KC-135 BL200.76 AIRFOIL (kc135c-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: KC-135 BL200.76 AIRFOIL (kc135c-il) Reynolds number: 500,000 Max Cl/Cd: 91.65 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-kc135c-il-500000.txt Download as CSV file: xf-kc135c-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: KC-135 BL200.76 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4802 0.08286 0.08057 -0.0415 1.0000 0.0164
-9.250 -0.4891 0.07843 0.07611 -0.0445 1.0000 0.0164
-9.000 -0.5019 0.07486 0.07251 -0.0453 1.0000 0.0164
-8.750 -0.5144 0.07215 0.06976 -0.0437 1.0000 0.0164
-8.500 -0.5333 0.06709 0.06465 -0.0424 1.0000 0.0166
-8.250 -0.5426 0.06319 0.06072 -0.0409 1.0000 0.0168
-8.000 -0.5460 0.06032 0.05780 -0.0392 1.0000 0.0169
-7.750 -0.5488 0.05776 0.05519 -0.0369 1.0000 0.0171
-7.250 -0.5294 0.05178 0.04906 -0.0375 0.9939 0.0177
-7.000 -0.5064 0.04803 0.04515 -0.0400 0.9868 0.0184
-6.750 -0.4815 0.04430 0.04121 -0.0421 0.9797 0.0196
-6.500 -0.4448 0.04223 0.03865 -0.0428 0.9729 0.0219
-6.250 -0.4284 0.03548 0.03148 -0.0437 0.9625 0.0225
-6.000 -0.4021 0.03229 0.02818 -0.0456 0.9544 0.0232
-5.750 -0.3763 0.03027 0.02605 -0.0466 0.9431 0.0242
-5.500 -0.3519 0.02848 0.02406 -0.0466 0.9300 0.0257
-5.250 -0.3216 0.02916 0.02442 -0.0455 0.9173 0.0294
-5.000 -0.3000 0.02788 0.02282 -0.0440 0.9043 0.0296
-4.750 -0.2866 0.02248 0.01709 -0.0425 0.8921 0.0314
-4.500 -0.2638 0.02102 0.01552 -0.0418 0.8817 0.0326
-4.250 -0.2404 0.01984 0.01418 -0.0409 0.8712 0.0346
-4.000 -0.2130 0.01646 0.01034 -0.0391 0.8628 0.0259
-3.750 -0.1868 0.01526 0.00897 -0.0384 0.8547 0.0264
-3.500 -0.1607 0.01377 0.00732 -0.0377 0.8463 0.0259
-3.250 -0.1353 0.01284 0.00628 -0.0370 0.8390 0.0264
-3.000 -0.1109 0.01212 0.00551 -0.0362 0.8309 0.0272
-2.750 -0.0863 0.01164 0.00493 -0.0355 0.8244 0.0284
-2.500 -0.0620 0.01122 0.00447 -0.0348 0.8168 0.0301
-2.250 -0.0407 0.01046 0.00363 -0.0335 0.8104 0.0325
-2.000 -0.0163 0.01013 0.00329 -0.0328 0.8034 0.0353
-1.750 0.0090 0.00988 0.00299 -0.0323 0.7974 0.0388
-1.500 0.0335 0.00953 0.00260 -0.0317 0.7911 0.0456
-1.250 0.0595 0.00936 0.00239 -0.0313 0.7850 0.0529
-1.000 0.0837 0.00895 0.00210 -0.0306 0.7793 0.1067
-0.750 0.0869 0.00689 0.00192 -0.0264 0.7721 0.6800
-0.500 0.1105 0.00677 0.00191 -0.0253 0.7658 0.7296
-0.250 0.1352 0.00669 0.00190 -0.0245 0.7578 0.7596
0.000 0.1607 0.00665 0.00188 -0.0238 0.7507 0.7821
0.250 0.1858 0.00658 0.00186 -0.0231 0.7418 0.8042
0.500 0.2112 0.00654 0.00185 -0.0224 0.7327 0.8250
0.750 0.2361 0.00651 0.00184 -0.0215 0.7235 0.8500
1.000 0.2623 0.00648 0.00186 -0.0211 0.7149 0.8666
1.250 0.2897 0.00649 0.00188 -0.0209 0.7091 0.8826
1.500 0.3177 0.00648 0.00193 -0.0209 0.7022 0.8995
1.750 0.3473 0.00652 0.00197 -0.0212 0.6952 0.9164
2.000 0.3792 0.00656 0.00205 -0.0220 0.6874 0.9319
