BOEING 707 .54 SPAN AIRFOIL (b707d-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 707 .54 SPAN AIRFOIL (b707d-il) Reynolds number: 100,000 Max Cl/Cd: 53.86 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707d-il-100000.txt Download as CSV file: xf-b707d-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .54 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4583 0.09256 0.08772 -0.0298 1.0000 0.0695
-8.500 -0.4680 0.08910 0.08436 -0.0340 1.0000 0.0703
-8.250 -0.4813 0.08612 0.08139 -0.0374 1.0000 0.0708
-8.000 -0.4943 0.08407 0.07923 -0.0397 1.0000 0.0713
-7.750 -0.5021 0.08231 0.07730 -0.0400 1.0000 0.0716
-7.500 -0.4743 0.07457 0.06996 -0.0360 1.0000 0.0778
-7.250 -0.4763 0.07145 0.06683 -0.0363 1.0000 0.0809
-7.000 -0.4834 0.06938 0.06457 -0.0372 1.0000 0.0846
-6.750 -0.4940 0.06833 0.06318 -0.0361 1.0000 0.0862
-6.500 -0.4826 0.06229 0.05750 -0.0344 1.0000 0.0892
-6.250 -0.4805 0.05976 0.05497 -0.0322 1.0000 0.0933
-5.750 -0.4833 0.05491 0.04982 -0.0281 1.0000 0.1029
-5.500 -0.4794 0.05259 0.04749 -0.0258 1.0000 0.1091
-5.250 -0.4774 0.05021 0.04489 -0.0240 1.0000 0.1170
-5.000 -0.4727 0.04880 0.04318 -0.0221 1.0000 0.1294
-4.750 -0.4644 0.04639 0.04077 -0.0204 1.0000 0.1410
-4.500 -0.4581 0.04380 0.03805 -0.0189 1.0000 0.1582
-4.250 -0.4493 0.04092 0.03528 -0.0173 1.0000 0.1761
-4.000 -0.4413 0.03874 0.03310 -0.0157 1.0000 0.2056
-3.750 -0.4344 0.03654 0.03096 -0.0137 1.0000 0.2473
-3.500 -0.4261 0.03448 0.02899 -0.0114 1.0000 0.2912
-3.250 -0.4153 0.03242 0.02700 -0.0092 1.0000 0.3328
-3.000 -0.3879 0.03016 0.02466 -0.0098 0.9966 0.3664
-2.750 -0.3429 0.02773 0.02169 -0.0137 0.9918 0.3518
-2.500 -0.2600 0.02753 0.01926 -0.0176 0.9865 0.1399
-2.250 -0.2137 0.02600 0.01709 -0.0184 0.9826 0.0955
-2.000 -0.1786 0.02484 0.01558 -0.0188 0.9769 0.0827
-1.750 -0.1410 0.02355 0.01412 -0.0204 0.9721 0.0777
-1.500 -0.1046 0.02269 0.01316 -0.0219 0.9673 0.0804
-1.250 -0.0727 0.02185 0.01225 -0.0226 0.9613 0.0789
-1.000 -0.0359 0.02112 0.01156 -0.0243 0.9565 0.0790
-0.750 -0.0088 0.02065 0.01110 -0.0243 0.9499 0.0810
-0.500 0.0244 0.02030 0.01066 -0.0255 0.9438 0.0867
-0.250 0.0555 0.01999 0.01030 -0.0265 0.9377 0.1030
0.000 0.0782 0.01775 0.01029 -0.0259 0.9323 0.7036
0.250 0.1698 0.01744 0.01021 -0.0376 0.9329 1.0000
0.500 0.2124 0.01771 0.01031 -0.0409 0.9274 1.0000
0.750 0.2363 0.01801 0.01050 -0.0408 0.9181 1.0000
1.000 0.2831 0.01821 0.01062 -0.0448 0.9128 1.0000
1.250 0.3089 0.01845 0.01082 -0.0448 0.9022 1.0000
1.500 0.3563 0.01835 0.01071 -0.0484 0.8907 1.0000
1.750 0.4150 0.01786 0.01025 -0.0535 0.8782 1.0000
2.000 0.4671 0.01741 0.00987 -0.0573 0.8689 1.0000
2.250 0.4977 0.01735 0.00986 -0.0574 0.8575 1.0000
2.500 0.5256 0.01732 0.00988 -0.0569 0.8454 1.0000
2.750 0.5569 0.01705 0.00967 -0.0566 0.8318 1.0000
3.000 0.5876 0.01665 0.00937 -0.0558 0.8166 1.0000
3.250 0.6183 0.01606 0.00883 -0.0546 0.7994 1.0000
3.500 0.6403 0.01558 0.00839 -0.0518 0.7754 1.0000
3.750 0.6640 0.01511 0.00795 -0.0494 0.7514 1.0000
4.000 0.6876 0.01484 0.00777 -0.0475 0.7299 1.0000
4.250 0.7094 0.01475 0.00778 -0.0456 0.7078 1.0000
4.500 0.7321 0.01461 0.00773 -0.0437 0.6828 1.0000
4.750 0.7529 0.01451 0.00770 -0.0416 0.6488 1.0000
5.000 0.7719 0.01442 0.00758 -0.0389 0.5956 1.0000
5.250 0.7874 0.01462 0.00748 -0.0357 0.5015 1.0000
5.500 0.8003 0.01539 0.00771 -0.0326 0.4049 1.0000
5.750 0.8134 0.01639 0.00833 -0.0301 0.3329 1.0000
6.000 0.8244 0.01762 0.00913 -0.0274 0.2429 1.0000
6.250 0.8290 0.01983 0.01033 -0.0238 0.1059 1.0000
6.500 0.8419 0.02140 0.01167 -0.0212 0.0818 1.0000
6.750 0.8550 0.02294 0.01319 -0.0188 0.0705 1.0000
7.000 0.8675 0.02494 0.01508 -0.0166 0.0620 1.0000
7.250 0.8882 0.02675 0.01698 -0.0152 0.0560 1.0000
7.500 0.9130 0.02976 0.01988 -0.0148 0.0496 1.0000
7.750 0.9383 0.03178 0.02219 -0.0140 0.0473 1.0000
8.000 0.9630 0.03448 0.02522 -0.0130 0.0462 1.0000
8.250 0.9838 0.03747 0.02865 -0.0116 0.0463 1.0000
8.500 1.0001 0.04086 0.03250 -0.0098 0.0471 1.0000
8.750 1.0123 0.04468 0.03675 -0.0078 0.0483 1.0000
9.000 1.0203 0.04857 0.04104 -0.0057 0.0490 1.0000
9.250 1.0240 0.05252 0.04538 -0.0033 0.0493 1.0000
9.500 1.0369 0.05723 0.05028 -0.0023 0.0513 1.0000
9.750 1.0085 0.06048 0.05453 0.0037 0.0579 1.0000
10.000 1.0001 0.06527 0.05954 0.0058 0.0605 1.0000
10.250 0.9111 0.06042 0.05527 0.0134 0.0653 1.0000
10.500 0.8649 0.06567 0.06080 0.0146 0.0673 1.0000
10.750 0.8289 0.07174 0.06703 0.0133 0.0682 1.0000
11.000 0.7951 0.07878 0.07414 0.0103 0.0683 1.0000
11.250 0.7632 0.08693 0.08236 0.0057 0.0678 1.0000
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Polar data table (+)
Polar graphs
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