BOEING 707 .99 SPAN AIRFOIL (b707e-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 707 .99 SPAN AIRFOIL (b707e-il) Reynolds number: 100,000 Max Cl/Cd: 53.93 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707e-il-100000.txt Download as CSV file: xf-b707e-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .99 SPAN AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3723 0.10084 0.09610 -0.0264 1.0000 0.0726
-9.500 -0.3778 0.09751 0.09281 -0.0285 1.0000 0.0744
-9.250 -0.4585 0.10113 0.09613 -0.0249 1.0000 0.0675
-9.000 -0.4573 0.09782 0.09287 -0.0261 1.0000 0.0707
-8.750 -0.4601 0.09446 0.08957 -0.0283 1.0000 0.0730
-8.500 -0.4674 0.09117 0.08635 -0.0316 1.0000 0.0743
-8.250 -0.4817 0.08781 0.08309 -0.0354 1.0000 0.0749
-8.000 -0.4957 0.08517 0.08041 -0.0379 1.0000 0.0755
-7.750 -0.5100 0.08350 0.07859 -0.0390 1.0000 0.0760
-7.500 -0.4894 0.07636 0.07171 -0.0364 1.0000 0.0790
-7.250 -0.4856 0.07320 0.06858 -0.0354 1.0000 0.0822
-7.000 -0.4869 0.07011 0.06546 -0.0355 1.0000 0.0855
-6.750 -0.4993 0.06962 0.06454 -0.0362 1.0000 0.0898
-6.500 -0.4933 0.06394 0.05903 -0.0351 1.0000 0.0919
-6.250 -0.4859 0.06062 0.05582 -0.0332 1.0000 0.0961
-6.000 -0.4913 0.06127 0.05591 -0.0314 1.0000 0.1041
-5.750 -0.4862 0.05560 0.05049 -0.0298 1.0000 0.1068
-5.500 -0.4830 0.05312 0.04806 -0.0272 1.0000 0.1123
-5.250 -0.4848 0.05129 0.04595 -0.0249 1.0000 0.1206
-5.000 -0.4798 0.04881 0.04352 -0.0226 1.0000 0.1279
-4.750 -0.4761 0.04646 0.04109 -0.0205 1.0000 0.1403
-4.250 -0.4648 0.04199 0.03642 -0.0170 1.0000 0.1792
-4.000 -0.4580 0.03979 0.03422 -0.0151 1.0000 0.2077
-3.500 -0.4418 0.03557 0.03002 -0.0112 1.0000 0.2771
-3.250 -0.4307 0.03357 0.02807 -0.0091 1.0000 0.3113
-3.000 -0.4063 0.03169 0.02599 -0.0095 0.9972 0.3436
-2.750 -0.3701 0.02920 0.02328 -0.0117 0.9924 0.3574
-2.500 -0.3057 0.02788 0.02052 -0.0168 0.9869 0.2461
-2.250 -0.2419 0.02734 0.01864 -0.0177 0.9828 0.1168
-2.000 -0.2072 0.02609 0.01703 -0.0181 0.9773 0.1024
-1.750 -0.1679 0.02547 0.01598 -0.0193 0.9720 0.0901
-1.500 -0.1305 0.02356 0.01410 -0.0211 0.9683 0.0843
-1.250 -0.0993 0.02277 0.01314 -0.0214 0.9622 0.0790
-1.000 -0.0614 0.02247 0.01266 -0.0231 0.9572 0.0762
-0.750 -0.0311 0.02164 0.01192 -0.0238 0.9517 0.0754
-0.500 -0.0012 0.02106 0.01142 -0.0244 0.9456 0.0753
-0.250 0.0348 0.02066 0.01097 -0.0262 0.9410 0.0764
0.000 0.0580 0.02035 0.01059 -0.0257 0.9333 0.0798
