KC-135 BL52.44 AIRFOIL (kc135a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: KC-135 BL52.44 AIRFOIL (kc135a-il) Reynolds number: 50,000 Max Cl/Cd: 24.94 at α=10.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-kc135a-il-50000.txt Download as CSV file: xf-kc135a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: KC-135 BL52.44 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3523 0.10948 0.10226 -0.0113 1.0000 0.3977
-9.750 -0.5318 0.08076 0.07393 -0.0429 1.0000 0.1890
-9.500 -0.6128 0.07457 0.06780 -0.0444 1.0000 0.1835
-9.250 -0.6357 0.06935 0.06256 -0.0432 1.0000 0.1730
-9.000 -0.6648 0.06558 0.05862 -0.0404 1.0000 0.1647
-8.750 -0.6877 0.06185 0.05470 -0.0370 1.0000 0.1579
-8.500 -0.7121 0.05939 0.05185 -0.0322 1.0000 0.1507
-8.250 -0.7131 0.05610 0.04848 -0.0294 1.0000 0.1465
-8.000 -0.7265 0.05367 0.04576 -0.0248 1.0000 0.1420
-7.750 -0.7524 0.05344 0.04478 -0.0177 1.0000 0.1368
-7.500 -0.7451 0.05036 0.04168 -0.0154 1.0000 0.1348
-7.250 -0.7428 0.04806 0.03922 -0.0123 1.0000 0.1329
-7.000 -0.7395 0.04604 0.03697 -0.0093 1.0000 0.1315
-6.750 -0.7345 0.04419 0.03482 -0.0064 1.0000 0.1312
-6.500 -0.7265 0.04248 0.03283 -0.0039 1.0000 0.1314
-6.250 -0.7158 0.04083 0.03089 -0.0016 1.0000 0.1319
-6.000 -0.7024 0.03926 0.02904 0.0004 1.0000 0.1326
-5.750 -0.6863 0.03781 0.02732 0.0022 1.0000 0.1335
-5.500 -0.6699 0.03670 0.02590 0.0039 1.0000 0.1356
-5.250 -0.6484 0.03509 0.02431 0.0047 1.0000 0.1397
-5.000 -0.6264 0.03392 0.02318 0.0057 1.0000 0.1441
-4.750 -0.6061 0.03307 0.02228 0.0070 1.0000 0.1500
-4.500 -0.5870 0.03209 0.02140 0.0084 1.0000 0.1574
-4.250 -0.5722 0.03135 0.02067 0.0104 1.0000 0.1655
-4.000 -0.5611 0.03046 0.01986 0.0126 1.0000 0.1759
-3.750 -0.5507 0.02954 0.01901 0.0148 0.9999 0.1902
-3.500 -0.5232 0.02798 0.01793 0.0133 0.9927 0.2360
-3.250 -0.5212 0.02695 0.01909 0.0188 0.9857 0.5940
-3.000 -0.5163 0.02903 0.02148 0.0285 0.9786 0.7086
-2.750 -0.2666 0.04130 0.03308 0.0156 0.9766 0.8910
-2.500 -0.2115 0.04113 0.03264 0.0085 0.9688 0.8990
-2.250 -0.1791 0.04092 0.03223 0.0051 0.9607 0.9069
-2.000 -0.1459 0.04069 0.03184 0.0016 0.9529 0.9137
-1.750 -0.1008 0.04061 0.03158 -0.0039 0.9455 0.9212
-1.500 -0.0786 0.04049 0.03136 -0.0054 0.9378 0.9290
-1.250 -0.0300 0.04047 0.03120 -0.0115 0.9303 0.9363
-1.000 -0.0048 0.04052 0.03116 -0.0135 0.9231 0.9447
-0.750 0.0399 0.04055 0.03110 -0.0189 0.9152 0.9523
-0.500 0.0809 0.04070 0.03116 -0.0237 0.9082 0.9607
-0.250 0.1145 0.04083 0.03124 -0.0273 0.9000 0.9693
0.000 0.1746 0.04096 0.03130 -0.0353 0.8928 0.9772
