BOEING AIRFOIL J (no closed TE) (bacj-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING AIRFOIL J (no closed TE) (bacj-il) Reynolds number: 50,000 Max Cl/Cd: 20.37 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-bacj-il-50000-n5.txt Download as CSV file: xf-bacj-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5574 0.10061 0.09271 -0.0392 1.0000 0.0669
-10.500 -0.5689 0.09498 0.08716 -0.0414 1.0000 0.0663
-10.250 -0.5905 0.08703 0.07925 -0.0455 1.0000 0.0659
-10.000 -0.6245 0.07855 0.07080 -0.0500 1.0000 0.0632
-9.500 -0.7217 0.06637 0.05839 -0.0515 1.0000 0.0602
-9.250 -0.7208 0.06523 0.05733 -0.0491 1.0000 0.0625
-9.000 -0.7775 0.06082 0.05251 -0.0433 1.0000 0.0596
-8.750 -0.7808 0.05903 0.05072 -0.0406 1.0000 0.0610
-8.500 -0.7856 0.05702 0.04865 -0.0376 1.0000 0.0625
-8.250 -0.7968 0.05416 0.04554 -0.0342 1.0000 0.0625
-8.000 -0.8027 0.05155 0.04268 -0.0309 1.0000 0.0631
-7.750 -0.8056 0.04890 0.03973 -0.0278 1.0000 0.0632
-7.500 -0.8051 0.04644 0.03697 -0.0248 1.0000 0.0646
-7.250 -0.8012 0.04402 0.03424 -0.0220 1.0000 0.0654
-7.000 -0.7948 0.04173 0.03159 -0.0193 1.0000 0.0671
-6.750 -0.7854 0.03951 0.02899 -0.0169 1.0000 0.0690
-6.500 -0.7723 0.03775 0.02684 -0.0148 1.0000 0.0703
-6.250 -0.7576 0.03612 0.02501 -0.0129 1.0000 0.0727
-6.000 -0.7427 0.03487 0.02377 -0.0111 1.0000 0.0763
-5.750 -0.7267 0.03356 0.02222 -0.0091 1.0000 0.0827
-5.500 -0.7090 0.03234 0.02092 -0.0076 1.0000 0.0874
-5.250 -0.6917 0.03132 0.01986 -0.0058 1.0000 0.0939
-5.000 -0.6712 0.03028 0.01858 -0.0043 1.0000 0.0999
-4.750 -0.6540 0.02924 0.01761 -0.0024 1.0000 0.1083
-4.500 -0.6386 0.02832 0.01672 -0.0004 1.0000 0.1224
-4.250 -0.6236 0.02746 0.01587 0.0018 1.0000 0.1348
-4.000 -0.6096 0.02664 0.01508 0.0040 1.0000 0.1498
-3.750 -0.5964 0.02578 0.01429 0.0065 1.0000 0.1710
-3.500 -0.5843 0.02492 0.01358 0.0089 1.0000 0.2027
-3.250 -0.5714 0.02397 0.01288 0.0110 1.0000 0.2452
-3.000 -0.5592 0.02284 0.01223 0.0130 1.0000 0.3178
-2.750 -0.5498 0.02179 0.01231 0.0166 1.0000 0.4570
-2.500 -0.5391 0.02199 0.01308 0.0214 1.0000 0.5936
-2.250 -0.5276 0.02226 0.01325 0.0254 1.0000 0.6570
-2.000 -0.5149 0.02253 0.01338 0.0292 1.0000 0.6979
-1.750 -0.5009 0.02277 0.01354 0.0325 1.0000 0.7281
-1.500 -0.4872 0.02297 0.01366 0.0359 1.0000 0.7572
-1.250 -0.4714 0.02305 0.01359 0.0388 1.0000 0.7814
-1.000 -0.4523 0.02306 0.01350 0.0410 1.0000 0.8041
-0.750 -0.4331 0.02302 0.01341 0.0433 1.0000 0.8306
