NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M13 AIRFOIL (m13-il) Reynolds number: 1,000,000 Max Cl/Cd: 86.34 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m13-il-1000000-n5.txt Download as CSV file: xf-m13-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4800 0.09682 0.09534 0.0043 1.0000 0.0033
-8.000 -0.4766 0.09298 0.09152 0.0021 1.0000 0.0034
-7.750 -0.4740 0.08910 0.08767 -0.0004 1.0000 0.0035
-7.500 -0.4651 0.08416 0.08272 -0.0058 0.9725 0.0038
-7.250 -0.4557 0.07844 0.07684 -0.0116 0.9035 0.0040
-7.000 -0.4454 0.07170 0.06996 -0.0181 0.8736 0.0042
-6.750 -0.4305 0.06394 0.06205 -0.0258 0.8504 0.0045
-6.500 -0.4107 0.05981 0.05777 -0.0297 0.8264 0.0048
-6.250 -0.3892 0.05517 0.05298 -0.0338 0.8049 0.0051
-6.000 -0.3655 0.04787 0.04546 -0.0394 0.7876 0.0057
-5.750 -0.3402 0.03496 0.03201 -0.0461 0.7756 0.0067
-5.500 -0.3149 0.03140 0.02823 -0.0472 0.7604 0.0076
-5.250 -0.2966 0.01498 0.01040 -0.0487 0.7527 0.0095
-4.750 -0.2423 0.01376 0.00887 -0.0487 0.7262 0.0128
-4.500 -0.2160 0.01185 0.00649 -0.0487 0.7152 0.0151
-4.250 -0.1878 0.01225 0.00689 -0.0487 0.7043 0.0157
-4.000 -0.1598 0.01248 0.00711 -0.0488 0.6939 0.0166
-3.750 -0.1319 0.01235 0.00689 -0.0489 0.6840 0.0177
-3.500 -0.1040 0.01210 0.00652 -0.0489 0.6748 0.0190
-3.250 -0.0762 0.01181 0.00610 -0.0489 0.6660 0.0202
-3.000 -0.0481 0.01162 0.00582 -0.0490 0.6575 0.0211
-2.750 -0.0201 0.01146 0.00556 -0.0490 0.6489 0.0217
-2.500 0.0080 0.01140 0.00542 -0.0491 0.6401 0.0222
-2.250 0.0362 0.01147 0.00541 -0.0492 0.6326 0.0225
-2.000 0.0633 0.01005 0.00382 -0.0492 0.6247 0.0241
-1.750 0.0912 0.00970 0.00341 -0.0493 0.6167 0.0248
-1.250 0.1471 0.00917 0.00277 -0.0494 0.6007 0.0256
-1.000 0.1751 0.00895 0.00249 -0.0495 0.5929 0.0259
-0.750 0.2031 0.00874 0.00224 -0.0495 0.5843 0.0260
-0.500 0.2312 0.00857 0.00202 -0.0496 0.5760 0.0262
-0.250 0.2593 0.00843 0.00184 -0.0497 0.5675 0.0266
0.000 0.2874 0.00830 0.00167 -0.0498 0.5590 0.0266
0.250 0.3155 0.00819 0.00151 -0.0498 0.5497 0.0263
0.500 0.3437 0.00810 0.00139 -0.0499 0.5401 0.0261
0.750 0.3719 0.00803 0.00129 -0.0500 0.5310 0.0259
1.000 0.4000 0.00799 0.00122 -0.0501 0.5206 0.0258
1.250 0.4282 0.00796 0.00116 -0.0503 0.5100 0.0257
1.500 0.4563 0.00795 0.00112 -0.0504 0.4989 0.0256
1.750 0.4844 0.00797 0.00111 -0.0505 0.4866 0.0256
2.000 0.5124 0.00800 0.00110 -0.0506 0.4736 0.0256
2.250 0.5405 0.00804 0.00111 -0.0507 0.4622 0.0258
2.500 0.5685 0.00807 0.00113 -0.0508 0.4527 0.0262
2.750 0.5966 0.00809 0.00116 -0.0509 0.4448 0.0272
3.000 0.6246 0.00814 0.00120 -0.0511 0.4360 0.0301
3.250 0.6527 0.00816 0.00129 -0.0512 0.4281 0.0574
