BOEING 707 .99 SPAN AIRFOIL (b707e-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 707 .99 SPAN AIRFOIL (b707e-il) Reynolds number: 1,000,000 Max Cl/Cd: 77.68 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b707e-il-1000000.txt Download as CSV file: xf-b707e-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 707 .99 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3718 0.08509 0.08356 -0.0306 1.0000 0.0108
-9.250 -0.3743 0.08065 0.07913 -0.0320 1.0000 0.0109
-9.000 -0.3778 0.07611 0.07460 -0.0334 1.0000 0.0109
-8.750 -0.3820 0.07165 0.07015 -0.0350 1.0000 0.0109
-8.500 -0.3881 0.06704 0.06556 -0.0371 1.0000 0.0109
-7.250 -0.4941 0.05427 0.05238 -0.0417 0.9965 0.0112
-7.000 -0.4756 0.05046 0.04848 -0.0449 0.9926 0.0116
-6.750 -0.4535 0.04775 0.04570 -0.0472 0.9878 0.0120
-6.500 -0.4276 0.04450 0.04231 -0.0501 0.9834 0.0125
-6.250 -0.4045 0.04145 0.03913 -0.0515 0.9748 0.0134
-5.750 -0.3342 0.03648 0.03376 -0.0557 0.9577 0.0160
-5.500 -0.3011 0.03348 0.03052 -0.0579 0.9394 0.0162
-5.250 -0.2745 0.03062 0.02734 -0.0585 0.9037 0.0162
-5.000 -0.2557 0.02829 0.02466 -0.0571 0.8725 0.0163
-4.750 -0.2371 0.02620 0.02226 -0.0555 0.8557 0.0163
-4.250 -0.2058 0.01905 0.01440 -0.0521 0.8353 0.0176
-4.000 -0.1827 0.01798 0.01319 -0.0515 0.8281 0.0181
-3.750 -0.1588 0.01700 0.01209 -0.0510 0.8213 0.0187
-3.500 -0.1344 0.01616 0.01112 -0.0504 0.8153 0.0196
-3.250 -0.1091 0.01525 0.01010 -0.0499 0.8097 0.0207
-3.000 -0.0833 0.01448 0.00919 -0.0494 0.8043 0.0219
-2.750 -0.0561 0.01453 0.00915 -0.0491 0.7993 0.0239
-2.500 -0.0291 0.01460 0.00913 -0.0487 0.7939 0.0245
-2.250 -0.0055 0.01176 0.00611 -0.0480 0.7888 0.0275
-2.000 0.0206 0.01124 0.00560 -0.0478 0.7830 0.0300
-1.750 0.0475 0.01111 0.00542 -0.0475 0.7768 0.0340
-1.500 0.0725 0.01019 0.00448 -0.0471 0.7708 0.0407
-1.250 0.0982 0.00917 0.00334 -0.0458 0.7644 0.0261
-1.000 0.1235 0.00890 0.00303 -0.0452 0.7521 0.0256
-0.750 0.1488 0.00869 0.00279 -0.0446 0.7371 0.0250
-0.500 0.1720 0.00810 0.00214 -0.0436 0.7263 0.0233
-0.250 0.1967 0.00779 0.00175 -0.0429 0.7117 0.0225
0.000 0.2215 0.00761 0.00146 -0.0422 0.6916 0.0219
0.250 0.2466 0.00751 0.00124 -0.0416 0.6679 0.0215
0.500 0.2717 0.00748 0.00109 -0.0410 0.6418 0.0213
0.750 0.2966 0.00751 0.00098 -0.0403 0.6101 0.0215
1.000 0.3203 0.00768 0.00091 -0.0395 0.5533 0.0222
1.250 0.3326 0.00892 0.00115 -0.0369 0.3069 0.0236
1.500 0.3403 0.00916 0.00143 -0.0338 0.0846 0.4185
1.750 0.3534 0.00848 0.00157 -0.0311 0.0395 0.7295
2.000 0.3704 0.00811 0.00170 -0.0286 0.0368 0.8735
2.250 0.4359 0.00821 0.00200 -0.0369 0.0354 0.9809
2.500 0.4849 0.00841 0.00219 -0.0417 0.0346 0.9913
2.750 0.5188 0.00857 0.00233 -0.0432 0.0343 0.9946
3.000 0.5547 0.00872 0.00248 -0.0452 0.0342 0.9971
3.250 0.5902 0.00889 0.00266 -0.0471 0.0337 0.9993
3.500 0.6171 0.00909 0.00286 -0.0471 0.0327 1.0000
3.750 0.6386 0.00929 0.00308 -0.0459 0.0321 1.0000
4.000 0.6607 0.00945 0.00324 -0.0449 0.0316 1.0000
4.250 0.6829 0.00962 0.00339 -0.0439 0.0302 1.0000
4.500 0.7050 0.00981 0.00359 -0.0428 0.0289 1.0000
4.750 0.7272 0.01002 0.00379 -0.0418 0.0271 1.0000
5.000 0.7481 0.01034 0.00412 -0.0406 0.0239 1.0000
5.250 0.7738 0.01026 0.00399 -0.0403 0.0221 1.0000
5.500 0.7985 0.01028 0.00395 -0.0398 0.0174 1.0000
5.750 0.8192 0.01066 0.00419 -0.0385 0.0097 1.0000
6.000 0.8412 0.01095 0.00451 -0.0374 0.0085 1.0000
6.250 0.8631 0.01126 0.00488 -0.0362 0.0079 1.0000
6.500 0.8841 0.01166 0.00534 -0.0350 0.0075 1.0000
6.750 0.9044 0.01213 0.00591 -0.0336 0.0070 1.0000
7.000 0.9242 0.01265 0.00649 -0.0321 0.0067 1.0000
7.250 0.9437 0.01319 0.00711 -0.0306 0.0065 1.0000
7.500 0.9622 0.01383 0.00782 -0.0290 0.0063 1.0000
7.750 0.9800 0.01452 0.00858 -0.0273 0.0060 1.0000
8.000 0.9951 0.01546 0.00959 -0.0252 0.0056 1.0000
8.250 1.0103 0.01642 0.01065 -0.0231 0.0055 1.0000
8.500 1.0261 0.01735 0.01167 -0.0212 0.0055 1.0000
8.750 1.0381 0.01882 0.01326 -0.0187 0.0054 1.0000
9.000 1.0530 0.02009 0.01465 -0.0168 0.0053 1.0000
9.250 1.0671 0.02164 0.01633 -0.0148 0.0053 1.0000
9.500 1.0823 0.02278 0.01767 -0.0130 0.0050 1.0000
9.750 1.0947 0.02451 0.01958 -0.0109 0.0050 1.0000
10.000 1.1104 0.02517 0.02039 -0.0092 0.0048 1.0000
10.250 1.1204 0.02686 0.02227 -0.0068 0.0047 1.0000
10.500 1.1284 0.02826 0.02386 -0.0041 0.0045 1.0000
10.750 1.1283 0.03038 0.02622 -0.0004 0.0045 1.0000
11.000 1.1238 0.03280 0.02889 0.0036 0.0045 1.0000
11.250 1.1152 0.03552 0.03186 0.0075 0.0045 1.0000
11.500 1.0974 0.03921 0.03584 0.0114 0.0046 1.0000
11.750 1.0813 0.04269 0.03956 0.0140 0.0047 1.0000
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Polar data table (+)
Polar graphs
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