BOEING 707 .40 SPAN AIRFOIL (b707c-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 707 .40 SPAN AIRFOIL (b707c-il) Reynolds number: 50,000 Max Cl/Cd: 33.65 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707c-il-50000.txt Download as CSV file: xf-b707c-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .40 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4982 0.12200 0.11456 -0.0121 1.0000 0.1765
-10.250 -0.4926 0.11825 0.11086 -0.0122 1.0000 0.1840
-10.000 -0.5025 0.11628 0.10898 -0.0141 1.0000 0.1932
-9.750 -0.4920 0.11243 0.10515 -0.0131 1.0000 0.2066
-9.500 -0.4853 0.10900 0.10177 -0.0125 1.0000 0.2203
-9.250 -0.4819 0.10589 0.09872 -0.0121 1.0000 0.2350
-9.000 -0.4842 0.10322 0.09613 -0.0119 1.0000 0.2517
-8.750 -0.4710 0.09987 0.09280 -0.0097 1.0000 0.2775
-8.500 -0.4669 0.09761 0.09060 -0.0076 1.0000 0.3089
-8.250 -0.4588 0.09585 0.08891 -0.0041 1.0000 0.3531
-8.000 -0.4641 0.09459 0.08775 -0.0014 1.0000 0.3862
-7.750 -0.4213 0.09048 0.08357 0.0032 1.0000 0.4476
-7.500 -0.4016 0.08761 0.08072 0.0063 1.0000 0.4972
-7.250 -0.3685 0.08386 0.07697 0.0090 1.0000 0.5574
-7.000 -0.3380 0.08060 0.07369 0.0112 1.0000 0.6219
-6.750 -0.3108 0.07766 0.07077 0.0131 1.0000 0.6907
-6.500 -0.2665 0.07261 0.06570 0.0116 1.0000 0.7618
-6.250 -0.2503 0.06931 0.06244 0.0108 1.0000 0.7794
-6.000 -0.2462 0.06724 0.06043 0.0118 1.0000 0.7928
-5.000 -0.3023 0.04420 0.03872 0.0104 1.0000 0.5680
-4.750 -0.3544 0.04161 0.03639 0.0143 1.0000 0.5357
-4.500 -0.4017 0.03795 0.03293 0.0153 1.0000 0.5039
-4.250 -0.4755 0.04388 0.03760 -0.0032 1.0000 0.4018
-4.000 -0.4395 0.03744 0.03001 -0.0126 1.0000 0.3065
-3.750 -0.4053 0.03493 0.02644 -0.0145 1.0000 0.2475
-3.500 -0.3748 0.03418 0.02484 -0.0138 1.0000 0.2075
-3.250 -0.3493 0.03266 0.02295 -0.0126 1.0000 0.1838
-3.000 -0.3242 0.03247 0.02215 -0.0107 1.0000 0.1632
-2.750 -0.3025 0.03069 0.02029 -0.0096 1.0000 0.1539
-2.500 -0.2812 0.03012 0.01938 -0.0079 1.0000 0.1462
-2.250 -0.2607 0.02894 0.01808 -0.0066 1.0000 0.1414
-2.000 -0.2403 0.02864 0.01745 -0.0050 1.0000 0.1349
-1.750 -0.2198 0.02784 0.01659 -0.0038 1.0000 0.1313
-1.500 -0.1991 0.02707 0.01582 -0.0026 1.0000 0.1283
-1.250 -0.1784 0.02652 0.01523 -0.0014 1.0000 0.1263
-1.000 -0.1581 0.02610 0.01476 -0.0002 1.0000 0.1263
-0.750 -0.1395 0.02576 0.01436 0.0010 1.0000 0.1305
-0.500 -0.1218 0.02558 0.01409 0.0022 1.0000 0.1340
-0.250 -0.1043 0.02514 0.01365 0.0032 1.0000 0.1384
0.000 -0.0862 0.02497 0.01336 0.0040 1.0000 0.1475
0.250 -0.0528 0.02465 0.01312 0.0020 0.9948 0.1703
0.500 0.0201 0.02232 0.01359 -0.0050 0.9847 1.0000
0.750 0.0682 0.02313 0.01386 -0.0096 0.9705 1.0000
1.000 0.1146 0.02391 0.01435 -0.0141 0.9574 1.0000
1.250 0.1535 0.02455 0.01483 -0.0173 0.9441 1.0000
1.500 0.1897 0.02517 0.01535 -0.0200 0.9308 1.0000
1.750 0.2264 0.02578 0.01589 -0.0226 0.9175 1.0000
2.000 0.2634 0.02639 0.01648 -0.0251 0.9051 1.0000
2.250 0.3056 0.02696 0.01707 -0.0284 0.8927 1.0000
2.500 0.3684 0.02664 0.01686 -0.0335 0.8648 1.0000
2.750 0.4352 0.02582 0.01623 -0.0384 0.8368 1.0000
3.000 0.5024 0.02458 0.01519 -0.0425 0.8119 1.0000
3.250 0.5358 0.02429 0.01507 -0.0421 0.7907 1.0000
3.500 0.5846 0.02312 0.01406 -0.0426 0.7653 1.0000
3.750 0.6169 0.02218 0.01319 -0.0402 0.7318 1.0000
4.000 0.6410 0.02154 0.01262 -0.0368 0.6944 1.0000
4.250 0.6640 0.02121 0.01234 -0.0338 0.6604 1.0000
4.500 0.6841 0.02087 0.01197 -0.0301 0.6170 1.0000
4.750 0.6945 0.02069 0.01176 -0.0250 0.5528 1.0000
5.000 0.6990 0.02077 0.01138 -0.0191 0.4084 1.0000
5.250 0.6968 0.02309 0.01212 -0.0142 0.2407 1.0000
5.500 0.7066 0.02514 0.01374 -0.0116 0.2061 1.0000
5.750 0.7238 0.02656 0.01504 -0.0099 0.1841 1.0000
6.000 0.7442 0.02754 0.01609 -0.0085 0.1660 1.0000
6.250 0.7649 0.02852 0.01707 -0.0073 0.1529 1.0000
6.500 0.7883 0.02972 0.01826 -0.0063 0.1436 1.0000
6.750 0.8145 0.03125 0.01968 -0.0057 0.1325 1.0000
7.000 0.8435 0.03340 0.02173 -0.0056 0.1185 1.0000
7.250 0.8785 0.03674 0.02493 -0.0064 0.1076 1.0000
7.500 0.9032 0.03955 0.02828 -0.0053 0.1045 1.0000
7.750 0.9242 0.04266 0.03186 -0.0040 0.1030 1.0000
8.000 0.9414 0.04601 0.03568 -0.0024 0.1029 1.0000
8.250 0.9545 0.04964 0.03975 -0.0006 0.1037 1.0000
8.500 0.9656 0.05352 0.04399 0.0011 0.1051 1.0000
8.750 0.9781 0.05803 0.04872 0.0022 0.1068 1.0000
9.000 0.9620 0.06155 0.05317 0.0061 0.1117 1.0000
9.250 0.9487 0.06640 0.05845 0.0083 0.1159 1.0000
9.500 0.9467 0.07112 0.06334 0.0094 0.1194 1.0000
9.750 0.9211 0.07578 0.06833 0.0110 0.1239 1.0000
10.000 0.8757 0.08092 0.07362 0.0118 0.1267 1.0000
10.250 0.8424 0.08739 0.08012 0.0090 0.1300 1.0000
10.500 0.8168 0.09557 0.08833 0.0040 0.1365 1.0000
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Polar data table (+)
Polar graphs
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