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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 100,000
Max Cl/Cd: 24.86 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m1-il-100000-n5.txt
Download as CSV file: xf-m1-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-M1 AIRFOIL                                 
  
 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  -0.7138   0.08829   0.08355   0.0098   1.0000   0.0221
  -8.500  -0.7195   0.08200   0.07731   0.0040   1.0000   0.0218
  -8.250  -0.7238   0.07567   0.07093  -0.0008   1.0000   0.0216
  -8.000  -0.7256   0.06941   0.06455  -0.0044   1.0000   0.0215
  -7.750  -0.7249   0.06334   0.05829  -0.0070   1.0000   0.0216
  -7.500  -0.7212   0.05743   0.05210  -0.0086   1.0000   0.0221
  -7.250  -0.7144   0.05167   0.04596  -0.0094   1.0000   0.0228
  -7.000  -0.7040   0.04630   0.04011  -0.0094   1.0000   0.0235
  -6.750  -0.6907   0.04129   0.03454  -0.0090   1.0000   0.0240
  -6.500  -0.6740   0.03681   0.02936  -0.0082   1.0000   0.0246
  -6.250  -0.6539   0.03327   0.02510  -0.0073   1.0000   0.0256
  -6.000  -0.6343   0.03029   0.02186  -0.0070   1.0000   0.0282
  -5.750  -0.6112   0.02823   0.01947  -0.0064   1.0000   0.0311
  -5.500  -0.5865   0.02574   0.01650  -0.0056   1.0000   0.0338
  -5.250  -0.5616   0.02377   0.01415  -0.0050   1.0000   0.0383
  -5.000  -0.5368   0.02215   0.01242  -0.0045   1.0000   0.0422
  -4.750  -0.5109   0.02111   0.01106  -0.0040   1.0000   0.0480
  -4.500  -0.4865   0.01957   0.00950  -0.0034   1.0000   0.0520
  -4.250  -0.4615   0.01865   0.00851  -0.0029   1.0000   0.0569
  -4.000  -0.4362   0.01788   0.00759  -0.0024   1.0000   0.0613
  -3.750  -0.4117   0.01695   0.00656  -0.0017   1.0000   0.0633
  -3.500  -0.3870   0.01618   0.00572  -0.0012   1.0000   0.0668
  -3.250  -0.3616   0.01557   0.00501  -0.0007   1.0000   0.0727
  -3.000  -0.3363   0.01489   0.00435  -0.0002   1.0000   0.0864
  -2.750  -0.3123   0.01388   0.00375   0.0003   1.0000   0.1680
  -2.500  -0.2904   0.01258   0.00335   0.0008   1.0000   0.3735
  -2.250  -0.2705   0.01146   0.00321   0.0025   1.0000   0.5958
  -2.000  -0.2518   0.01080   0.00341   0.0062   1.0000   0.8225
  -1.750  -0.1821   0.01094   0.00357  -0.0005   1.0000   0.9774
  -1.500  -0.1278   0.01089   0.00333  -0.0062   1.0000   1.0000
  -1.250  -0.1061   0.01076   0.00309  -0.0053   1.0000   1.0000
  -1.000  -0.0846   0.01065   0.00291  -0.0044   1.0000   1.0000
  -0.750  -0.0631   0.01058   0.00278  -0.0034   1.0000   1.0000
  -0.500  -0.0419   0.01053   0.00269  -0.0023   1.0000   1.0000
  -0.250  -0.0209   0.01049   0.00263  -0.0012   1.0000   1.0000
   0.000   0.0000   0.01048   0.00261   0.0000   1.0000   1.0000
   0.250   0.0209   0.01049   0.00263   0.0012   1.0000   1.0000
   0.500   0.0419   0.01053   0.00269   0.0023   1.0000   1.0000
   0.750   0.0631   0.01058   0.00278   0.0034   1.0000   1.0000
   1.000   0.0846   0.01065   0.00291   0.0044   1.0000   1.0000
   1.250   0.1062   0.01076   0.00309   0.0053   1.0000   1.0000
   1.500   0.1278   0.01089   0.00333   0.0062   1.0000   1.0000
   1.750   0.1821   0.01094   0.00357   0.0005   0.9773   1.0000
   2.000   0.2518   0.01080   0.00341  -0.0062   0.8223   1.0000
   2.250   0.2705   0.01146   0.00320  -0.0025   0.5947   1.0000
   2.500   0.2904   0.01258   0.00335  -0.0008   0.3723   1.0000
   2.750   0.3124   0.01388   0.00375  -0.0003   0.1673   1.0000
   3.000   0.3363   0.01489   0.00435   0.0002   0.0863   1.0000
   3.250   0.3616   0.01557   0.00501   0.0007   0.0726   1.0000
   3.500   0.3870   0.01618   0.00572   0.0012   0.0667   1.0000
   3.750   0.4118   0.01696   0.00656   0.0017   0.0632   1.0000
   4.000   0.4362   0.01788   0.00759   0.0024   0.0613   1.0000
   4.250   0.4615   0.01866   0.00851   0.0029   0.0570   1.0000
   4.500   0.4865   0.01957   0.00950   0.0034   0.0520   1.0000
   4.750   0.5109   0.02111   0.01106   0.0040   0.0480   1.0000
   5.000   0.5368   0.02215   0.01242   0.0045   0.0422   1.0000
   5.250   0.5616   0.02377   0.01415   0.0050   0.0383   1.0000
   5.500   0.5865   0.02574   0.01650   0.0056   0.0338   1.0000
   5.750   0.6112   0.02822   0.01947   0.0064   0.0311   1.0000
   6.000   0.6343   0.03029   0.02185   0.0069   0.0282   1.0000
   6.250   0.6539   0.03326   0.02510   0.0073   0.0256   1.0000
   6.500   0.6741   0.03681   0.02936   0.0082   0.0246   1.0000
   6.750   0.6907   0.04130   0.03455   0.0090   0.0240   1.0000
   7.000   0.7040   0.04630   0.04011   0.0094   0.0235   1.0000
   7.250   0.7144   0.05167   0.04596   0.0094   0.0228   1.0000
   7.500   0.7213   0.05743   0.05211   0.0086   0.0221   1.0000
   7.750   0.7250   0.06335   0.05831   0.0069   0.0216   1.0000
   8.000   0.7258   0.06943   0.06457   0.0044   0.0215   1.0000
   8.250   0.7240   0.07569   0.07095   0.0007   0.0216   1.0000
   8.500   0.7198   0.08203   0.07733  -0.0041   0.0218   1.0000
   8.750   0.7141   0.08833   0.08360  -0.0099   0.0221   1.0000
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