BOEING 707 .08 SPAN AIRFOIL (b707a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 707 .08 SPAN AIRFOIL (b707a-il) Reynolds number: 100,000 Max Cl/Cd: 26.4 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707a-il-100000.txt Download as CSV file: xf-b707a-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6627 0.11554 0.11018 0.0063 1.0000 0.1717
-10.750 -0.6340 0.11385 0.10846 0.0095 1.0000 0.1768
-10.500 -0.9515 0.06352 0.05671 -0.0224 1.0000 0.0645
-10.250 -0.9438 0.05961 0.05275 -0.0216 1.0000 0.0633
-10.000 -0.9475 0.05600 0.04891 -0.0196 1.0000 0.0619
-9.750 -0.9547 0.05222 0.04473 -0.0169 1.0000 0.0602
-9.500 -0.9717 0.04914 0.04068 -0.0119 1.0000 0.0579
-9.250 -0.9601 0.04616 0.03746 -0.0103 1.0000 0.0577
-9.000 -0.9484 0.04345 0.03444 -0.0083 1.0000 0.0577
-8.750 -0.9348 0.04091 0.03161 -0.0065 1.0000 0.0581
-8.500 -0.9125 0.03851 0.02934 -0.0061 1.0000 0.0596
-8.250 -0.8949 0.03666 0.02736 -0.0047 1.0000 0.0616
-8.000 -0.8780 0.03484 0.02524 -0.0029 1.0000 0.0635
-7.750 -0.8599 0.03295 0.02304 -0.0011 1.0000 0.0656
-7.500 -0.8352 0.03088 0.02111 -0.0005 1.0000 0.0685
-7.250 -0.8135 0.02924 0.01926 0.0009 1.0000 0.0724
-7.000 -0.7898 0.02710 0.01729 0.0018 1.0000 0.0786
-6.750 -0.7703 0.02507 0.01540 0.0034 1.0000 0.0899
-6.500 -0.7537 0.02357 0.01405 0.0054 1.0000 0.1008
-6.250 -0.7382 0.02247 0.01309 0.0074 1.0000 0.1104
-6.000 -0.7224 0.02157 0.01228 0.0093 1.0000 0.1202
-5.750 -0.7049 0.02084 0.01155 0.0112 1.0000 0.1287
-5.500 -0.6884 0.02004 0.01081 0.0131 1.0000 0.1375
-5.250 -0.6712 0.01926 0.01009 0.0149 1.0000 0.1465
-5.000 -0.6530 0.01852 0.00939 0.0166 1.0000 0.1583
-4.750 -0.6351 0.01765 0.00869 0.0182 1.0000 0.1773
-4.500 -0.6188 0.01642 0.00778 0.0200 1.0000 0.2358
-4.250 -0.6060 0.01585 0.00803 0.0228 1.0000 0.4111
-4.000 -0.5857 0.01611 0.00826 0.0249 1.0000 0.4518
-3.750 -0.5651 0.01634 0.00846 0.0270 1.0000 0.4811
-3.500 -0.5435 0.01645 0.00857 0.0288 1.0000 0.5023
-3.250 -0.5209 0.01645 0.00864 0.0305 1.0000 0.5186
-3.000 -0.4982 0.01639 0.00858 0.0319 1.0000 0.5322
-2.750 -0.4756 0.01635 0.00848 0.0333 1.0000 0.5456
-2.500 -0.4521 0.01621 0.00844 0.0346 1.0000 0.5557
-2.250 -0.4292 0.01612 0.00828 0.0358 1.0000 0.5666
-2.000 -0.4056 0.01595 0.00821 0.0370 1.0000 0.5753
-1.750 -0.3818 0.01579 0.00806 0.0380 1.0000 0.5826
-1.500 -0.3584 0.01564 0.00789 0.0389 1.0000 0.5906
-1.250 -0.3349 0.01550 0.00786 0.0401 1.0000 0.5988
-1.000 -0.3115 0.01538 0.00775 0.0410 1.0000 0.6074
-0.750 -0.2883 0.01522 0.00773 0.0420 1.0000 0.6153
-0.500 -0.2655 0.01514 0.00775 0.0431 1.0000 0.6252
-0.250 -0.1995 0.01515 0.00799 0.0357 0.9709 0.6402