3.750 0.7085 0.00830 0.00150 -0.0515 0.4103 0.0711
4.000 0.7355 0.00857 0.00163 -0.0515 0.3708 0.0739
4.250 0.7624 0.00883 0.00180 -0.0516 0.3340 0.0785
4.500 0.7877 0.00949 0.00210 -0.0516 0.2510 0.0836
4.750 0.8114 0.01046 0.00260 -0.0515 0.1445 0.0904
5.000 0.8343 0.01161 0.00327 -0.0512 0.0338 0.0965
5.500 0.8832 0.01054 0.00400 -0.0507 0.0126 1.0000
5.750 0.9097 0.01083 0.00431 -0.0506 0.0105 1.0000
6.000 0.9356 0.01124 0.00473 -0.0505 0.0080 1.0000
6.250 0.9619 0.01154 0.00508 -0.0504 0.0074 1.0000
6.500 0.9879 0.01189 0.00546 -0.0502 0.0066 1.0000
6.750 1.0135 0.01228 0.00587 -0.0501 0.0059 1.0000
7.000 1.0381 0.01286 0.00649 -0.0498 0.0051 1.0000
7.250 1.0635 0.01324 0.00691 -0.0496 0.0047 1.0000
7.500 1.0882 0.01371 0.00746 -0.0493 0.0043 1.0000
7.750 1.1125 0.01423 0.00803 -0.0490 0.0040 1.0000
8.000 1.1365 0.01476 0.00860 -0.0486 0.0037 1.0000
8.250 1.1600 0.01532 0.00922 -0.0482 0.0035 1.0000
8.500 1.1819 0.01610 0.01006 -0.0476 0.0033 1.0000
8.750 1.2019 0.01712 0.01119 -0.0467 0.0031 1.0000
9.000 1.2232 0.01789 0.01205 -0.0460 0.0029 1.0000
9.250 1.2432 0.01879 0.01305 -0.0451 0.0028 1.0000
9.500 1.2621 0.01979 0.01419 -0.0441 0.0027 1.0000
9.750 1.2807 0.02076 0.01528 -0.0431 0.0025 1.0000
10.000 1.2990 0.02171 0.01632 -0.0420 0.0024 1.0000
10.250 1.3173 0.02259 0.01729 -0.0411 0.0023 1.0000
10.500 1.3348 0.02349 0.01827 -0.0400 0.0022 1.0000
10.750 1.3509 0.02449 0.01937 -0.0388 0.0021 1.0000
11.000 1.3652 0.02561 0.02059 -0.0375 0.0020 1.0000
11.250 1.3769 0.02688 0.02198 -0.0358 0.0020 1.0000
11.500 1.3842 0.02826 0.02350 -0.0335 0.0019 1.0000
11.750 1.3860 0.02998 0.02537 -0.0307 0.0019 1.0000
12.000 1.3851 0.03217 0.02777 -0.0283 0.0018 1.0000
12.250 1.3785 0.03523 0.03107 -0.0262 0.0018 1.0000
12.500 1.3704 0.03872 0.03480 -0.0246 0.0017 1.0000
12.750 1.3658 0.04195 0.03824 -0.0239 0.0017 1.0000
13.000 1.3617 0.04528 0.04176 -0.0237 0.0017 1.0000
13.250 1.3531 0.04939 0.04607 -0.0238 0.0017 1.0000
13.500 1.3418 0.05405 0.05094 -0.0245 0.0017 1.0000
13.750 1.3273 0.05940 0.05650 -0.0258 0.0017 1.0000
14.000 1.3112 0.06521 0.06250 -0.0276 0.0017 1.0000
14.250 1.2920 0.07186 0.06933 -0.0302 0.0017 1.0000
14.500 1.2717 0.07898 0.07663 -0.0332 0.0017 1.0000
14.750 1.2497 0.08698 0.08480 -0.0370 0.0017 1.0000
15.000 1.2261 0.09587 0.09385 -0.0416 0.0017 1.0000
15.250 1.2026 0.10537 0.10350 -0.0467 0.0017 1.0000
15.500 1.1792 0.11551 0.11377 -0.0522 0.0017 1.0000
15.750 1.1512 0.12752 0.12590 -0.0587 0.0017 1.0000
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Polar data table (+)
Polar graphs
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