BOEING 707 .54 SPAN AIRFOIL (b707d-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING 707 .54 SPAN AIRFOIL (b707d-il) Reynolds number: 1,000,000 Max Cl/Cd: 76.91 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b707d-il-1000000-n5.txt Download as CSV file: xf-b707d-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4852 0.08711 0.08550 -0.0260 1.0000 0.0061
-9.250 -0.4916 0.08094 0.07936 -0.0299 1.0000 0.0064
-9.000 -0.5016 0.07540 0.07385 -0.0341 1.0000 0.0065
-8.750 -0.5099 0.07170 0.07016 -0.0376 1.0000 0.0063
-8.500 -0.5246 0.06646 0.06487 -0.0395 1.0000 0.0066
-8.250 -0.5251 0.06196 0.06030 -0.0414 0.9918 0.0066
-8.000 -0.5052 0.05821 0.05647 -0.0459 0.9811 0.0068
-7.750 -0.4817 0.05361 0.05173 -0.0510 0.9689 0.0069
-7.500 -0.4551 0.04996 0.04795 -0.0553 0.9544 0.0072
-7.250 -0.4341 0.04629 0.04409 -0.0578 0.9332 0.0075
-7.000 -0.4217 0.04270 0.04026 -0.0575 0.9100 0.0076
-6.750 -0.4084 0.03976 0.03710 -0.0566 0.8916 0.0079
-6.500 -0.3950 0.03636 0.03345 -0.0552 0.8763 0.0083
-6.250 -0.3807 0.03248 0.02926 -0.0535 0.8630 0.0088
-6.000 -0.3626 0.02927 0.02576 -0.0519 0.8520 0.0094
-5.750 -0.3431 0.02673 0.02299 -0.0506 0.8422 0.0095
-5.000 -0.2844 0.01849 0.01382 -0.0464 0.8157 0.0100
-4.750 -0.2607 0.01711 0.01223 -0.0456 0.8080 0.0099
-4.500 -0.2362 0.01592 0.01087 -0.0451 0.8009 0.0098
-4.250 -0.2124 0.01421 0.00891 -0.0442 0.7937 0.0098
-4.000 -0.1883 0.01243 0.00689 -0.0435 0.7868 0.0098
-3.750 -0.1651 0.01083 0.00508 -0.0426 0.7798 0.0104
-3.500 -0.1395 0.01035 0.00454 -0.0422 0.7735 0.0107
-3.250 -0.1135 0.01004 0.00419 -0.0420 0.7669 0.0111
-3.000 -0.0876 0.00969 0.00379 -0.0416 0.7607 0.0115
-2.750 -0.0619 0.00934 0.00339 -0.0413 0.7540 0.0119
-2.500 -0.0362 0.00903 0.00303 -0.0409 0.7480 0.0123
-2.250 -0.0104 0.00871 0.00267 -0.0405 0.7418 0.0126
-2.000 0.0154 0.00844 0.00234 -0.0402 0.7357 0.0129
-1.750 0.0415 0.00817 0.00203 -0.0399 0.7298 0.0132
-1.500 0.0681 0.00802 0.00186 -0.0397 0.7230 0.0138
-1.250 0.0932 0.00792 0.00163 -0.0391 0.6958 0.0141
-1.000 0.1162 0.00796 0.00139 -0.0381 0.6349 0.0143
-0.750 0.1406 0.00801 0.00123 -0.0375 0.5859 0.0146
-0.500 0.1635 0.00817 0.00104 -0.0366 0.5055 0.0163
-0.250 0.1794 0.00927 0.00125 -0.0348 0.2640 0.0188
0.000 0.1997 0.01001 0.00139 -0.0337 0.0987 0.0214
0.250 0.2251 0.01018 0.00142 -0.0334 0.0479 0.0253
0.500 0.2512 0.01026 0.00144 -0.0331 0.0177 0.0315
0.750 0.2713 0.00925 0.00138 -0.0322 0.0169 0.4183
1.000 0.2914 0.00858 0.00144 -0.0310 0.0163 0.6549
1.250 0.3164 0.00852 0.00150 -0.0305 0.0161 0.7009
1.500 0.3415 0.00847 0.00157 -0.0300 0.0159 0.7401
1.750 0.3652 0.00839 0.00165 -0.0292 0.0158 0.7956
2.000 0.3883 0.00831 0.00176 -0.0281 0.0157 0.8547
2.250 0.4132 0.00824 0.00189 -0.0273 0.0157 0.9283
2.500 0.4461 0.00836 0.00202 -0.0286 0.0157 0.9502