0.250 0.0966 0.02019 0.01033 -0.0279 0.9283 0.0890
0.500 0.1194 0.01844 0.01037 -0.0279 0.9219 0.5974
0.750 0.2148 0.01765 0.01035 -0.0402 0.9228 1.0000
1.000 0.2535 0.01783 0.01040 -0.0427 0.9158 1.0000
1.250 0.2844 0.01803 0.01052 -0.0436 0.9067 1.0000
1.500 0.3417 0.01789 0.01035 -0.0492 0.9002 1.0000
1.750 0.3821 0.01778 0.01023 -0.0515 0.8888 1.0000
2.000 0.4277 0.01749 0.00997 -0.0544 0.8773 1.0000
2.250 0.5022 0.01584 0.00839 -0.0603 0.8559 1.0000
2.500 0.5412 0.01494 0.00747 -0.0599 0.8308 1.0000
2.750 0.5651 0.01475 0.00727 -0.0580 0.8104 1.0000
3.000 0.5890 0.01463 0.00719 -0.0562 0.7919 1.0000
3.250 0.6122 0.01463 0.00723 -0.0547 0.7766 1.0000
3.500 0.6351 0.01463 0.00728 -0.0531 0.7612 1.0000
3.750 0.6581 0.01460 0.00732 -0.0516 0.7453 1.0000
4.000 0.6775 0.01458 0.00743 -0.0493 0.7245 1.0000
4.250 0.6984 0.01439 0.00729 -0.0471 0.6985 1.0000
4.500 0.7203 0.01430 0.00727 -0.0452 0.6742 1.0000
4.750 0.7411 0.01427 0.00732 -0.0432 0.6443 1.0000
5.000 0.7610 0.01427 0.00733 -0.0409 0.6028 1.0000
5.250 0.7777 0.01442 0.00738 -0.0381 0.5279 1.0000
5.500 0.7906 0.01504 0.00745 -0.0347 0.4250 1.0000
5.750 0.8032 0.01597 0.00798 -0.0319 0.3475 1.0000
6.000 0.8147 0.01708 0.00868 -0.0292 0.2729 1.0000
6.250 0.8240 0.01850 0.00951 -0.0264 0.1723 1.0000
6.500 0.8358 0.01993 0.01043 -0.0239 0.1111 1.0000
6.750 0.8499 0.02119 0.01151 -0.0217 0.0916 1.0000
7.000 0.8624 0.02268 0.01301 -0.0192 0.0774 1.0000
7.250 0.8762 0.02423 0.01458 -0.0169 0.0620 1.0000
7.500 0.8898 0.02634 0.01657 -0.0147 0.0527 1.0000
7.750 0.9127 0.02840 0.01876 -0.0134 0.0485 1.0000
8.000 0.9389 0.03089 0.02134 -0.0128 0.0457 1.0000
8.250 0.9655 0.03390 0.02451 -0.0124 0.0441 1.0000
8.500 0.9885 0.03784 0.02867 -0.0119 0.0428 1.0000
8.750 1.0006 0.04067 0.03206 -0.0095 0.0414 1.0000
9.000 1.0120 0.04382 0.03566 -0.0073 0.0412 1.0000
9.250 1.0222 0.04845 0.04061 -0.0056 0.0421 1.0000
9.500 1.0311 0.05033 0.04306 -0.0024 0.0443 1.0000
9.750 1.0165 0.05514 0.04868 0.0021 0.0483 1.0000
10.000 1.0117 0.05985 0.05367 0.0046 0.0509 1.0000
10.250 1.0018 0.06402 0.05824 0.0076 0.0555 1.0000
10.500 0.9659 0.06827 0.06285 0.0116 0.0579 1.0000
10.750 0.9392 0.07261 0.06738 0.0130 0.0592 1.0000
11.000 0.9139 0.07755 0.07246 0.0125 0.0602 1.0000
11.250 0.8898 0.08314 0.07816 0.0104 0.0609 1.0000
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Polar data table (+)
Polar graphs
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