BOEING 707 .19 SPAN AIRFOIL (b707b-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING 707 .19 SPAN AIRFOIL (b707b-il) Reynolds number: 50,000 Max Cl/Cd: 22.61 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b707b-il-50000-n5.txt Download as CSV file: xf-b707b-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6945 0.10355 0.09665 -0.0111 1.0000 0.0625
-11.750 -0.6962 0.09859 0.09169 -0.0131 1.0000 0.0620
-11.500 -0.7058 0.09295 0.08603 -0.0160 1.0000 0.0614
-11.250 -0.7208 0.08730 0.08034 -0.0189 1.0000 0.0609
-11.000 -0.7391 0.08203 0.07500 -0.0211 1.0000 0.0603
-10.750 -0.7586 0.07739 0.07026 -0.0222 1.0000 0.0599
-10.500 -0.7767 0.07339 0.06614 -0.0220 1.0000 0.0595
-10.250 -0.7930 0.06986 0.06247 -0.0206 1.0000 0.0591
-10.000 -0.8061 0.06661 0.05904 -0.0186 1.0000 0.0588
-9.750 -0.8135 0.06332 0.05553 -0.0169 1.0000 0.0584
-9.500 -0.8169 0.06011 0.05206 -0.0151 1.0000 0.0580
-9.250 -0.8161 0.05703 0.04870 -0.0134 1.0000 0.0577
-9.000 -0.8113 0.05407 0.04545 -0.0118 1.0000 0.0574
-8.750 -0.8031 0.05126 0.04232 -0.0103 1.0000 0.0571
-8.500 -0.7914 0.04860 0.03936 -0.0091 1.0000 0.0569
-8.250 -0.7768 0.04610 0.03656 -0.0080 1.0000 0.0567
-8.000 -0.7591 0.04376 0.03393 -0.0073 1.0000 0.0566
-7.750 -0.7391 0.04159 0.03151 -0.0067 1.0000 0.0566
-7.500 -0.7170 0.03960 0.02927 -0.0063 1.0000 0.0568
-7.250 -0.6934 0.03779 0.02727 -0.0061 1.0000 0.0570
-7.000 -0.6637 0.03605 0.02535 -0.0069 0.9787 0.0573
-6.750 -0.6243 0.03432 0.02342 -0.0091 0.9538 0.0580
-6.500 -0.5834 0.03282 0.02175 -0.0113 0.9394 0.0590
-6.250 -0.5440 0.03155 0.02032 -0.0132 0.9283 0.0602
-6.000 -0.5089 0.03050 0.01906 -0.0145 0.9184 0.0618
-5.750 -0.4800 0.02944 0.01794 -0.0151 0.9093 0.0643
-5.500 -0.4556 0.02851 0.01692 -0.0151 0.9011 0.0677
-5.250 -0.4334 0.02768 0.01592 -0.0147 0.8941 0.0719
-5.000 -0.4136 0.02674 0.01487 -0.0140 0.8880 0.0775
-4.750 -0.3943 0.02585 0.01388 -0.0131 0.8828 0.0859
-4.500 -0.3769 0.02482 0.01295 -0.0121 0.8778 0.1002
-4.250 -0.3604 0.02358 0.01204 -0.0111 0.8738 0.1534
-4.000 -0.3624 0.02134 0.01165 -0.0071 0.8696 0.4541
-3.750 -0.3481 0.02132 0.01232 -0.0033 0.8663 0.6074
-3.500 -0.3264 0.02154 0.01254 -0.0013 0.8631 0.6590
-3.250 -0.2993 0.02189 0.01287 0.0001 0.8601 0.6989
-3.000 -0.2700 0.02234 0.01323 0.0012 0.8570 0.7334
-2.750 -0.2405 0.02282 0.01364 0.0022 0.8534 0.7632
-2.500 -0.2066 0.02349 0.01423 0.0029 0.8498 0.7930
-2.250 -0.1652 0.02400 0.01460 0.0018 0.8462 0.8125
-2.000 -0.1328 0.02409 0.01455 0.0012 0.8422 0.8255
-1.750 -0.1050 0.02409 0.01444 0.0011 0.8380 0.8373
-1.500 -0.0712 0.02412 0.01436 -0.0001 0.8328 0.8461
-1.250 -0.0395 0.02409 0.01421 -0.0006 0.8261 0.8553
-1.000 -0.0125 0.02406 0.01411 -0.0005 0.8177 0.8659
-0.750 0.0213 0.02400 0.01395 -0.0012 0.8088 0.8745
