NACA M26 AIRFOIL (m26-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M26 AIRFOIL (m26-il) Reynolds number: 1,000,000 Max Cl/Cd: 141.1 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m26-il-1000000.txt Download as CSV file: xf-m26-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M26 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3003 0.10594 0.10290 0.0128 0.5196 0.0099
-7.750 -0.2935 0.10293 0.09990 0.0104 0.5184 0.0099
-7.500 -0.2882 0.10003 0.09701 0.0080 0.5172 0.0099
-7.250 -0.2772 0.09673 0.09370 0.0050 0.5161 0.0099
-7.000 -0.2680 0.09260 0.08956 0.0028 0.5149 0.0100
-6.750 -0.2563 0.08949 0.08644 0.0010 0.5137 0.0102
-6.500 -0.2420 0.08665 0.08357 -0.0014 0.5125 0.0104
-6.250 -0.2259 0.08383 0.08072 -0.0040 0.5111 0.0107
-6.000 -0.2084 0.08093 0.07778 -0.0068 0.5095 0.0112
-5.750 -0.1871 0.07760 0.07441 -0.0104 0.5084 0.0126
-5.500 -0.1611 0.07416 0.07093 -0.0151 0.5078 0.0128
-5.250 -0.1365 0.07073 0.06746 -0.0189 0.5070 0.0129
-5.000 -0.1115 0.06735 0.06402 -0.0223 0.5061 0.0129
-4.750 -0.0983 0.06355 0.06022 -0.0230 0.5051 0.0132
-4.500 -0.0764 0.06110 0.05772 -0.0249 0.5039 0.0136
-4.250 -0.0522 0.05863 0.05521 -0.0270 0.5027 0.0142
-4.000 -0.0234 0.05593 0.05244 -0.0295 0.5015 0.0160
-3.750 0.0128 0.05311 0.04949 -0.0330 0.5004 0.0163
-3.500 0.0439 0.05018 0.04646 -0.0354 0.4994 0.0164
-3.250 0.0724 0.02814 0.02463 -0.0363 0.4967 0.0168
-3.000 0.0943 0.02612 0.02256 -0.0371 0.4954 0.0172
-2.750 0.1186 0.02423 0.02059 -0.0382 0.4937 0.0179
-2.500 0.1451 0.02221 0.01850 -0.0392 0.4931 0.0194
-2.250 0.1816 0.02059 0.01673 -0.0408 0.4923 0.0204
-2.000 0.2096 0.01852 0.01455 -0.0414 0.4914 0.0205
-1.750 0.2368 0.01657 0.01249 -0.0419 0.4905 0.0205
-1.500 0.2585 0.01372 0.00954 -0.0424 0.4896 0.0210
-1.250 0.2826 0.01252 0.00829 -0.0427 0.4885 0.0214
-1.000 0.3081 0.01151 0.00722 -0.0431 0.4873 0.0219
-0.750 0.3349 0.01054 0.00616 -0.0434 0.4861 0.0229
-0.500 0.3674 0.00994 0.00541 -0.0434 0.4848 0.0255
-0.250 0.3967 0.00913 0.00447 -0.0435 0.4833 0.0258
0.000 0.4246 0.00829 0.00349 -0.0435 0.4815 0.0258
0.250 0.4525 0.00754 0.00259 -0.0435 0.4793 0.0259
0.750 0.5205 0.01806 0.01254 -0.0446 0.4783 0.0269
1.000 0.5483 0.01731 0.01174 -0.0449 0.4773 0.0276
1.250 0.5766 0.01659 0.01096 -0.0451 0.4762 0.0287
1.500 0.6060 0.01583 0.01010 -0.0451 0.4749 0.0319
1.750 0.6357 0.01545 0.00962 -0.0450 0.4735 0.0331
2.000 0.6637 0.01330 0.00720 -0.0448 0.4723 0.0347
2.250 0.6919 0.01288 0.00676 -0.0451 0.4709 0.0358
2.500 0.7202 0.01256 0.00641 -0.0454 0.4694 0.0375
2.750 0.7487 0.01217 0.00595 -0.0456 0.4679 0.0399
3.000 0.7775 0.01212 0.00579 -0.0457 0.4663 0.0436
3.250 0.8051 0.01148 0.00512 -0.0460 0.4643 0.0475
3.500 0.8336 0.01123 0.00488 -0.0462 0.4633 0.0509
3.750 0.8619 0.01040 0.00391 -0.0458 0.4620 0.0333
4.000 0.8893 0.01010 0.00362 -0.0458 0.4605 0.0335
4.250 0.9168 0.00992 0.00344 -0.0459 0.4589 0.0337
