NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M13 AIRFOIL (m13-il) Reynolds number: 500,000 Max Cl/Cd: 94.96 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m13-il-500000-n5.txt Download as CSV file: xf-m13-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M13 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4686 0.09868 0.09658 0.0033 1.0000 0.0064
-8.000 -0.4645 0.09493 0.09286 0.0012 1.0000 0.0064
-7.750 -0.4607 0.09130 0.08925 -0.0009 1.0000 0.0064
-7.500 -0.4562 0.08749 0.08546 -0.0039 1.0000 0.0063
-7.250 -0.4472 0.08309 0.08108 -0.0082 1.0000 0.0063
-7.000 -0.4359 0.07839 0.07638 -0.0129 1.0000 0.0064
-6.750 -0.4225 0.07317 0.07115 -0.0182 1.0000 0.0064
-6.250 -0.3771 0.06220 0.06005 -0.0312 0.9441 0.0069
-6.000 -0.3554 0.05848 0.05616 -0.0347 0.9087 0.0072
-5.750 -0.3349 0.05516 0.05268 -0.0372 0.8818 0.0078
-5.500 -0.3116 0.05028 0.04760 -0.0407 0.8598 0.0087
-5.250 -0.2849 0.04199 0.03897 -0.0455 0.8421 0.0096
-5.000 -0.2593 0.03722 0.03391 -0.0476 0.8245 0.0103
-4.750 -0.2324 0.01972 0.01515 -0.0505 0.8148 0.0147
-4.500 -0.2059 0.02049 0.01588 -0.0504 0.7971 0.0153
-4.250 -0.1793 0.02132 0.01667 -0.0504 0.7804 0.0157
-4.000 -0.1529 0.02124 0.01646 -0.0504 0.7655 0.0167
-3.750 -0.1261 0.01887 0.01372 -0.0503 0.7526 0.0194
-3.500 -0.0986 0.01663 0.01094 -0.0501 0.7406 0.0216
-3.250 -0.0707 0.01665 0.01074 -0.0500 0.7284 0.0225
-3.000 -0.0444 0.01441 0.00815 -0.0502 0.7181 0.0240
-2.750 -0.0170 0.01369 0.00725 -0.0502 0.7078 0.0249
-2.500 0.0105 0.01316 0.00658 -0.0502 0.6974 0.0262
-2.250 0.0383 0.01252 0.00575 -0.0502 0.6879 0.0272
-2.000 0.0660 0.01197 0.00504 -0.0501 0.6790 0.0281
-1.750 0.0938 0.01153 0.00447 -0.0501 0.6694 0.0288
-1.500 0.1216 0.01114 0.00397 -0.0501 0.6606 0.0292
-1.000 0.1771 0.01052 0.00317 -0.0500 0.6434 0.0298
-0.750 0.2049 0.01031 0.00289 -0.0500 0.6347 0.0303
-0.500 0.2328 0.01011 0.00264 -0.0500 0.6258 0.0306
-0.250 0.2606 0.00986 0.00235 -0.0501 0.6177 0.0304
0.000 0.2884 0.00968 0.00212 -0.0501 0.6089 0.0303
0.250 0.3164 0.00952 0.00192 -0.0501 0.5999 0.0302
0.500 0.3443 0.00940 0.00177 -0.0502 0.5914 0.0300
0.750 0.3723 0.00931 0.00165 -0.0502 0.5822 0.0300
1.000 0.4003 0.00924 0.00156 -0.0503 0.5730 0.0300
1.250 0.4283 0.00921 0.00149 -0.0503 0.5637 0.0300
1.500 0.4563 0.00918 0.00145 -0.0504 0.5538 0.0301
1.750 0.4843 0.00918 0.00142 -0.0505 0.5437 0.0303
2.000 0.5122 0.00919 0.00141 -0.0505 0.5336 0.0305
2.250 0.5401 0.00923 0.00141 -0.0506 0.5224 0.0310
2.500 0.5680 0.00926 0.00144 -0.0507 0.5102 0.0332
