BOEING AIRFOIL J (no closed TE) (bacj-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: BOEING AIRFOIL J (no closed TE) (bacj-il) Reynolds number: 100,000 Max Cl/Cd: 25.03 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-bacj-il-100000-n5.txt Download as CSV file: xf-bacj-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING AIRFOIL J
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6982 0.07046 0.06463 -0.0582 1.0000 0.0370
-11.000 -0.7280 0.06537 0.05940 -0.0597 1.0000 0.0371
-10.750 -0.7557 0.06116 0.05505 -0.0595 1.0000 0.0370
-10.500 -0.7678 0.05844 0.05230 -0.0582 1.0000 0.0364
-10.250 -0.8025 0.05536 0.04898 -0.0556 1.0000 0.0369
-10.000 -0.8163 0.05332 0.04689 -0.0525 1.0000 0.0364
-9.750 -0.8430 0.05178 0.04514 -0.0470 1.0000 0.0368
-9.500 -0.8565 0.04972 0.04297 -0.0428 1.0000 0.0364
-9.250 -0.8785 0.04743 0.04029 -0.0375 1.0000 0.0372
-9.000 -0.8837 0.04531 0.03796 -0.0340 1.0000 0.0371
-8.750 -0.8890 0.04310 0.03543 -0.0302 1.0000 0.0372
-8.500 -0.8906 0.04047 0.03264 -0.0270 1.0000 0.0381
-8.250 -0.8869 0.03854 0.03049 -0.0241 1.0000 0.0382
-8.000 -0.8802 0.03698 0.02878 -0.0215 1.0000 0.0390
-7.750 -0.8719 0.03523 0.02678 -0.0189 1.0000 0.0389
-7.500 -0.8626 0.03380 0.02523 -0.0164 1.0000 0.0402
-7.250 -0.8512 0.03246 0.02372 -0.0141 1.0000 0.0410
-6.250 -0.7867 0.02792 0.01842 -0.0075 0.9975 0.0463
-6.000 -0.7639 0.02680 0.01718 -0.0070 0.9948 0.0483
-5.750 -0.7426 0.02579 0.01618 -0.0063 0.9921 0.0516
-5.500 -0.7209 0.02509 0.01542 -0.0056 0.9894 0.0546
-5.250 -0.6979 0.02438 0.01457 -0.0049 0.9865 0.0574
-5.000 -0.6751 0.02374 0.01393 -0.0045 0.9838 0.0608
-4.750 -0.6539 0.02321 0.01337 -0.0036 0.9811 0.0682
-4.500 -0.6341 0.02263 0.01282 -0.0025 0.9778 0.0787
-4.250 -0.6125 0.02212 0.01230 -0.0016 0.9743 0.0875
-4.000 -0.5892 0.02171 0.01188 -0.0011 0.9714 0.0988
-3.750 -0.5646 0.02134 0.01155 -0.0007 0.9686 0.1131
-3.500 -0.5472 0.02081 0.01122 0.0009 0.9648 0.1422
-3.250 -0.5256 0.02042 0.01089 0.0019 0.9611 0.1711
-3.000 -0.5021 0.02005 0.01071 0.0023 0.9577 0.2115
-2.750 -0.4760 0.01965 0.01058 0.0022 0.9549 0.2681
-2.500 -0.4587 0.01891 0.01034 0.0035 0.9511 0.3603
-2.250 -0.4441 0.01836 0.01091 0.0065 0.9468 0.5583
-2.000 -0.4205 0.01857 0.01116 0.0079 0.9434 0.6191
-1.750 -0.3938 0.01890 0.01146 0.0085 0.9400 0.6582
-1.500 -0.3724 0.01905 0.01155 0.0098 0.9359 0.6797
-1.250 -0.3509 0.01924 0.01167 0.0115 0.9309 0.7035
0.750 -0.1441 0.02039 0.01274 0.0168 0.8972 0.8155
1.000 -0.1099 0.02046 0.01278 0.0153 0.8944 0.8211
1.250 -0.0830 0.02044 0.01276 0.0152 0.8884 0.8266
1.500 -0.0499 0.02043 0.01282 0.0139 0.8834 0.8307
1.750 -0.0121 0.02043 0.01283 0.0119 0.8794 0.8353
2.000 0.0157 0.02043 0.01289 0.0116 0.8726 0.8403
2.250 0.0573 0.02020 0.01271 0.0092 0.8650 0.8431
2.500 0.0960 0.01965 0.01226 0.0079 0.8509 0.8479
2.750 0.1305 0.01928 0.01197 0.0071 0.8382 0.8523
3.000 0.1663 0.01897 0.01177 0.0058 0.8286 0.8566
3.250 0.2031 0.01826 0.01120 0.0052 0.8106 0.8609
3.500 0.2355 0.01690 0.00992 0.0065 0.7666 0.8656
3.750 0.2723 0.01581 0.00883 0.0069 0.6957 0.8699
4.000 0.3249 0.01625 0.00723 0.0047 0.2881 0.8715
4.250 0.3502 0.01743 0.00761 0.0043 0.1309 0.8765
4.500 0.3770 0.01829 0.00816 0.0041 0.0839 0.8821
4.750 0.4053 0.01907 0.00875 0.0035 0.0614 0.8874
5.000 0.4344 0.01964 0.00941 0.0029 0.0576 0.8933
5.250 0.4628 0.02033 0.01020 0.0026 0.0536 0.9000
5.500 0.4945 0.02102 0.01095 0.0013 0.0495 0.9068
5.750 0.5255 0.02202 0.01195 0.0001 0.0478 0.9137
6.000 0.5651 0.02317 0.01320 -0.0024 0.0449 0.9201
6.250 0.6064 0.02438 0.01452 -0.0053 0.0443 0.9268
6.500 0.6530 0.02611 0.01641 -0.0091 0.0423 0.9326
6.750 0.6992 0.02794 0.01842 -0.0130 0.0407 0.9389
7.000 0.7407 0.02969 0.02040 -0.0163 0.0391 0.9475
7.250 0.7780 0.03198 0.02283 -0.0192 0.0366 0.9596
7.500 0.8035 0.03405 0.02522 -0.0193 0.0362 1.0000
7.750 0.8187 0.03596 0.02743 -0.0174 0.0360 1.0000
8.000 0.8303 0.03814 0.02997 -0.0151 0.0355 1.0000
8.250 0.8411 0.04038 0.03250 -0.0128 0.0355 1.0000
8.500 0.8463 0.04282 0.03534 -0.0096 0.0345 1.0000
8.750 0.8503 0.04585 0.03868 -0.0068 0.0350 1.0000
9.000 0.8510 0.04847 0.04161 -0.0038 0.0348 1.0000
9.250 0.8475 0.05152 0.04492 -0.0004 0.0350 1.0000
9.500 0.8402 0.05411 0.04784 0.0030 0.0331 1.0000
9.750 0.8317 0.05720 0.05118 0.0058 0.0320 1.0000
10.000 0.8239 0.06098 0.05514 0.0075 0.0342 1.0000
10.250 0.8058 0.06518 0.05956 0.0090 0.0333 1.0000
10.500 0.7864 0.06997 0.06457 0.0089 0.0329 1.0000
10.750 0.7712 0.07504 0.06975 0.0075 0.0344 1.0000
11.000 0.7596 0.08006 0.07485 0.0052 0.0352 1.0000
11.250 0.7346 0.08914 0.08404 -0.0015 0.0354 1.0000
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Polar data table (+)
Polar graphs
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