739   0.09521   0.0090   1.0000   0.0193
  -8.250  -0.5837   0.09320   0.09105   0.0056   1.0000   0.0193
  -8.000  -0.5802   0.08857   0.08642   0.0009   1.0000   0.0194
  -7.750  -0.5721   0.08343   0.08127  -0.0042   1.0000   0.0194
  -7.500  -0.5625   0.07815   0.07595  -0.0088   1.0000   0.0194
  -7.250  -0.5514   0.07275   0.07050  -0.0129   1.0000   0.0194
  -7.000  -0.5390   0.06721   0.06488  -0.0165   1.0000   0.0195
  -6.750  -0.5374   0.05765   0.05519  -0.0209   1.0000   0.0200
  -6.500  -0.5249   0.05385   0.05131  -0.0221   1.0000   0.0206
  -6.250  -0.5083   0.05120   0.04860  -0.0228   1.0000   0.0213
  -6.000  -0.4895   0.04808   0.04538  -0.0239   1.0000   0.0224
  -5.750  -0.4684   0.04417   0.04130  -0.0251   1.0000   0.0243
  -5.500  -0.4365   0.04115   0.03797  -0.0255   1.0000   0.0274
  -5.250  -0.4298   0.02552   0.02137  -0.0258   1.0000   0.0213
  -5.000  -0.4062   0.02291   0.01847  -0.0255   1.0000   0.0223
  -4.750  -0.3817   0.02017   0.01535  -0.0249   1.0000   0.0236
  -4.500  -0.3559   0.01843   0.01329  -0.0244   1.0000   0.0258
  -4.250  -0.3295   0.01754   0.01212  -0.0239   1.0000   0.0281
  -4.000  -0.3055   0.01477   0.00906  -0.0235   1.0000   0.0303
  -3.750  -0.2799   0.01409   0.00830  -0.0231   1.0000   0.0324
  -3.500  -0.2546   0.01359   0.00773  -0.0226   1.0000   0.0351
  -3.250  -0.2298   0.01283   0.00685  -0.0219   1.0000   0.0367
  -3.000  -0.2056   0.01224   0.00617  -0.0211   1.0000   0.0379
  -2.750  -0.1824   0.01155   0.00541  -0.0202   1.0000   0.0395
  -2.500  -0.1485   0.01039   0.00419  -0.0216   0.9964   0.0412
  -2.250  -0.1113   0.00975   0.00352  -0.0237   0.9907   0.0426
  -2.000  -0.0740   0.00928   0.00304  -0.0258   0.9828   0.0443
  -1.750  -0.0387   0.00892   0.00268  -0.0275   0.9704   0.0468
  -1.500  -0.0058   0.00859   0.00232  -0.0284   0.9534   0.0490
  -1.250   0.0228   0.00839   0.00207  -0.0283   0.9313   0.0511
  -1.000   0.0483   0.00815   0.00177  -0.0275   0.9074   0.0564
  -0.750   0.0734   0.00803   0.00161  -0.0266   0.8841   0.0656
  -0.500   0.0983   0.00746   0.00148  -0.0261   0.8616   0.2090
  -0.250   0.1094   0.00526   0.00148  -0.0227   0.8415   0.8872
   0.000   0.1641   0.00524   0.00148  -0.0280   0.8234   1.0000
   0.250   0.1896   0.00532   0.00141  -0.0274   0.8028   1.0000
   0.500   0.2154   0.00539   0.00137  -0.0269   0.7835   1.0000
   0.750   0.2414   0.00547   0.00135  -0.0264   0.7644   1.0000
   1.000   0.2675   0.00556   0.00134  -0.0259   0.7460   1.0000
   1.250   0.2938   0.00565   0.00134  -0.0255   0.7281   1.0000
   1.500   0.3200   0.00575   0.00136  -0.0251   0.7055   1.0000
   1.750   0.3462   0.00588   0.00137  -0.0247   0.6782   1.0000
   2.000   0.3727   0.00600   0.00140  -0.0243   0.6533   1.0000
   2.250   0.3994   0.00613   0.00145  -0.0240   0.6317   1.0000
   2.500   0.4262   0.00627   0.00152  -0.0238   0.6055   1.0000
   2.750   0.4527   0.00646   0.00159  -0.0235   0.5707   1.0000
   3.000   0.4792   0.00668   0.00167  -0.0232   0.5316   1.0000
   3.250   0.5056   0.00694   0.00177  -0.0230   0.4832   1.0000
   3.500   0.5317   0.00731   0.00193  -0.0228   0.4222   1.0000
   3.750   0.5565   0.00804   0.00218  -0.0226   0.3076   1.0000
   4.000   0.5776   0.00987   0.00289  -0.0224   0.0683   1.0000
   4.250   0.6038   0.01038   0.00333  -0.0222   0.0481   1.0000
   4.500   0.6300   0.01092   0.00387  -0.0219   0.0403   1.0000
   4.750   0.6563   0.01135   0.00438  -0.0217   0.0357   1.0000
   5.000   0.6822   0.01190   0.00496  -0.0215   0.0309   1.0000
   5.250   0.7060   0.01301   0.00617  -0.0209   0.0271   1.0000
   5.500   0.7321   0.01343   0.00664  -0.0206   0.0245   1.0000
   5.750   0.7576   0.01395   0.00720  -0.0203   0.0218   1.0000
   6.000   0.7806   0.01516   0.00845  -0.0196   0.0198   1.0000
   6.250   0.8026   0.01686   0.01027  -0.0187   0.0186   1.0000
   6.500   0.8276   0.01775   0.01127  -0.0181   0.0176   1.0000
   6.750   0.8525   0.01852   0.01213  -0.0177   0.0161   1.0000
   7.000   0.8769   0.01939   0.01308  -0.0172   0.0150   1.0000
   7.250   0.9005   0.02052   0.01433  -0.0167   0.0142   1.0000
   7.500   0.9224   0.02241   0.01636  -0.0160   0.0136   1.0000
   7.750   0.9402   0.02641   0.02074  -0.0147   0.0131   1.0000
   8.000   0.9563   0.03039   0.02514  -0.0133   0.0131   1.0000
   8.250   0.9751   0.03274   0.02777  -0.0122   0.0133   1.0000
  19.000   0.9785   0.25475   0.25242  -0.1043   0.0086   1.0000
  19.250   0.9858   0.25863   0.25630  -0.1076   0.0085   1.0000
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