-0.500 0.0537 0.02399 0.01389 -0.0021 0.8000 0.8834
-0.250 0.0844 0.02395 0.01381 -0.0026 0.7935 0.8931
0.000 0.1145 0.02396 0.01381 -0.0033 0.7876 0.9036
0.250 0.1505 0.02395 0.01380 -0.0048 0.7796 0.9134
0.500 0.1852 0.02387 0.01371 -0.0057 0.7675 0.9242
0.750 0.2193 0.02374 0.01357 -0.0064 0.7530 0.9355
1.000 0.2527 0.02355 0.01337 -0.0070 0.7369 0.9464
1.250 0.2864 0.02335 0.01319 -0.0078 0.7216 0.9568
1.500 0.3198 0.02311 0.01299 -0.0086 0.7045 0.9668
1.750 0.3516 0.02275 0.01264 -0.0091 0.6805 0.9765
2.000 0.3828 0.02231 0.01222 -0.0094 0.6511 0.9857
2.250 0.4152 0.02188 0.01186 -0.0103 0.6178 0.9946
2.500 0.4427 0.02164 0.01183 -0.0111 0.5735 1.0000
2.750 0.4537 0.02115 0.01093 -0.0073 0.4403 1.0000
3.000 0.4640 0.02145 0.01053 -0.0041 0.3975 1.0000
3.250 0.4778 0.02198 0.01086 -0.0023 0.3656 1.0000
3.500 0.4932 0.02247 0.01127 -0.0008 0.3377 1.0000
4.000 0.5212 0.02352 0.01214 0.0027 0.2837 1.0000
4.250 0.5334 0.02415 0.01259 0.0047 0.2602 1.0000
4.500 0.5469 0.02467 0.01312 0.0066 0.2309 1.0000
4.750 0.5610 0.02514 0.01361 0.0086 0.2006 1.0000
5.000 0.5736 0.02575 0.01404 0.0107 0.1816 1.0000
5.250 0.5861 0.02641 0.01452 0.0128 0.1705 1.0000
5.500 0.6000 0.02708 0.01513 0.0149 0.1618 1.0000
5.750 0.6144 0.02778 0.01575 0.0169 0.1558 1.0000
6.000 0.6305 0.02849 0.01642 0.0188 0.1515 1.0000
6.250 0.6491 0.02918 0.01720 0.0204 0.1474 1.0000
6.500 0.6687 0.02991 0.01799 0.0218 0.1441 1.0000
6.750 0.6893 0.03068 0.01880 0.0230 0.1414 1.0000
7.000 0.7108 0.03151 0.01968 0.0241 0.1392 1.0000
7.250 0.7328 0.03241 0.02058 0.0251 0.1373 1.0000
7.500 0.7542 0.03338 0.02172 0.0261 0.1351 1.0000
7.750 0.7737 0.03441 0.02296 0.0272 0.1321 1.0000
8.000 0.7915 0.03542 0.02413 0.0284 0.1288 1.0000
8.250 0.8080 0.03632 0.02512 0.0297 0.1252 1.0000
8.500 0.8253 0.03724 0.02596 0.0308 0.1220 1.0000
8.750 0.8398 0.03851 0.02757 0.0322 0.1192 1.0000
9.000 0.8544 0.03984 0.02915 0.0336 0.1169 1.0000
9.250 0.8676 0.04110 0.03064 0.0350 0.1143 1.0000
9.500 0.8806 0.04225 0.03191 0.0363 0.1117 1.0000
9.750 0.8949 0.04327 0.03297 0.0376 0.1094 1.0000
10.000 0.9074 0.04476 0.03460 0.0389 0.1079 1.0000
10.250 0.9146 0.04675 0.03694 0.0404 0.1071 1.0000
10.500 0.9188 0.04893 0.03943 0.0421 0.1063 1.0000
10.750 0.9187 0.05134 0.04215 0.0438 0.1055 1.0000
11.000 0.9134 0.05400 0.04509 0.0457 0.1049 1.0000
11.250 0.9000 0.05690 0.04823 0.0479 0.1044 1.0000
11.500 0.8802 0.06054 0.05208 0.0492 0.1042 1.0000
11.750 0.8512 0.06557 0.05731 0.0488 0.1042 1.0000
12.000 0.8084 0.07317 0.06505 0.0461 0.1049 1.0000
12.250 0.7499 0.08525 0.07715 0.0393 0.1060 1.0000
12.500 0.7060 0.09857 0.09044 0.0312 0.1066 1.0000
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Polar data table (+)
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