NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Reynolds number: 200,000 Max Cl/Cd: 41.82 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-a63a108c-il-200000-n5.txt Download as CSV file: xf-a63a108c-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES 63A108 MOD C 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7432 0.08327 0.07976 0.0147 1.0000 0.0158
-8.500 -0.7591 0.07357 0.07001 0.0046 1.0000 0.0155
-8.250 -0.7718 0.06426 0.06051 -0.0013 1.0000 0.0152
-8.000 -0.7811 0.05490 0.05079 -0.0046 1.0000 0.0150
-7.750 -0.7841 0.04606 0.04141 -0.0060 1.0000 0.0150
-7.500 -0.7784 0.03891 0.03362 -0.0062 1.0000 0.0152
-7.250 -0.7648 0.03364 0.02769 -0.0059 1.0000 0.0158
-7.000 -0.7470 0.02910 0.02234 -0.0054 1.0000 0.0168
-6.750 -0.7228 0.02792 0.02107 -0.0054 1.0000 0.0174
-6.500 -0.6988 0.02589 0.01871 -0.0052 1.0000 0.0179
-6.250 -0.6739 0.02374 0.01619 -0.0049 1.0000 0.0185
-6.000 -0.6481 0.02182 0.01389 -0.0046 1.0000 0.0192
-5.750 -0.6219 0.02048 0.01239 -0.0044 1.0000 0.0198
-5.500 -0.5952 0.01949 0.01134 -0.0043 1.0000 0.0206
-5.250 -0.5681 0.01855 0.01023 -0.0042 1.0000 0.0222
-5.000 -0.5411 0.01775 0.00941 -0.0042 1.0000 0.0235
-4.750 -0.5139 0.01691 0.00851 -0.0042 1.0000 0.0249
-4.500 -0.4867 0.01606 0.00762 -0.0041 1.0000 0.0262
-4.250 -0.4592 0.01538 0.00694 -0.0041 1.0000 0.0278
-4.000 -0.4315 0.01478 0.00633 -0.0042 1.0000 0.0303
-3.750 -0.4036 0.01422 0.00575 -0.0043 1.0000 0.0334
-3.500 -0.3756 0.01368 0.00521 -0.0044 1.0000 0.0367
-3.250 -0.3474 0.01320 0.00475 -0.0046 1.0000 0.0417
-3.000 -0.3191 0.01274 0.00434 -0.0048 1.0000 0.0485
-2.750 -0.2909 0.01229 0.00395 -0.0050 1.0000 0.0599
-2.500 -0.2628 0.01181 0.00361 -0.0052 1.0000 0.0828
-2.250 -0.2352 0.01124 0.00332 -0.0055 1.0000 0.1374
-2.000 -0.2086 0.01045 0.00305 -0.0058 1.0000 0.2534
-1.750 -0.1841 0.00945 0.00288 -0.0058 1.0000 0.4484
-1.250 -0.1257 0.00857 0.00296 -0.0062 0.9826 0.6908
-1.000 -0.0941 0.00839 0.00299 -0.0066 0.9712 0.7592
-0.750 -0.0653 0.00825 0.00303 -0.0061 0.9575 0.8151
-0.500 -0.0414 0.00816 0.00310 -0.0041 0.9382 0.8756
-0.250 -0.0145 0.00813 0.00311 -0.0027 0.9161 0.9320
0.000 0.0184 0.00813 0.00306 -0.0031 0.8947 0.9660
0.250 0.0528 0.00813 0.00299 -0.0041 0.8753 0.9941
0.500 0.0793 0.00812 0.00290 -0.0035 0.8533 1.0000
0.750 0.1038 0.00814 0.00283 -0.0025 0.8288 1.0000
1.000 0.1281 0.00817 0.00275 -0.0014 0.8023 1.0000
1.250 0.1533 0.00822 0.00269 -0.0005 0.7682 1.0000
1.500 0.1784 0.00832 0.00263 0.0006 0.7241 1.0000
1.750 0.2037 0.00851 0.00257 0.0015 0.6645 1.0000
2.000 0.2292 0.00886 0.00253 0.0022 0.5780 1.0000
2.250 0.2555 0.00939 0.00259 0.0025 0.4728 1.0000
2.500 0.2826 0.00999 0.00272 0.0024 0.3727 1.0000
3.000 0.3381 0.01098 0.00309 0.0020 0.2426 1.0000
3.250 0.3660 0.01139 0.00332 0.0018 0.2014 1.0000
3.500 0.3939 0.01178 0.00355 0.0016 0.1689 1.0000
4.000 0.4498 0.01254 0.00408 0.0013 0.1227 1.0000
4.250 0.4776 0.01294 0.00440 0.0012 0.1061 1.0000
4.500 0.5054 0.01333 0.00474 0.0011 0.0928 1.0000
4.750 0.5331 0.01374 0.00510 0.0010 0.0817 1.0000
5.000 0.5607 0.01417 0.00551 0.0009 0.0728 1.0000
5.250 0.5881 0.01462 0.00594 0.0009 0.0653 1.0000
5.500 0.6155 0.01510 0.00640 0.0008 0.0589 1.0000
5.750 0.6427 0.01560 0.00691 0.0008 0.0536 1.0000
6.000 0.6697 0.01613 0.00746 0.0008 0.0489 1.0000
6.250 0.6965 0.01670 0.00805 0.0008 0.0451 1.0000
6.500 0.7231 0.01729 0.00864 0.0009 0.0418 1.0000
6.750 0.7496 0.01793 0.00935 0.0009 0.0389 1.0000
7.000 0.7757 0.01864 0.01011 0.0010 0.0365 1.0000
7.250 0.8016 0.01931 0.01081 0.0011 0.0343 1.0000
7.500 0.8273 0.02010 0.01172 0.0013 0.0323 1.0000
7.750 0.8526 0.02090 0.01257 0.0014 0.0308 1.0000
8.000 0.8773 0.02189 0.01365 0.0017 0.0295 1.0000
8.250 0.9019 0.02288 0.01479 0.0019 0.0282 1.0000
8.500 0.9262 0.02376 0.01570 0.0021 0.0271 1.0000
8.750 0.9498 0.02495 0.01710 0.0024 0.0258 1.0000
9.000 0.9729 0.02613 0.01846 0.0027 0.0248 1.0000
9.250 0.9955 0.02730 0.01974 0.0030 0.0241 1.0000
9.500 1.0171 0.02862 0.02115 0.0034 0.0236 1.0000
9.750 1.0363 0.03069 0.02363 0.0040 0.0229 1.0000
10.000 1.0540 0.03281 0.02608 0.0045 0.0222 1.0000
10.250 1.0709 0.03475 0.02828 0.0051 0.0216 1.0000
10.500 1.0879 0.03633 0.03003 0.0056 0.0211 1.0000
10.750 1.1044 0.03775 0.03152 0.0061 0.0207 1.0000
11.000 1.1062 0.04175 0.03612 0.0071 0.0201 1.0000
11.250 1.1010 0.04612 0.04099 0.0079 0.0197 1.0000
11.500 1.0860 0.05081 0.04609 0.0087 0.0195 1.0000
11.750 1.0599 0.05628 0.05187 0.0079 0.0195 1.0000
12.000 1.0249 0.06569 0.06160 0.0007 0.0197 1.0000
12.250 0.9329 0.09620 0.09241 -0.0237 0.0207 1.0000
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Polar data table (+)
Polar graphs
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