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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 1,000,000
Max Cl/Cd: 56.01 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m1-il-1000000.txt
Download as CSV file: xf-m1-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.8721   0.02331   0.01955  -0.0049   1.0000   0.0076
  -8.000  -0.8510   0.02137   0.01733  -0.0043   1.0000   0.0078
  -7.750  -0.8348   0.01730   0.01264  -0.0030   1.0000   0.0082
  -7.500  -0.8128   0.01537   0.01044  -0.0023   1.0000   0.0088
  -7.250  -0.7871   0.01479   0.00979  -0.0021   1.0000   0.0094
  -7.000  -0.7610   0.01432   0.00925  -0.0019   1.0000   0.0100
  -6.750  -0.7347   0.01390   0.00876  -0.0018   1.0000   0.0108
  -6.500  -0.7079   0.01361   0.00842  -0.0016   1.0000   0.0115
  -6.250  -0.6835   0.01235   0.00698  -0.0011   1.0000   0.0124
  -6.000  -0.6574   0.01175   0.00636  -0.0009   1.0000   0.0136
  -5.750  -0.6302   0.01159   0.00618  -0.0009   1.0000   0.0147
  -5.500  -0.6031   0.01133   0.00590  -0.0008   1.0000   0.0160
  -5.250  -0.5756   0.01125   0.00579  -0.0007   1.0000   0.0170
  -5.000  -0.5502   0.01035   0.00483  -0.0004   1.0000   0.0193
  -4.750  -0.5229   0.01018   0.00468  -0.0004   1.0000   0.0210
  -4.500  -0.4957   0.00995   0.00443  -0.0003   1.0000   0.0229
  -4.250  -0.4684   0.00978   0.00424  -0.0002   1.0000   0.0243
  -4.000  -0.4405   0.00989   0.00437  -0.0002   1.0000   0.0251
  -3.750  -0.4155   0.00878   0.00316   0.0002   1.0000   0.0285
  -3.500  -0.3886   0.00846   0.00280   0.0004   1.0000   0.0301
  -3.250  -0.3617   0.00820   0.00253   0.0005   1.0000   0.0319
  -3.000  -0.3348   0.00795   0.00227   0.0007   1.0000   0.0336
  -2.750  -0.3081   0.00774   0.00204   0.0010   1.0000   0.0351
  -2.500  -0.2816   0.00759   0.00189   0.0013   1.0000   0.0361
  -2.250  -0.2554   0.00737   0.00163   0.0017   1.0000   0.0382
  -2.000  -0.2296   0.00720   0.00146   0.0021   1.0000   0.0408
  -1.750  -0.2041   0.00705   0.00135   0.0026   1.0000   0.0454
  -1.500  -0.1792   0.00665   0.00122   0.0031   1.0000   0.1157
  -1.250  -0.1548   0.00584   0.00109   0.0033   1.0000   0.3045
  -1.000  -0.1239   0.00506   0.00100   0.0020   0.9981   0.4986
  -0.750  -0.0840   0.00434   0.00092  -0.0010   0.9908   0.6799
  -0.500  -0.0492   0.00393   0.00091  -0.0025   0.9762   0.7892
  -0.250  -0.0212   0.00377   0.00093  -0.0020   0.9381   0.8530
   0.000   0.0001   0.00375   0.00093   0.0000   0.8953   0.8944
   0.250   0.0213   0.00376   0.00093   0.0020   0.8542   0.9384
   0.500   0.0492   0.00393   0.00092   0.0025   0.7892   0.9761
   0.750   0.0842   0.00434   0.00092   0.0009   0.6798   0.9911
   1.000   0.1244   0.00507   0.00100  -0.0021   0.4974   0.9983
   1.250   0.1548   0.00584   0.00109  -0.0033   0.3057   1.0000
   1.500   0.1792   0.00662   0.00122  -0.0031   0.1203   1.0000
   1.750   0.2041   0.00705   0.00135  -0.0026   0.0454   1.0000
   2.000   0.2296   0.00720   0.00146  -0.0021   0.0409   1.0000
   2.250   0.2554   0.00737   0.00163  -0.0017   0.0381   1.0000
   2.500   0.2816   0.00759   0.00189  -0.0013   0.0361   1.0000
   2.750   0.3081   0.00774   0.00204  -0.0010   0.0351   1.0000
   3.000   0.3348   0.00795   0.00227  -0.0007   0.0336   1.0000
   3.250   0.3617   0.00820   0.00254  -0.0005   0.0319   1.0000
   3.500   0.3886   0.00845   0.00280  -0.0004   0.0301   1.0000
   3.750   0.4155   0.00878   0.00316  -0.0002   0.0285   1.0000
   4.000   0.4405   0.00989   0.00437   0.0002   0.0251   1.0000
   4.250   0.4684   0.00978   0.00425   0.0002   0.0243   1.0000
   4.500   0.4957   0.00995   0.00443   0.0003   0.0229   1.0000
   4.750   0.5229   0.01019   0.00469   0.0004   0.0210   1.0000
   5.000   0.5502   0.01035   0.00484   0.0004   0.0193   1.0000
   5.250   0.5755   0.01126   0.00581   0.0007   0.0170   1.0000
   5.500   0.6031   0.01133   0.00590   0.0008   0.0160   1.0000
   5.750   0.6302   0.01158   0.00618   0.0009   0.0147   1.0000
   6.000   0.6575   0.01174   0.00635   0.0009   0.0136   1.0000
   6.250   0.6835   0.01235   0.00699   0.0011   0.0124   1.0000
   6.500   0.7079   0.01363   0.00844   0.0016   0.0115   1.0000
   6.750   0.7347   0.01391   0.00877   0.0018   0.0108   1.0000
   7.000   0.7610   0.01432   0.00925   0.0019   0.0100   1.0000
   7.250   0.7872   0.01477   0.00977   0.0021   0.0093   1.0000
   7.500   0.8128   0.01536   0.01043   0.0023   0.0088   1.0000
   7.750   0.8347   0.01735   0.01269   0.0030   0.0082   1.0000
   8.000   0.8510   0.02139   0.01734   0.0043   0.0078   1.0000
   8.250   0.8722   0.02329   0.01953   0.0049   0.0076   1.0000
   8.500   0.8831   0.02848   0.02533   0.0063   0.0073   1.0000
  10.500   0.7262   0.11837   0.11676  -0.0325   0.0116   1.0000
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