4.500 0.9444 0.00975 0.00328 -0.0461 0.4572 0.0347
4.750 0.9725 0.00962 0.00316 -0.0463 0.4541 0.0352
5.000 1.0004 0.00960 0.00310 -0.0466 0.4494 0.0358
5.250 1.0290 0.00946 0.00303 -0.0469 0.4451 0.0375
5.500 1.0574 0.00938 0.00294 -0.0473 0.4382 0.0416
5.750 1.0855 0.00936 0.00301 -0.0476 0.4333 0.0911
6.000 1.1074 0.00839 0.00316 -0.0471 0.4267 0.7911
6.250 1.1500 0.00815 0.00339 -0.0504 0.4077 0.9915
6.500 1.2112 0.01015 0.00463 -0.0602 0.2804 0.9980
6.750 1.2340 0.01523 0.00835 -0.0652 0.0203 1.0000
7.000 1.2551 0.01573 0.00887 -0.0646 0.0172 1.0000
7.250 1.2749 0.01632 0.00952 -0.0639 0.0149 1.0000
7.500 1.2941 0.01685 0.01009 -0.0632 0.0143 1.0000
7.750 1.3117 0.01747 0.01075 -0.0622 0.0136 1.0000
8.000 1.3268 0.01821 0.01154 -0.0610 0.0129 1.0000
8.250 1.3388 0.01908 0.01247 -0.0594 0.0123 1.0000
8.500 1.3454 0.02028 0.01374 -0.0577 0.0116 1.0000
8.750 1.3359 0.02220 0.01574 -0.0544 0.0112 1.0000
9.000 1.3268 0.02464 0.01827 -0.0519 0.0108 1.0000
9.250 1.3155 0.02817 0.02195 -0.0506 0.0104 1.0000
9.500 1.3180 0.03050 0.02435 -0.0501 0.0103 1.0000
9.750 1.3220 0.03273 0.02663 -0.0496 0.0101 1.0000
10.000 1.3235 0.03522 0.02919 -0.0491 0.0099 1.0000
10.250 1.3230 0.03793 0.03197 -0.0486 0.0097 1.0000
10.500 1.3218 0.04073 0.03483 -0.0480 0.0095 1.0000
10.750 1.3192 0.04364 0.03783 -0.0475 0.0093 1.0000
11.000 1.3160 0.04665 0.04091 -0.0469 0.0091 1.0000
11.250 1.3131 0.04966 0.04398 -0.0464 0.0089 1.0000
11.500 1.3106 0.05274 0.04713 -0.0461 0.0087 1.0000
11.750 1.3090 0.05574 0.05019 -0.0457 0.0085 1.0000
12.000 1.3077 0.05877 0.05327 -0.0455 0.0083 1.0000
12.250 1.3081 0.06161 0.05615 -0.0452 0.0080 1.0000
12.500 1.3071 0.06464 0.05923 -0.0451 0.0078 1.0000
12.750 1.3047 0.06780 0.06242 -0.0448 0.0075 1.0000
13.000 1.2958 0.07158 0.06624 -0.0442 0.0073 1.0000
13.250 1.2935 0.07398 0.06867 -0.0428 0.0071 1.0000
13.500 1.2996 0.07628 0.07102 -0.0428 0.0070 1.0000
13.750 1.3062 0.07838 0.07317 -0.0425 0.0069 1.0000
14.000 1.3134 0.08024 0.07508 -0.0420 0.0067 1.0000
14.250 1.3213 0.08188 0.07676 -0.0414 0.0066 1.0000
14.500 1.3304 0.08328 0.07820 -0.0405 0.0065 1.0000
14.750 1.3409 0.08443 0.07938 -0.0395 0.0063 1.0000
15.000 1.3532 0.08522 0.08023 -0.0383 0.0061 1.0000
15.250 1.3661 0.08590 0.08096 -0.0368 0.0059 1.0000
15.500 1.3807 0.08622 0.08134 -0.0349 0.0057 1.0000
15.750 1.3954 0.08654 0.08174 -0.0328 0.0057 1.0000
16.000 1.4030 0.08834 0.08360 -0.0321 0.0055 1.0000
16.250 1.4065 0.09091 0.08623 -0.0325 0.0053 1.0000
16.500 1.4086 0.09379 0.08914 -0.0334 0.0052 1.0000
16.750 1.4105 0.09662 0.09202 -0.0339 0.0050 1.0000
17.000 1.4174 0.09824 0.09369 -0.0329 0.0049 1.0000
17.250 1.4236 0.10021 0.09580 -0.0317 0.0049 1.0000
19.000 1.3368 0.13995 0.13713 -0.0389 0.0063 1.0000
19.250 1.3240 0.14633 0.14366 -0.0422 0.0062 1.0000
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