NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA-M1 AIRFOIL (m1-il) Reynolds number: 1,000,000 Max Cl/Cd: 63.36 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m1-il-1000000-n5.txt Download as CSV file: xf-m1-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7560 0.12933 0.12775 0.0395 1.0000 0.0027
-11.250 -0.7594 0.12348 0.12191 0.0372 1.0000 0.0028
-10.750 -1.0826 0.03428 0.03166 -0.0077 1.0000 0.0022
-10.500 -1.0800 0.02911 0.02601 -0.0066 1.0000 0.0023
-10.250 -1.0668 0.02617 0.02274 -0.0057 1.0000 0.0023
-10.000 -1.0492 0.02411 0.02044 -0.0050 1.0000 0.0024
-9.750 -1.0295 0.02244 0.01855 -0.0044 1.0000 0.0025
-9.500 -1.0083 0.02107 0.01699 -0.0039 1.0000 0.0026
-9.250 -0.9859 0.01991 0.01567 -0.0035 1.0000 0.0027
-9.000 -0.9629 0.01886 0.01447 -0.0031 1.0000 0.0029
-8.750 -0.9394 0.01787 0.01332 -0.0027 1.0000 0.0031
-8.500 -0.9156 0.01689 0.01218 -0.0024 1.0000 0.0033
-8.250 -0.8913 0.01600 0.01114 -0.0020 1.0000 0.0034
-8.000 -0.8666 0.01519 0.01019 -0.0017 1.0000 0.0035
-7.750 -0.8422 0.01420 0.00903 -0.0013 1.0000 0.0037
-7.500 -0.8173 0.01338 0.00808 -0.0010 1.0000 0.0040
-7.250 -0.7916 0.01281 0.00743 -0.0008 1.0000 0.0044
-7.000 -0.7656 0.01229 0.00684 -0.0006 1.0000 0.0047
-6.750 -0.7394 0.01179 0.00627 -0.0004 1.0000 0.0051
-6.500 -0.7131 0.01134 0.00575 -0.0003 1.0000 0.0054
-6.250 -0.6867 0.01087 0.00520 -0.0001 1.0000 0.0061
-6.000 -0.6602 0.01042 0.00472 0.0000 1.0000 0.0071
-5.750 -0.6334 0.01006 0.00430 0.0001 1.0000 0.0081
-5.500 -0.6065 0.00973 0.00391 0.0002 1.0000 0.0090
-5.250 -0.5796 0.00939 0.00361 0.0003 1.0000 0.0119
-5.000 -0.5524 0.00915 0.00341 0.0004 1.0000 0.0158
-4.750 -0.5250 0.00903 0.00335 0.0004 1.0000 0.0208
-4.500 -0.4975 0.00892 0.00322 0.0004 1.0000 0.0232
-4.250 -0.4700 0.00885 0.00314 0.0004 1.0000 0.0246
-4.000 -0.4425 0.00878 0.00309 0.0004 1.0000 0.0253
-3.750 -0.4152 0.00868 0.00299 0.0005 1.0000 0.0256
-3.500 -0.3880 0.00855 0.00284 0.0006 1.0000 0.0258
-3.250 -0.3613 0.00821 0.00246 0.0007 1.0000 0.0267
-3.000 -0.3345 0.00795 0.00216 0.0009 1.0000 0.0275
-2.750 -0.3077 0.00775 0.00193 0.0012 1.0000 0.0275
-2.500 -0.2813 0.00757 0.00174 0.0014 1.0000 0.0276
-2.250 -0.2492 0.00740 0.00155 0.0004 0.9960 0.0279
-2.000 -0.2166 0.00724 0.00137 -0.0007 0.9899 0.0287
-1.750 -0.1840 0.00710 0.00124 -0.0017 0.9816 0.0311
-1.500 -0.1519 0.00697 0.00115 -0.0027 0.9687 0.0355
-1.250 -0.1250 0.00687 0.00106 -0.0022 0.9208 0.0620
-1.000 -0.1030 0.00683 0.00095 -0.0007 0.8461 0.1116
-0.750 -0.0774 0.00668 0.00085 -0.0003 0.7935 0.1847
