BOEING 707 .19 SPAN AIRFOIL (b707b-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING 707 .19 SPAN AIRFOIL (b707b-il) Reynolds number: 100,000 Max Cl/Cd: 29.99 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707b-il-100000.txt Download as CSV file: xf-b707b-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .19 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7070 0.09325 0.08893 -0.0094 1.0000 0.1816
-10.000 -0.6276 0.09457 0.09016 0.0000 1.0000 0.1868
-9.750 -0.7199 0.08500 0.08075 -0.0108 1.0000 0.1969
-9.500 -0.6277 0.08711 0.08282 -0.0018 1.0000 0.2060
-9.250 -0.6359 0.08282 0.07861 -0.0033 1.0000 0.2184
-9.000 -0.6562 0.07773 0.07363 -0.0060 1.0000 0.2323
-8.750 -0.6618 0.07436 0.07033 -0.0062 1.0000 0.2477
-8.000 -0.8065 0.04928 0.04299 -0.0062 1.0000 0.1219
-7.750 -0.7733 0.04223 0.03520 -0.0071 0.9739 0.0891
-7.500 -0.7434 0.03974 0.03173 -0.0073 0.9552 0.0795
-7.250 -0.7108 0.03589 0.02772 -0.0092 0.9453 0.0774
-7.000 -0.6783 0.03326 0.02480 -0.0105 0.9374 0.0756
-6.750 -0.6481 0.03116 0.02239 -0.0111 0.9295 0.0741
-6.500 -0.6162 0.02932 0.02027 -0.0119 0.9237 0.0731
-6.250 -0.5842 0.02771 0.01846 -0.0126 0.9186 0.0728
-6.000 -0.5520 0.02628 0.01690 -0.0134 0.9131 0.0730
-5.750 -0.5196 0.02500 0.01553 -0.0140 0.9079 0.0737
-5.500 -0.4892 0.02389 0.01438 -0.0144 0.9029 0.0752
-5.250 -0.4614 0.02299 0.01344 -0.0144 0.8975 0.0778
-5.000 -0.4361 0.02202 0.01249 -0.0140 0.8930 0.0811
-4.750 -0.4167 0.02107 0.01166 -0.0128 0.8892 0.0859
-4.500 -0.3980 0.02040 0.01094 -0.0114 0.8852 0.0920
-4.250 -0.3819 0.01957 0.01017 -0.0096 0.8818 0.1020
-4.000 -0.3686 0.01826 0.00924 -0.0074 0.8788 0.1613
-3.750 -0.3786 0.01598 0.00905 -0.0012 0.8760 0.5314
-3.500 -0.3633 0.01606 0.00942 0.0022 0.8739 0.6395
-3.250 -0.3432 0.01635 0.00984 0.0044 0.8718 0.6946
-3.000 -0.3178 0.01660 0.01007 0.0054 0.8701 0.7232
-2.750 -0.2920 0.01696 0.01043 0.0066 0.8679 0.7522
-2.500 -0.2655 0.01728 0.01072 0.0076 0.8645 0.7756
-2.250 -0.2326 0.01779 0.01119 0.0083 0.8604 0.7992
-2.000 -0.1966 0.01846 0.01182 0.0086 0.8561 0.8217
-1.750 -0.1611 0.01904 0.01235 0.0084 0.8518 0.8404
-1.500 -0.1159 0.01978 0.01302 0.0071 0.8469 0.8588
-1.250 -0.0544 0.02040 0.01354 0.0028 0.8412 0.8694
-1.000 -0.0086 0.02065 0.01372 0.0003 0.8330 0.8800
-0.750 0.0196 0.02067 0.01366 0.0008 0.8245 0.8916
-0.500 0.0602 0.02062 0.01355 -0.0009 0.8129 0.8983
-0.250 0.0883 0.02058 0.01348 -0.0010 0.8053 0.9083
0.000 0.1245 0.02048 0.01331 -0.0019 0.7950 0.9158
0.250 0.1520 0.02030 0.01310 -0.0015 0.7809 0.9248
0.500 0.1837 0.02008 0.01282 -0.0015 0.7648 0.9329
0.750 0.2130 0.01981 0.01248 -0.0012 0.7481 0.9411
1.000 0.2450 0.01953 0.01216 -0.0016 0.7319 0.9498