2.250 0.4130 0.00664 0.00211 -0.0233 0.6798 0.9447
2.500 0.4468 0.00668 0.00219 -0.0246 0.6704 0.9562
2.750 0.4847 0.00675 0.00226 -0.0269 0.6607 0.9627
3.000 0.5193 0.00683 0.00233 -0.0284 0.6489 0.9709
3.250 0.5581 0.00690 0.00240 -0.0309 0.6348 0.9752
3.500 0.5937 0.00698 0.00248 -0.0327 0.6184 0.9815
3.750 0.6305 0.00707 0.00256 -0.0348 0.5961 0.9862
4.000 0.6663 0.00727 0.00261 -0.0367 0.5506 0.9914
4.250 0.7001 0.00770 0.00273 -0.0384 0.4696 0.9966
4.500 0.7263 0.00852 0.00306 -0.0389 0.3629 1.0000
4.750 0.7406 0.00907 0.00334 -0.0367 0.3061 1.0000
5.000 0.7572 0.00951 0.00360 -0.0348 0.2661 1.0000
5.250 0.7751 0.00991 0.00387 -0.0331 0.2371 1.0000
5.500 0.7939 0.01025 0.00413 -0.0316 0.2148 1.0000
5.750 0.8126 0.01063 0.00443 -0.0300 0.1914 1.0000
6.000 0.8311 0.01104 0.00472 -0.0285 0.1581 1.0000
6.250 0.8445 0.01185 0.00517 -0.0262 0.0968 1.0000
6.500 0.8559 0.01285 0.00585 -0.0235 0.0473 1.0000
6.750 0.8722 0.01351 0.00648 -0.0215 0.0376 1.0000
7.000 0.8909 0.01398 0.00701 -0.0199 0.0335 1.0000
7.250 0.9047 0.01484 0.00789 -0.0176 0.0286 1.0000
7.500 0.9240 0.01527 0.00838 -0.0161 0.0265 1.0000
7.750 0.9413 0.01585 0.00901 -0.0144 0.0242 1.0000
8.000 0.9542 0.01678 0.00998 -0.0121 0.0220 1.0000
8.250 0.9657 0.01784 0.01114 -0.0095 0.0208 1.0000
8.500 0.9821 0.01851 0.01189 -0.0077 0.0198 1.0000
8.750 0.9982 0.01919 0.01263 -0.0059 0.0185 1.0000
9.000 1.0147 0.01978 0.01324 -0.0044 0.0172 1.0000
9.250 1.0268 0.02075 0.01427 -0.0021 0.0165 1.0000
9.500 1.0320 0.02277 0.01639 0.0010 0.0156 1.0000
9.750 1.0453 0.02408 0.01781 0.0030 0.0153 1.0000
10.000 1.0588 0.02490 0.01874 0.0050 0.0149 1.0000
10.250 1.0720 0.02612 0.02009 0.0069 0.0145 1.0000
10.500 1.0847 0.02753 0.02165 0.0087 0.0141 1.0000
10.750 1.0966 0.02909 0.02336 0.0105 0.0138 1.0000
11.000 1.1061 0.03054 0.02495 0.0125 0.0134 1.0000
11.250 1.1142 0.03198 0.02652 0.0146 0.0130 1.0000
11.500 1.1209 0.03331 0.02798 0.0165 0.0126 1.0000
11.750 1.1259 0.03440 0.02914 0.0184 0.0121 1.0000
12.000 1.1293 0.03632 0.03121 0.0202 0.0119 1.0000
12.250 1.1296 0.03869 0.03371 0.0220 0.0117 1.0000
12.500 1.1253 0.04168 0.03690 0.0238 0.0115 1.0000
12.750 1.1187 0.04472 0.04015 0.0253 0.0115 1.0000
13.000 1.1101 0.04790 0.04354 0.0266 0.0115 1.0000
13.250 1.0989 0.05145 0.04729 0.0273 0.0115 1.0000
13.500 1.0804 0.05614 0.05220 0.0274 0.0114 1.0000
13.750 1.0601 0.06124 0.05752 0.0268 0.0113 1.0000
14.000 1.0516 0.06493 0.06135 0.0258 0.0114 1.0000
14.250 1.0338 0.07029 0.06691 0.0238 0.0115 1.0000
14.500 1.0172 0.07595 0.07274 0.0211 0.0115 1.0000
14.750 0.9983 0.08262 0.07958 0.0174 0.0116 1.0000
15.000 0.9745 0.09123 0.08839 0.0120 0.0117 1.0000
15.250 0.9540 0.09996 0.09726 0.0063 0.0117 1.0000
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Polar data table (+)
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