0.250 0.2016 0.04120 0.03153 -0.0380 0.8841 0.9867
0.500 0.2679 0.04128 0.03158 -0.0471 0.8765 0.9951
0.750 0.2757 0.04171 0.03201 -0.0466 0.8677 1.0000
1.000 0.2975 0.04202 0.03231 -0.0478 0.8599 1.0000
1.250 0.2792 0.04256 0.03283 -0.0428 0.8517 1.0000
1.500 0.2901 0.04291 0.03318 -0.0422 0.8435 1.0000
1.750 0.2728 0.04332 0.03357 -0.0374 0.8359 1.0000
2.000 0.2636 0.04355 0.03378 -0.0338 0.8278 1.0000
2.250 0.2485 0.04372 0.03390 -0.0293 0.8204 1.0000
2.500 0.2159 0.04368 0.03377 -0.0221 0.8134 1.0000
2.750 0.2298 0.04432 0.03431 -0.0212 0.8033 1.0000
3.000 0.2354 0.04500 0.03490 -0.0191 0.7926 1.0000
3.250 0.2452 0.04587 0.03570 -0.0176 0.7808 1.0000
3.500 0.2707 0.04681 0.03661 -0.0179 0.7672 1.0000
3.750 0.3025 0.04771 0.03750 -0.0189 0.7523 1.0000
4.000 0.3320 0.04855 0.03833 -0.0194 0.7368 1.0000
4.250 0.3421 0.04948 0.03926 -0.0176 0.7214 1.0000
4.500 0.3611 0.05040 0.04017 -0.0167 0.7050 1.0000
4.750 0.3824 0.05129 0.04109 -0.0161 0.6881 1.0000
5.000 0.4042 0.05218 0.04200 -0.0154 0.6710 1.0000
5.250 0.4281 0.05298 0.04285 -0.0149 0.6534 1.0000
5.500 0.4535 0.05369 0.04362 -0.0143 0.6356 1.0000
5.750 0.4823 0.05420 0.04420 -0.0138 0.6176 1.0000
6.000 0.5109 0.05463 0.04471 -0.0133 0.6000 1.0000
6.250 0.5434 0.05473 0.04490 -0.0127 0.5824 1.0000
6.500 0.5751 0.05474 0.04503 -0.0119 0.5652 1.0000
6.750 0.5863 0.05569 0.04602 -0.0099 0.5464 1.0000
7.750 0.6533 0.05856 0.04918 -0.0034 0.4772 1.0000
8.000 0.6766 0.05865 0.04937 -0.0019 0.4612 1.0000
8.250 0.8029 0.04905 0.04019 -0.0013 0.4483 1.0000
8.500 0.8304 0.04837 0.03962 0.0005 0.4314 1.0000
8.750 0.8590 0.04764 0.03899 0.0022 0.4141 1.0000
9.000 0.8993 0.04619 0.03766 0.0032 0.3954 1.0000
9.250 0.9964 0.04132 0.03277 0.0003 0.3685 1.0000
9.500 1.0121 0.04211 0.03359 0.0024 0.3507 1.0000
9.750 1.0285 0.04291 0.03440 0.0043 0.3326 1.0000
10.000 1.0472 0.04367 0.03513 0.0060 0.3137 1.0000
10.250 1.0865 0.04356 0.03479 0.0060 0.2895 1.0000
10.500 1.0931 0.04499 0.03624 0.0086 0.2725 1.0000
10.750 1.0997 0.04653 0.03779 0.0111 0.2559 1.0000
11.000 1.1055 0.04816 0.03942 0.0135 0.2398 1.0000
11.250 1.1114 0.04994 0.04121 0.0159 0.2245 1.0000
11.500 1.1174 0.05182 0.04307 0.0181 0.2097 1.0000
11.750 1.0900 0.05482 0.04633 0.0228 0.2051 1.0000
12.000 1.0938 0.05703 0.04853 0.0248 0.1930 1.0000
12.250 1.0608 0.06110 0.05284 0.0278 0.1908 1.0000
12.500 1.0200 0.06661 0.05857 0.0292 0.1909 1.0000
12.750 0.9688 0.07448 0.06660 0.0284 0.1935 1.0000
13.000 0.9157 0.08458 0.07674 0.0254 0.1958 1.0000
|
Polar data table (+)
Polar graphs
<< Back to KC-135 BL52.44 AIRFOIL (kc135a-il)