NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M2 AIRFOIL (m2-il) Reynolds number: 1,000,000 Max Cl/Cd: 70.24 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m2-il-1000000-n5.txt Download as CSV file: xf-m2-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M2 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
------ -------- --------- --------- -------- -------- --------
-15.000 -1.2588 0.05836 0.05624 -0.0133 1.0000 0.0035
-14.750 -1.3003 0.04837 0.04595 -0.0194 1.0000 0.0034
-14.500 -1.3214 0.04327 0.04063 -0.0201 1.0000 0.0034
-14.250 -1.3344 0.03981 0.03699 -0.0188 1.0000 0.0035
-14.000 -1.3429 0.03714 0.03414 -0.0164 1.0000 0.0035
-13.750 -1.3436 0.03487 0.03170 -0.0142 1.0000 0.0035
-13.500 -1.3379 0.03279 0.02944 -0.0127 1.0000 0.0036
-13.250 -1.3287 0.03093 0.02741 -0.0113 1.0000 0.0037
-13.000 -1.3167 0.02926 0.02556 -0.0101 1.0000 0.0038
-12.750 -1.3027 0.02773 0.02387 -0.0090 1.0000 0.0039
-12.500 -1.2870 0.02633 0.02231 -0.0080 1.0000 0.0040
-12.250 -1.2697 0.02504 0.02087 -0.0072 1.0000 0.0041
-12.000 -1.2512 0.02387 0.01955 -0.0064 1.0000 0.0042
-11.750 -1.2317 0.02279 0.01833 -0.0057 1.0000 0.0043
-11.500 -1.2111 0.02181 0.01722 -0.0050 1.0000 0.0044
-11.250 -1.1898 0.02088 0.01617 -0.0045 1.0000 0.0045
-11.000 -1.1695 0.01971 0.01483 -0.0038 1.0000 0.0048
-10.750 -1.1474 0.01880 0.01383 -0.0033 1.0000 0.0050
-10.500 -1.1243 0.01803 0.01298 -0.0029 1.0000 0.0053
-10.250 -1.1008 0.01734 0.01220 -0.0025 1.0000 0.0056
-10.000 -1.0767 0.01670 0.01148 -0.0022 1.0000 0.0059
-9.750 -1.0522 0.01611 0.01080 -0.0019 1.0000 0.0062
-9.500 -1.0274 0.01556 0.01018 -0.0017 1.0000 0.0065
-9.250 -1.0029 0.01491 0.00945 -0.0014 1.0000 0.0070
-9.000 -0.9780 0.01431 0.00878 -0.0011 1.0000 0.0076
-8.750 -0.9526 0.01380 0.00823 -0.0009 1.0000 0.0082
-8.500 -0.9270 0.01334 0.00771 -0.0008 1.0000 0.0088
-8.250 -0.9010 0.01292 0.00723 -0.0006 1.0000 0.0093
-8.000 -0.8751 0.01246 0.00675 -0.0005 1.0000 0.0107
-7.750 -0.8489 0.01209 0.00638 -0.0004 1.0000 0.0122
-7.500 -0.8224 0.01175 0.00603 -0.0003 1.0000 0.0137
-7.250 -0.7957 0.01147 0.00578 -0.0003 1.0000 0.0161
-7.000 -0.7687 0.01124 0.00556 -0.0003 1.0000 0.0179
-6.750 -0.7416 0.01104 0.00533 -0.0003 1.0000 0.0191
-6.500 -0.7144 0.01089 0.00517 -0.0003 1.0000 0.0197
-6.250 -0.6880 0.01053 0.00477 -0.0001 1.0000 0.0209
-6.000 -0.6616 0.01022 0.00444 0.0000 1.0000 0.0224
-5.750 -0.6353 0.00998 0.00419 0.0002 1.0000 0.0236
-5.500 -0.6093 0.00976 0.00397 0.0005 1.0000 0.0248
-5.250 -0.5783 0.00953 0.00372 -0.0004 0.9969 0.0260
-5.000 -0.5463 0.00932 0.00347 -0.0014 0.9935 0.0271
-4.750 -0.5148 0.00913 0.00327 -0.0023 0.9887 0.0280
-4.500 -0.4836 0.00894 0.00307 -0.0032 0.9836 0.0285
-4.250 -0.4543 0.00879 0.00290 -0.0035 0.9772 0.0288
-4.000 -0.4269 0.00858 0.00264 -0.0034 0.9701 0.0299
-3.750 -0.4003 0.00840 0.00244 -0.0031 0.9635 0.0311
-3.500 -0.3734 0.00824 0.00226 -0.0028 0.9576 0.0323
-3.250 -0.3479 0.00811 0.00211 -0.0022 0.9493 0.0336
-3.000 -0.3222 0.00798 0.00197 -0.0016 0.9371 0.0350
-2.750 -0.2958 0.00785 0.00183 -0.0013 0.9245 0.0367
-2.500 -0.2688 0.00773 0.00170 -0.0010 0.9120 0.0391
-2.250 -0.2428 0.00753 0.00156 -0.0006 0.8924 0.0616
-2.000 -0.2175 0.00744 0.00145 0.0001 0.8570 0.0793
-1.750 -0.1914 0.00748 0.00135 0.0006 0.8124 0.0887
-1.500 -0.1641 0.00751 0.00128 0.0007 0.7769 0.0963
-1.250 -0.1364 0.00755 0.00122 0.0008 0.7380 0.1027