1.250 0.2816 0.01918 0.01180 -0.0029 0.7151 0.9584
1.500 0.3155 0.01883 0.01146 -0.0038 0.6958 0.9688
1.750 0.3555 0.01827 0.01090 -0.0058 0.6728 0.9775
2.000 0.3925 0.01760 0.01021 -0.0073 0.6428 0.9861
2.250 0.4310 0.01686 0.00948 -0.0093 0.6042 0.9949
2.500 0.4597 0.01611 0.00840 -0.0095 0.4476 1.0000
2.750 0.4717 0.01668 0.00831 -0.0074 0.3920 1.0000
3.000 0.4876 0.01703 0.00855 -0.0060 0.3607 1.0000
3.250 0.5040 0.01730 0.00875 -0.0047 0.3298 1.0000
3.500 0.5198 0.01762 0.00893 -0.0032 0.2914 1.0000
3.750 0.5337 0.01819 0.00924 -0.0015 0.2436 1.0000
4.000 0.5495 0.01859 0.00945 0.0002 0.2012 1.0000
4.250 0.5653 0.01897 0.00962 0.0020 0.1836 1.0000
4.500 0.5813 0.01943 0.00992 0.0038 0.1738 1.0000
4.750 0.5973 0.01994 0.01031 0.0055 0.1668 1.0000
5.000 0.6140 0.02052 0.01083 0.0072 0.1613 1.0000
5.250 0.6315 0.02106 0.01138 0.0089 0.1566 1.0000
5.500 0.6493 0.02167 0.01193 0.0104 0.1531 1.0000
5.750 0.6679 0.02248 0.01263 0.0118 0.1505 1.0000
6.000 0.6871 0.02320 0.01339 0.0132 0.1481 1.0000
6.250 0.7059 0.02381 0.01414 0.0147 0.1447 1.0000
6.500 0.7244 0.02443 0.01480 0.0161 0.1408 1.0000
6.750 0.7432 0.02514 0.01545 0.0174 0.1372 1.0000
7.000 0.7633 0.02620 0.01645 0.0184 0.1347 1.0000
7.250 0.7824 0.02698 0.01741 0.0198 0.1333 1.0000
7.500 0.8013 0.02782 0.01845 0.0211 0.1316 1.0000
7.750 0.8197 0.02863 0.01942 0.0225 0.1292 1.0000
8.000 0.8382 0.02952 0.02044 0.0238 0.1271 1.0000
8.250 0.8566 0.03039 0.02139 0.0250 0.1250 1.0000
8.500 0.8760 0.03126 0.02226 0.0261 0.1225 1.0000
8.750 0.8950 0.03256 0.02360 0.0270 0.1207 1.0000
9.000 0.9106 0.03369 0.02503 0.0286 0.1196 1.0000
9.250 0.9253 0.03494 0.02656 0.0302 0.1184 1.0000
9.500 0.9387 0.03605 0.02796 0.0319 0.1161 1.0000
9.750 0.9539 0.03697 0.02903 0.0334 0.1133 1.0000
10.000 0.9705 0.03780 0.02993 0.0346 0.1107 1.0000
10.250 0.9904 0.03897 0.03100 0.0352 0.1076 1.0000
10.500 0.9949 0.04049 0.03297 0.0377 0.1057 1.0000
10.750 1.0000 0.04186 0.03466 0.0400 0.1023 1.0000
11.000 1.0234 0.04161 0.03419 0.0405 0.0964 1.0000
11.250 1.0234 0.04351 0.03638 0.0430 0.0928 1.0000
11.500 1.0224 0.04514 0.03830 0.0456 0.0886 1.0000
11.750 1.0416 0.04668 0.03961 0.0460 0.0822 1.0000
12.000 1.0188 0.04915 0.04256 0.0501 0.0800 1.0000
12.250 0.9946 0.05183 0.04553 0.0538 0.0780 1.0000
12.500 0.9762 0.05481 0.04870 0.0557 0.0756 1.0000
12.750 0.9904 0.05692 0.05062 0.0562 0.0702 1.0000
13.000 0.9600 0.06109 0.05506 0.0568 0.0699 1.0000
13.250 0.9293 0.06625 0.06043 0.0559 0.0699 1.0000
13.500 0.8978 0.07235 0.06672 0.0538 0.0701 1.0000
13.750 0.8633 0.07979 0.07429 0.0502 0.0705 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING 707 .19 SPAN AIRFOIL (b707b-il)