2.750 0.4805 0.00851 0.00216 -0.0302 0.0158 0.9625
3.000 0.5148 0.00867 0.00231 -0.0318 0.0159 0.9710
3.250 0.5478 0.00884 0.00247 -0.0332 0.0159 0.9780
3.500 0.5809 0.00902 0.00263 -0.0346 0.0160 0.9829
3.750 0.6126 0.00921 0.00283 -0.0357 0.0159 0.9876
4.000 0.6437 0.00944 0.00307 -0.0367 0.0147 0.9922
4.250 0.6740 0.00966 0.00329 -0.0375 0.0144 0.9966
4.500 0.7054 0.00990 0.00354 -0.0386 0.0142 0.9996
4.750 0.7297 0.01016 0.00383 -0.0381 0.0135 1.0000
5.000 0.7511 0.01053 0.00424 -0.0369 0.0123 1.0000
5.250 0.7748 0.01067 0.00436 -0.0363 0.0116 1.0000
5.500 0.7996 0.01071 0.00438 -0.0359 0.0104 1.0000
5.750 0.8227 0.01095 0.00466 -0.0350 0.0073 1.0000
6.000 0.8471 0.01104 0.00467 -0.0346 0.0062 1.0000
6.250 0.8699 0.01131 0.00493 -0.0338 0.0054 1.0000
6.500 0.8922 0.01163 0.00530 -0.0329 0.0050 1.0000
6.750 0.9143 0.01200 0.00571 -0.0319 0.0047 1.0000
7.000 0.9360 0.01241 0.00616 -0.0310 0.0043 1.0000
7.250 0.9575 0.01284 0.00665 -0.0300 0.0040 1.0000
7.500 0.9781 0.01337 0.00724 -0.0289 0.0037 1.0000
7.750 0.9969 0.01410 0.00807 -0.0274 0.0033 1.0000
8.000 1.0182 0.01454 0.00856 -0.0265 0.0032 1.0000
8.250 1.0387 0.01505 0.00914 -0.0254 0.0032 1.0000
8.500 1.0580 0.01568 0.00984 -0.0242 0.0031 1.0000
8.750 1.0775 0.01627 0.01053 -0.0230 0.0030 1.0000
9.000 1.0954 0.01700 0.01136 -0.0216 0.0030 1.0000
9.250 1.1133 0.01771 0.01214 -0.0202 0.0029 1.0000
9.500 1.1302 0.01849 0.01303 -0.0187 0.0028 1.0000
9.750 1.1463 0.01932 0.01395 -0.0171 0.0027 1.0000
10.000 1.1609 0.02025 0.01500 -0.0153 0.0026 1.0000
10.250 1.1748 0.02119 0.01605 -0.0135 0.0025 1.0000
10.500 1.1880 0.02207 0.01704 -0.0115 0.0024 1.0000
10.750 1.1980 0.02299 0.01807 -0.0091 0.0023 1.0000
11.000 1.2061 0.02405 0.01926 -0.0064 0.0023 1.0000
11.250 1.2135 0.02516 0.02050 -0.0039 0.0022 1.0000
11.500 1.2216 0.02621 0.02164 -0.0016 0.0021 1.0000
11.750 1.2261 0.02762 0.02320 0.0008 0.0021 1.0000
12.000 1.2297 0.02914 0.02485 0.0031 0.0020 1.0000
12.250 1.2315 0.03084 0.02669 0.0052 0.0020 1.0000
12.500 1.2344 0.03248 0.02847 0.0069 0.0020 1.0000
12.750 1.2332 0.03461 0.03075 0.0087 0.0019 1.0000
13.000 1.2252 0.03749 0.03380 0.0103 0.0018 1.0000
13.250 1.2215 0.04015 0.03663 0.0113 0.0019 1.0000
13.500 1.2093 0.04387 0.04053 0.0119 0.0018 1.0000
13.750 1.1950 0.04815 0.04501 0.0118 0.0018 1.0000
14.000 1.1853 0.05221 0.04924 0.0112 0.0018 1.0000
14.250 1.1623 0.05846 0.05568 0.0092 0.0018 1.0000
14.500 1.1374 0.06578 0.06321 0.0059 0.0018 1.0000
14.750 1.1386 0.06929 0.06680 0.0042 0.0018 1.0000
15.000 1.1155 0.07781 0.07548 -0.0008 0.0018 1.0000
15.250 1.0952 0.08668 0.08451 -0.0061 0.0018 1.0000
15.500 1.0682 0.09764 0.09565 -0.0124 0.0019 1.0000
15.750 1.0361 0.11017 0.10831 -0.0193 0.0019 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING 707 .54 SPAN AIRFOIL (b707d-il)