2.750 0.5958 0.00927 0.00153 -0.0507 0.4981 0.0620
3.000 0.6236 0.00935 0.00163 -0.0508 0.4869 0.0722
3.250 0.6513 0.00942 0.00171 -0.0509 0.4769 0.0773
3.500 0.6790 0.00949 0.00183 -0.0510 0.4672 0.0830
3.750 0.7068 0.00955 0.00195 -0.0511 0.4581 0.0924
4.000 0.7316 0.00781 0.00218 -0.0509 0.4494 1.0000
4.250 0.7587 0.00799 0.00234 -0.0509 0.4320 1.0000
4.500 0.7853 0.00828 0.00251 -0.0508 0.3987 1.0000
4.750 0.8113 0.00866 0.00271 -0.0507 0.3538 1.0000
5.000 0.8352 0.00949 0.00308 -0.0505 0.2588 1.0000
5.250 0.8551 0.01115 0.00395 -0.0502 0.1005 1.0000
5.500 0.8775 0.01229 0.00469 -0.0498 0.0229 1.0000
5.750 0.9030 0.01275 0.00515 -0.0495 0.0152 1.0000
6.000 0.9285 0.01323 0.00573 -0.0492 0.0126 1.0000
6.250 0.9539 0.01367 0.00624 -0.0490 0.0110 1.0000
6.500 0.9787 0.01416 0.00680 -0.0487 0.0094 1.0000
6.750 1.0015 0.01504 0.00777 -0.0481 0.0081 1.0000
7.000 1.0247 0.01578 0.00862 -0.0476 0.0076 1.0000
7.250 1.0473 0.01656 0.00949 -0.0470 0.0071 1.0000
7.500 1.0693 0.01738 0.01040 -0.0463 0.0065 1.0000
7.750 1.0918 0.01805 0.01115 -0.0457 0.0059 1.0000
8.000 1.1136 0.01876 0.01189 -0.0452 0.0054 1.0000
8.250 1.1299 0.02028 0.01351 -0.0438 0.0050 1.0000
8.500 1.1491 0.02135 0.01471 -0.0427 0.0048 1.0000
8.750 1.1669 0.02260 0.01611 -0.0415 0.0046 1.0000
9.000 1.1837 0.02405 0.01771 -0.0401 0.0044 1.0000
9.250 1.1996 0.02568 0.01950 -0.0386 0.0042 1.0000
9.500 1.2148 0.02746 0.02147 -0.0371 0.0040 1.0000
9.750 1.2290 0.02943 0.02367 -0.0356 0.0038 1.0000
10.000 1.2416 0.03161 0.02608 -0.0339 0.0037 1.0000
10.250 1.2522 0.03384 0.02854 -0.0321 0.0036 1.0000
10.500 1.2614 0.03569 0.03058 -0.0304 0.0035 1.0000
10.750 1.2680 0.03725 0.03230 -0.0285 0.0034 1.0000
11.000 1.2695 0.03885 0.03403 -0.0260 0.0033 1.0000
11.250 1.2668 0.04120 0.03656 -0.0238 0.0032 1.0000
11.500 1.2557 0.04487 0.04048 -0.0219 0.0031 1.0000
11.750 1.2462 0.04863 0.04452 -0.0208 0.0031 1.0000
12.000 1.2356 0.05283 0.04899 -0.0204 0.0030 1.0000
12.250 1.2206 0.05787 0.05430 -0.0208 0.0030 1.0000
12.500 1.2039 0.06344 0.06010 -0.0222 0.0030 1.0000
12.750 1.1867 0.06939 0.06624 -0.0243 0.0030 1.0000
13.000 1.1682 0.07600 0.07305 -0.0271 0.0030 1.0000
13.250 1.1497 0.08297 0.08019 -0.0306 0.0030 1.0000
13.500 1.1312 0.09043 0.08780 -0.0347 0.0030 1.0000
13.750 1.1128 0.09837 0.09586 -0.0393 0.0030 1.0000
14.000 1.0950 0.10694 0.10456 -0.0444 0.0030 1.0000
14.250 1.0778 0.11608 0.11379 -0.0499 0.0031 1.0000
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Polar data table (+)
Polar graphs
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