NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA-M1 AIRFOIL (m1-il) Reynolds number: 200,000 Max Cl/Cd: 32.16 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m1-il-200000-n5.txt Download as CSV file: xf-m1-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.7323 0.08747 0.08413 0.0130 1.0000 0.0114
-8.750 -0.7396 0.08070 0.07741 0.0069 1.0000 0.0113
-8.500 -0.7485 0.07302 0.06969 -0.0001 1.0000 0.0110
-8.250 -0.7526 0.06570 0.06224 -0.0046 1.0000 0.0109
-8.000 -0.7532 0.05884 0.05517 -0.0072 1.0000 0.0111
-7.750 -0.7505 0.05221 0.04826 -0.0086 1.0000 0.0114
-7.500 -0.7449 0.04560 0.04125 -0.0090 1.0000 0.0117
-7.250 -0.7369 0.03905 0.03417 -0.0085 1.0000 0.0119
-7.000 -0.7245 0.03335 0.02784 -0.0076 1.0000 0.0122
-6.750 -0.7074 0.02890 0.02276 -0.0065 1.0000 0.0128
-6.500 -0.6865 0.02567 0.01897 -0.0056 1.0000 0.0136
-6.250 -0.6647 0.02296 0.01583 -0.0049 1.0000 0.0152
-6.000 -0.6407 0.02189 0.01457 -0.0046 1.0000 0.0169
-5.750 -0.6159 0.02038 0.01281 -0.0041 1.0000 0.0188
-5.500 -0.5905 0.01903 0.01119 -0.0035 1.0000 0.0220
-5.250 -0.5643 0.01829 0.01022 -0.0031 1.0000 0.0252
-5.000 -0.5400 0.01679 0.00862 -0.0026 1.0000 0.0283
-4.750 -0.5144 0.01618 0.00792 -0.0023 1.0000 0.0329
-4.500 -0.4885 0.01552 0.00712 -0.0019 1.0000 0.0360
-4.250 -0.4637 0.01462 0.00618 -0.0015 1.0000 0.0398
-4.000 -0.4382 0.01399 0.00550 -0.0012 1.0000 0.0420
-3.750 -0.4123 0.01351 0.00497 -0.0009 1.0000 0.0450
-3.500 -0.3864 0.01301 0.00437 -0.0005 1.0000 0.0463
-3.250 -0.3604 0.01257 0.00383 -0.0002 1.0000 0.0473
-3.000 -0.3342 0.01220 0.00339 0.0002 1.0000 0.0487
-2.750 -0.3082 0.01179 0.00294 0.0005 1.0000 0.0536
-2.500 -0.2823 0.01141 0.00261 0.0008 1.0000 0.0676
-2.250 -0.2575 0.01071 0.00229 0.0011 1.0000 0.1534
-2.000 -0.2335 0.00990 0.00205 0.0013 1.0000 0.2920
-1.750 -0.2113 0.00894 0.00189 0.0020 1.0000 0.4870
-1.500 -0.1898 0.00825 0.00185 0.0033 1.0000 0.6374
-1.250 -0.1697 0.00777 0.00191 0.0054 1.0000 0.7690
-1.000 -0.1466 0.00757 0.00205 0.0075 1.0000 0.8900
-0.750 -0.0899 0.00764 0.00219 0.0021 1.0000 0.9802
-0.500 -0.0410 0.00767 0.00217 -0.0026 1.0000 1.0000
-0.250 -0.0202 0.00763 0.00211 -0.0014 1.0000 1.0000
0.000 0.0000 0.00762 0.00209 0.0000 1.0000 1.0000
0.250 0.0202 0.00763 0.00211 0.0014 1.0000 1.0000
0.500 0.0410 0.00767 0.00217 0.0026 1.0000 1.0000
0.750 0.0904 0.00764 0.00218 -0.0022 0.9797 1.0000
1.000 0.1467 0.00757 0.00205 -0.0075 0.8894 1.0000
1.250 0.1697 0.00777 0.00190 -0.0054 0.7675 1.0000
1.500 0.1899 0.00825 0.00185 -0.0033 0.6375 1.0000
1.750 0.2114 0.00894 0.00189 -0.0020 0.4871 1.0000
2.000 0.2336 0.00990 0.00205 -0.0014 0.2920 1.0000
2.250 0.2575 0.01072 0.00229 -0.0011 0.1520 1.0000
2.500 0.2823 0.01141 0.00261 -0.0008 0.0671 1.0000
2.750 0.3083 0.01179 0.00294 -0.0005 0.0535 1.0000
3.000 0.3343 0.01220 0.00339 -0.0002 0.0486 1.0000
3.250 0.3605 0.01257 0.00384 0.0002 0.0472 1.0000
3.500 0.3865 0.01301 0.00437 0.0005 0.0463 1.0000
3.750 0.4123 0.01352 0.00497 0.0009 0.0450 1.0000
4.000 0.4382 0.01400 0.00551 0.0012 0.0420 1.0000
4.250 0.4637 0.01462 0.00618 0.0015 0.0398 1.0000
4.500 0.4885 0.01552 0.00712 0.0019 0.0360 1.0000
4.750 0.5144 0.01618 0.00793 0.0023 0.0329 1.0000
5.000 0.5400 0.01679 0.00862 0.0026 0.0283 1.0000
5.250 0.5644 0.01829 0.01021 0.0031 0.0252 1.0000
5.500 0.5905 0.01903 0.01120 0.0035 0.0220 1.0000
5.750 0.6159 0.02038 0.01280 0.0041 0.0188 1.0000
6.000 0.6407 0.02188 0.01456 0.0046 0.0169 1.0000
6.250 0.6647 0.02297 0.01584 0.0049 0.0152 1.0000
6.500 0.6865 0.02567 0.01897 0.0056 0.0136 1.0000
6.750 0.7074 0.02890 0.02276 0.0065 0.0128 1.0000
7.000 0.7245 0.03334 0.02784 0.0076 0.0122 1.0000
7.250 0.7369 0.03904 0.03416 0.0085 0.0119 1.0000
7.500 0.7449 0.04560 0.04125 0.0090 0.0117 1.0000
7.750 0.7505 0.05222 0.04826 0.0086 0.0114 1.0000
8.000 0.7533 0.05884 0.05518 0.0072 0.0111 1.0000
8.250 0.7529 0.06566 0.06220 0.0046 0.0109 1.0000
8.500 0.7486 0.07305 0.06971 0.0001 0.0110 1.0000
8.750 0.7399 0.08072 0.07743 -0.0070 0.0113 1.0000
9.000 0.7326 0.08751 0.08418 -0.0131 0.0114 1.0000
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Polar data table (+)
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