0.000 -0.1393 0.01475 0.00764 0.0301 0.8730 0.6501
0.250 -0.0452 0.01534 0.00720 0.0188 0.6046 0.6555
0.500 -0.0172 0.01562 0.00715 0.0190 0.5472 0.6604
0.750 0.0106 0.01580 0.00713 0.0192 0.5111 0.6665
1.000 0.0377 0.01594 0.00710 0.0194 0.4827 0.6730
1.250 0.0647 0.01598 0.00713 0.0199 0.4556 0.6792
1.500 0.0911 0.01607 0.00714 0.0203 0.4302 0.6872
1.750 0.1174 0.01626 0.00718 0.0208 0.4058 0.6958
2.000 0.1440 0.01659 0.00742 0.0212 0.3813 0.7060
2.250 0.1708 0.01713 0.00782 0.0215 0.3584 0.7187
2.500 0.1981 0.01781 0.00840 0.0217 0.3382 0.7358
2.750 0.2251 0.01840 0.00903 0.0220 0.3204 0.7576
3.000 0.2516 0.01874 0.00957 0.0226 0.3035 0.7891
3.250 0.2821 0.01895 0.01008 0.0227 0.2872 0.8404
3.500 0.3351 0.01914 0.01055 0.0185 0.2673 0.9072
3.750 0.4001 0.01908 0.01063 0.0115 0.2457 0.9640
4.000 0.4706 0.01871 0.01015 0.0030 0.2265 1.0000
4.250 0.4863 0.01883 0.01014 0.0049 0.2150 1.0000
4.500 0.5050 0.01931 0.01047 0.0064 0.2012 1.0000
4.750 0.5246 0.02010 0.01117 0.0079 0.1862 1.0000
5.000 0.5446 0.02099 0.01203 0.0093 0.1726 1.0000
5.250 0.5655 0.02189 0.01284 0.0105 0.1626 1.0000
5.500 0.5867 0.02255 0.01346 0.0117 0.1543 1.0000
5.750 0.6077 0.02342 0.01442 0.0130 0.1473 1.0000
6.000 0.6291 0.02403 0.01501 0.0142 0.1413 1.0000
6.250 0.6501 0.02481 0.01581 0.0154 0.1358 1.0000
6.500 0.6709 0.02555 0.01667 0.0166 0.1308 1.0000
6.750 0.6930 0.02625 0.01729 0.0176 0.1272 1.0000
7.000 0.7132 0.02741 0.01861 0.0189 0.1242 1.0000
7.250 0.7324 0.02854 0.02000 0.0202 0.1209 1.0000
7.500 0.7529 0.02943 0.02100 0.0214 0.1178 1.0000
7.750 0.7741 0.03040 0.02198 0.0224 0.1156 1.0000
8.000 0.7928 0.03174 0.02347 0.0236 0.1133 1.0000
8.250 0.8074 0.03327 0.02541 0.0253 0.1101 1.0000
8.500 0.8266 0.03419 0.02643 0.0265 0.1068 1.0000
8.750 0.8506 0.03460 0.02666 0.0270 0.1038 1.0000
9.000 0.8587 0.03675 0.02936 0.0292 0.1005 1.0000
9.250 0.8797 0.03688 0.02949 0.0302 0.0957 1.0000
9.500 0.8930 0.03791 0.03072 0.0317 0.0907 1.0000
9.750 0.9191 0.03726 0.02985 0.0324 0.0855 1.0000
10.000 0.9242 0.03916 0.03218 0.0345 0.0799 1.0000
10.250 0.9356 0.04016 0.03327 0.0362 0.0739 1.0000
10.500 0.9476 0.04116 0.03427 0.0378 0.0685 1.0000
10.750 0.9619 0.04209 0.03511 0.0391 0.0642 1.0000
11.000 0.9484 0.04592 0.03939 0.0416 0.0619 1.0000
11.250 0.9299 0.04968 0.04348 0.0440 0.0604 1.0000
11.500 0.9586 0.04974 0.04320 0.0446 0.0575 1.0000
11.750 0.9331 0.05356 0.04730 0.0469 0.0574 1.0000
12.000 0.9001 0.05870 0.05271 0.0462 0.0574 1.0000
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Polar data table (+)
Polar graphs
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