-0.500 -0.4102 0.02289 0.01322 0.0447 1.0000 0.8575
-0.250 -0.3783 0.02276 0.01309 0.0444 1.0000 0.8863
0.000 -0.3325 0.02276 0.01300 0.0410 1.0000 0.9105
0.250 -0.2907 0.02282 0.01298 0.0376 1.0000 0.9217
0.500 -0.2504 0.02301 0.01312 0.0341 0.9972 0.9294
0.750 -0.2039 0.02333 0.01339 0.0295 0.9925 0.9347
1.000 -0.1599 0.02364 0.01366 0.0253 0.9876 0.9408
1.250 -0.1111 0.02415 0.01416 0.0203 0.9829 0.9460
1.500 -0.0686 0.02440 0.01449 0.0163 0.9776 0.9525
1.750 -0.0240 0.02481 0.01494 0.0121 0.9722 0.9587
2.000 0.0194 0.02519 0.01541 0.0080 0.9656 0.9656
2.250 0.0632 0.02556 0.01588 0.0040 0.9580 0.9723
2.500 0.1062 0.02590 0.01638 0.0000 0.9496 0.9826
2.750 0.1535 0.02632 0.01696 -0.0044 0.9385 0.9916
3.000 0.1862 0.02705 0.01778 -0.0060 0.9230 1.0000
3.250 0.2258 0.02679 0.01770 -0.0082 0.9010 1.0000
3.500 0.2665 0.02656 0.01771 -0.0104 0.8818 1.0000
3.750 0.3075 0.02585 0.01731 -0.0118 0.8543 1.0000
4.000 0.3612 0.02265 0.01437 -0.0107 0.7566 1.0000
4.250 0.4437 0.02290 0.01181 -0.0142 0.1935 1.0000
4.500 0.4452 0.02385 0.01225 -0.0103 0.1359 1.0000
4.750 0.4551 0.02476 0.01290 -0.0077 0.1098 1.0000
5.000 0.4705 0.02557 0.01362 -0.0060 0.1011 1.0000
5.250 0.4917 0.02648 0.01454 -0.0051 0.0902 1.0000
5.500 0.5196 0.02745 0.01561 -0.0052 0.0822 1.0000
5.750 0.5575 0.02862 0.01684 -0.0069 0.0772 1.0000
6.000 0.6155 0.03068 0.01898 -0.0122 0.0703 1.0000
6.250 0.6596 0.03238 0.02090 -0.0151 0.0668 1.0000
6.500 0.6907 0.03405 0.02266 -0.0160 0.0626 1.0000
6.750 0.7232 0.03628 0.02500 -0.0172 0.0608 1.0000
7.000 0.7483 0.03863 0.02762 -0.0171 0.0598 1.0000
7.250 0.7654 0.04083 0.03022 -0.0154 0.0591 1.0000
7.500 0.7795 0.04302 0.03272 -0.0134 0.0588 1.0000
7.750 0.7902 0.04542 0.03546 -0.0110 0.0585 1.0000
8.000 0.7958 0.04790 0.03833 -0.0079 0.0577 1.0000
8.250 0.8014 0.05049 0.04123 -0.0052 0.0577 1.0000
8.500 0.8012 0.05327 0.04437 -0.0019 0.0568 1.0000
8.750 0.7982 0.05606 0.04749 0.0013 0.0552 1.0000
9.000 0.7950 0.05887 0.05054 0.0041 0.0549 1.0000
9.250 0.7867 0.06200 0.05389 0.0070 0.0542 1.0000
9.500 0.7776 0.06558 0.05770 0.0093 0.0548 1.0000
9.750 0.7631 0.06963 0.06197 0.0108 0.0548 1.0000
10.000 0.7543 0.07328 0.06574 0.0111 0.0558 1.0000
10.250 0.7350 0.07833 0.07093 0.0102 0.0556 1.0000
10.500 0.7183 0.08400 0.07671 0.0078 0.0564 1.0000
10.750 0.7066 0.08974 0.08246 0.0044 0.0573 1.0000
11.000 0.7002 0.09541 0.08815 0.0009 0.0582 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING AIRFOIL J (no closed TE) (bacj-il)