BOEING 707 .08 SPAN AIRFOIL (b707a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: BOEING 707 .08 SPAN AIRFOIL (b707a-il) Reynolds number: 50,000 Max Cl/Cd: 16.35 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707a-il-50000.txt Download as CSV file: xf-b707a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .08 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7787 0.08883 0.08120 -0.0156 1.0000 0.1287
-10.500 -0.7986 0.08206 0.07437 -0.0188 1.0000 0.1251
-10.250 -0.8251 0.07603 0.06829 -0.0209 1.0000 0.1225
-10.000 -0.8636 0.07100 0.06312 -0.0204 1.0000 0.1201
-9.750 -0.9067 0.06601 0.05771 -0.0186 1.0000 0.1169
-9.500 -0.9292 0.06223 0.05336 -0.0161 1.0000 0.1143
-9.250 -0.9252 0.05825 0.04916 -0.0147 1.0000 0.1133
-9.000 -0.9216 0.05453 0.04512 -0.0129 1.0000 0.1126
-8.750 -0.9152 0.05106 0.04130 -0.0110 1.0000 0.1123
-8.500 -0.9054 0.04783 0.03768 -0.0092 1.0000 0.1126
-8.250 -0.8945 0.04495 0.03432 -0.0071 1.0000 0.1141
-8.000 -0.8746 0.04186 0.03118 -0.0061 1.0000 0.1180
-7.750 -0.8541 0.03938 0.02857 -0.0048 1.0000 0.1239
-7.500 -0.8313 0.03664 0.02572 -0.0036 1.0000 0.1317
-7.250 -0.8061 0.03414 0.02322 -0.0025 1.0000 0.1441
-7.000 -0.7852 0.03217 0.02142 -0.0010 1.0000 0.1609
-6.750 -0.7726 0.03080 0.02009 0.0010 1.0000 0.1830
-6.500 -0.7569 0.02936 0.01876 0.0028 1.0000 0.2037
-6.250 -0.7366 0.02785 0.01732 0.0044 1.0000 0.2207
-6.000 -0.7162 0.02632 0.01587 0.0061 1.0000 0.2365
-5.750 -0.7000 0.02465 0.01448 0.0082 1.0000 0.2557
-5.500 -0.6908 0.02309 0.01387 0.0113 1.0000 0.2913
-5.250 -0.6816 0.02327 0.01457 0.0158 1.0000 0.4243
-5.000 -0.6699 0.02448 0.01587 0.0216 1.0000 0.5016
-4.750 -0.6525 0.02554 0.01696 0.0264 1.0000 0.5481
-4.500 -0.6293 0.02664 0.01814 0.0308 1.0000 0.5809
-4.250 -0.6108 0.02721 0.01867 0.0349 1.0000 0.6103
-4.000 -0.5859 0.02823 0.01974 0.0393 1.0000 0.6369
-3.750 -0.5659 0.02826 0.01970 0.0423 1.0000 0.6553
-3.500 -0.5486 0.02795 0.01929 0.0449 1.0000 0.6717
-3.250 -0.5229 0.02816 0.01948 0.0476 1.0000 0.6866
-3.000 -0.5025 0.02783 0.01911 0.0498 1.0000 0.6995
-2.750 -0.4857 0.02721 0.01843 0.0517 1.0000 0.7116
-2.500 -0.4616 0.02706 0.01828 0.0538 1.0000 0.7238
-2.250 -0.4456 0.02647 0.01766 0.0560 1.0000 0.7373
-2.000 -0.4253 0.02625 0.01746 0.0585 1.0000 0.7532
-1.750 -0.3997 0.02617 0.01743 0.0605 1.0000 0.7683
-1.500 -0.3848 0.02541 0.01668 0.0623 1.0000 0.7787
-1.250 -0.3603 0.02493 0.01622 0.0630 1.0000 0.7863
-1.000 -0.3430 0.02408 0.01539 0.0638 1.0000 0.7917
-0.750 -0.3270 0.02317 0.01449 0.0644 1.0000 0.7968
-0.500 -0.3023 0.02260 0.01400 0.0644 1.0000 0.8011
-0.250 -0.2788 0.02204 0.01352 0.0644 1.0000 0.8066
0.000 -0.2579 0.02144 0.01302 0.0645 1.0000 0.8135
0.250 -0.2368 0.02094 0.01265 0.0646 1.0000 0.8205
0.500 -0.2149 0.02061 0.01254 0.0646 1.0000 0.8284
0.750 -0.1099 0.02089 0.01322 0.0478 0.9125 0.8357
1.000 0.0305 0.02133 0.01293 0.0302 0.7005 0.8401
1.250 0.0668 0.02190 0.01310 0.0294 0.6323 0.8486
1.500 0.0957 0.02224 0.01318 0.0294 0.5853 0.8609
1.750 0.1341 0.02282 0.01355 0.0279 0.5396 0.8732
2.000 0.1765 0.02347 0.01411 0.0255 0.4974 0.8876
2.250 0.2221 0.02414 0.01467 0.0223 0.4621 0.9044
2.500 0.2744 0.02486 0.01543 0.0177 0.4303 0.9231
2.750 0.3323 0.02572 0.01626 0.0118 0.4030 0.9448
3.000 0.3942 0.02675 0.01733 0.0045 0.3768 0.9720
3.250 0.4535 0.02774 0.01828 -0.0034 0.3457 1.0000
3.500 0.4458 0.02790 0.01816 -0.0002 0.3371 1.0000
3.750 0.4538 0.02873 0.01886 0.0022 0.3206 1.0000
4.000 0.4673 0.02988 0.02005 0.0041 0.3041 1.0000
4.250 0.4827 0.03122 0.02149 0.0058 0.2888 1.0000
4.500 0.4984 0.03269 0.02310 0.0075 0.2738 1.0000
4.750 0.5145 0.03417 0.02473 0.0092 0.2603 1.0000
5.000 0.5315 0.03563 0.02626 0.0108 0.2488 1.0000
5.250 0.5484 0.03731 0.02808 0.0123 0.2402 1.0000
5.500 0.5609 0.03992 0.03099 0.0137 0.2341 1.0000
5.750 0.5753 0.04221 0.03344 0.0150 0.2287 1.0000
6.000 0.5919 0.04424 0.03549 0.0162 0.2245 1.0000
6.250 0.5966 0.04824 0.03987 0.0170 0.2228 1.0000
6.500 0.6011 0.05239 0.04427 0.0173 0.2230 1.0000
6.750 0.6069 0.05649 0.04855 0.0175 0.2243 1.0000
7.000 0.5799 0.06511 0.05751 0.0139 0.2357 1.0000
7.250 0.5697 0.07227 0.06474 0.0107 0.2507 1.0000
7.500 0.4481 0.08804 0.08034 -0.0099 0.3818 1.0000
7.750 0.4628 0.09127 0.08356 -0.0098 0.3750 1.0000
8.000 0.4644 0.09447 0.08670 -0.0100 0.3687 1.0000
8.250 0.4671 0.09699 0.08920 -0.0100 0.3588 1.0000
8.500 0.5036 0.10245 0.09471 -0.0097 0.3540 1.0000
8.750 0.4728 0.10290 0.09503 -0.0105 0.3440 1.0000
9.000 0.4961 0.10707 0.09923 -0.0103 0.3385 1.0000
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Polar data table (+)
Polar graphs
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