-0.500 -0.0521 0.00658 0.00077 0.0000 0.7043 0.2880
-0.250 -0.0267 0.00672 0.00072 0.0001 0.5537 0.3876
0.000 0.0000 0.00679 0.00071 0.0000 0.4593 0.4628
0.250 0.0268 0.00672 0.00072 -0.0001 0.3873 0.5559
0.500 0.0522 0.00658 0.00077 0.0000 0.2883 0.7041
0.750 0.0775 0.00669 0.00085 0.0003 0.1835 0.7925
1.000 0.1031 0.00684 0.00095 0.0007 0.1106 0.8453
1.250 0.1249 0.00689 0.00106 0.0022 0.0542 0.9221
1.500 0.1516 0.00697 0.00115 0.0027 0.0358 0.9675
1.750 0.1840 0.00710 0.00124 0.0017 0.0312 0.9815
2.000 0.2166 0.00724 0.00137 0.0007 0.0287 0.9898
2.250 0.2493 0.00740 0.00155 -0.0005 0.0279 0.9960
2.500 0.2813 0.00758 0.00174 -0.0015 0.0276 1.0000
2.750 0.3078 0.00775 0.00194 -0.0012 0.0275 1.0000
3.000 0.3346 0.00795 0.00216 -0.0010 0.0275 1.0000
3.250 0.3613 0.00821 0.00246 -0.0008 0.0267 1.0000
3.500 0.3880 0.00855 0.00285 -0.0006 0.0258 1.0000
3.750 0.4152 0.00870 0.00301 -0.0005 0.0256 1.0000
4.000 0.4426 0.00879 0.00309 -0.0004 0.0253 1.0000
4.250 0.4700 0.00885 0.00315 -0.0004 0.0246 1.0000
4.500 0.4976 0.00893 0.00322 -0.0004 0.0233 1.0000
4.750 0.5250 0.00903 0.00335 -0.0004 0.0208 1.0000
5.000 0.5525 0.00915 0.00341 -0.0004 0.0158 1.0000
5.250 0.5796 0.00939 0.00361 -0.0003 0.0119 1.0000
5.500 0.6065 0.00973 0.00391 -0.0002 0.0090 1.0000
5.750 0.6335 0.01006 0.00430 -0.0001 0.0081 1.0000
6.000 0.6602 0.01042 0.00472 0.0000 0.0071 1.0000
6.250 0.6867 0.01087 0.00520 0.0001 0.0061 1.0000
6.500 0.7131 0.01134 0.00574 0.0003 0.0054 1.0000
6.750 0.7394 0.01179 0.00627 0.0004 0.0051 1.0000
7.000 0.7656 0.01229 0.00684 0.0006 0.0047 1.0000
7.250 0.7915 0.01281 0.00744 0.0008 0.0044 1.0000
7.500 0.8173 0.01339 0.00808 0.0010 0.0041 1.0000
7.750 0.8423 0.01419 0.00903 0.0013 0.0038 1.0000
8.000 0.8666 0.01519 0.01019 0.0017 0.0035 1.0000
8.250 0.8913 0.01599 0.01113 0.0020 0.0034 1.0000
8.500 0.9156 0.01689 0.01219 0.0024 0.0033 1.0000
8.750 0.9394 0.01786 0.01331 0.0027 0.0031 1.0000
9.000 0.9629 0.01885 0.01446 0.0031 0.0029 1.0000
9.250 0.9859 0.01991 0.01568 0.0035 0.0027 1.0000
9.500 1.0083 0.02106 0.01698 0.0039 0.0026 1.0000
9.750 1.0294 0.02245 0.01857 0.0044 0.0025 1.0000
10.000 1.0492 0.02410 0.02042 0.0050 0.0024 1.0000
10.250 1.0667 0.02616 0.02274 0.0057 0.0023 1.0000
10.500 1.0799 0.02912 0.02602 0.0066 0.0023 1.0000
10.750 1.0824 0.03433 0.03171 0.0077 0.0022 1.0000
11.250 0.7602 0.12354 0.12196 -0.0373 0.0028 1.0000
11.500 0.7569 0.12940 0.12781 -0.0397 0.0027 1.0000
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Polar data table (+)
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