-1.000 -0.1085 0.00758 0.00116 0.0007 0.7014 0.1101
-0.750 -0.0808 0.00771 0.00111 0.0007 0.6410 0.1180
-0.500 -0.0534 0.00793 0.00108 0.0006 0.5510 0.1324
-0.250 -0.0261 0.00793 0.00104 0.0003 0.4730 0.2020
0.000 0.0000 0.00733 0.00094 0.0000 0.4179 0.4178
0.250 0.0261 0.00793 0.00104 -0.0004 0.2008 0.4738
0.500 0.0534 0.00794 0.00108 -0.0006 0.1326 0.5491
0.750 0.0809 0.00771 0.00111 -0.0007 0.1180 0.6396
1.000 0.1085 0.00758 0.00116 -0.0007 0.1102 0.7011
1.250 0.1365 0.00755 0.00122 -0.0008 0.1028 0.7379
1.500 0.1642 0.00751 0.00128 -0.0007 0.0966 0.7772
1.750 0.1915 0.00748 0.00135 -0.0006 0.0888 0.8123
2.000 0.2175 0.00744 0.00145 -0.0001 0.0790 0.8570
2.250 0.2429 0.00753 0.00156 0.0005 0.0616 0.8925
2.500 0.2689 0.00773 0.00170 0.0010 0.0391 0.9122
2.750 0.2958 0.00785 0.00183 0.0012 0.0368 0.9244
3.000 0.3223 0.00798 0.00197 0.0016 0.0351 0.9371
3.250 0.3479 0.00811 0.00211 0.0022 0.0336 0.9493
3.500 0.3735 0.00824 0.00226 0.0028 0.0323 0.9576
3.750 0.4004 0.00840 0.00244 0.0030 0.0311 0.9635
4.000 0.4269 0.00858 0.00265 0.0034 0.0299 0.9701
4.250 0.4543 0.00879 0.00291 0.0035 0.0288 0.9772
4.500 0.4836 0.00895 0.00307 0.0031 0.0285 0.9835
4.750 0.5148 0.00913 0.00327 0.0023 0.0280 0.9887
5.000 0.5463 0.00932 0.00348 0.0014 0.0271 0.9935
5.250 0.5784 0.00953 0.00372 0.0004 0.0260 0.9970
5.500 0.6093 0.00976 0.00397 -0.0005 0.0248 1.0000
5.750 0.6353 0.00998 0.00419 -0.0002 0.0236 1.0000
6.000 0.6617 0.01022 0.00444 0.0000 0.0223 1.0000
6.250 0.6881 0.01053 0.00478 0.0001 0.0210 1.0000
6.500 0.7144 0.01089 0.00517 0.0003 0.0197 1.0000
6.750 0.7417 0.01104 0.00534 0.0003 0.0191 1.0000
7.000 0.7688 0.01125 0.00556 0.0003 0.0179 1.0000
7.250 0.7958 0.01147 0.00578 0.0003 0.0161 1.0000
7.500 0.8225 0.01175 0.00603 0.0003 0.0137 1.0000
7.750 0.8490 0.01209 0.00639 0.0004 0.0122 1.0000
8.000 0.8752 0.01246 0.00675 0.0004 0.0107 1.0000
8.250 0.9011 0.01292 0.00723 0.0006 0.0093 1.0000
8.500 0.9271 0.01334 0.00771 0.0007 0.0088 1.0000
8.750 0.9528 0.01381 0.00823 0.0009 0.0082 1.0000
9.000 0.9781 0.01431 0.00878 0.0011 0.0076 1.0000
9.250 1.0030 0.01491 0.00945 0.0014 0.0070 1.0000
9.500 1.0276 0.01556 0.01018 0.0017 0.0065 1.0000
9.750 1.0524 0.01611 0.01080 0.0019 0.0062 1.0000
10.000 1.0768 0.01671 0.01149 0.0022 0.0059 1.0000
10.250 1.1009 0.01734 0.01220 0.0025 0.0056 1.0000
10.500 1.1245 0.01804 0.01298 0.0029 0.0053 1.0000
10.750 1.1476 0.01880 0.01383 0.0033 0.0050 1.0000
11.000 1.1697 0.01971 0.01484 0.0038 0.0048 1.0000
11.250 1.1900 0.02089 0.01618 0.0044 0.0045 1.0000
11.500 1.2113 0.02181 0.01722 0.0050 0.0044 1.0000
11.750 1.2319 0.02279 0.01834 0.0056 0.0043 1.0000
12.000 1.2515 0.02387 0.01955 0.0063 0.0042 1.0000
12.250 1.2700 0.02505 0.02088 0.0071 0.0041 1.0000
12.500 1.2873 0.02633 0.02231 0.0080 0.0040 1.0000
12.750 1.3030 0.02773 0.02387 0.0089 0.0039 1.0000
13.000 1.3171 0.02926 0.02556 0.0100 0.0038 1.0000
13.250 1.3291 0.03094 0.02742 0.0112 0.0037 1.0000
13.500 1.3384 0.03279 0.02944 0.0126 0.0036 1.0000
13.750 1.3442 0.03487 0.03170 0.0141 0.0035 1.0000
14.000 1.3436 0.03715 0.03416 0.0162 0.0035 1.0000
14.250 1.3356 0.03979 0.03697 0.0186 0.0035 1.0000
14.500 1.3221 0.04330 0.04067 0.0200 0.0034 1.0000
14.750 1.3011 0.04842 0.04600 0.0192 0.0034 1.0000
15.000 1.2597 0.05843 0.05631 0.0131 0.0035 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M2 AIRFOIL (m2-il)