KC-135 BL200.76 AIRFOIL (kc135c-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: KC-135 BL200.76 AIRFOIL (kc135c-il) Reynolds number: 100,000 Max Cl/Cd: 50.23 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-kc135c-il-100000-n5.txt Download as CSV file: xf-kc135c-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4922 0.08990 0.08493 -0.0398 1.0000 0.0494
-9.000 -0.5081 0.08656 0.08161 -0.0417 1.0000 0.0496
-8.750 -0.5213 0.08381 0.07881 -0.0414 1.0000 0.0497
-8.500 -0.5319 0.08110 0.07601 -0.0409 1.0000 0.0498
-8.250 -0.5384 0.07841 0.07319 -0.0400 1.0000 0.0499
-7.750 -0.5460 0.07303 0.06751 -0.0366 1.0000 0.0501
-7.500 -0.5427 0.06629 0.06093 -0.0362 1.0000 0.0510
-7.250 -0.5397 0.06248 0.05716 -0.0347 1.0000 0.0517
-7.000 -0.5388 0.05950 0.05416 -0.0325 1.0000 0.0523
-6.750 -0.5401 0.05689 0.05148 -0.0297 1.0000 0.0529
-6.500 -0.5370 0.05320 0.04747 -0.0257 1.0000 0.0357
-6.250 -0.5150 0.04851 0.04252 -0.0277 0.9913 0.0333
-6.000 -0.4895 0.04487 0.03850 -0.0292 0.9826 0.0337
-5.750 -0.4627 0.04154 0.03476 -0.0304 0.9745 0.0344
-5.500 -0.4356 0.03810 0.03089 -0.0313 0.9668 0.0334
-5.250 -0.4084 0.03513 0.02751 -0.0317 0.9587 0.0324
-5.000 -0.3769 0.03228 0.02422 -0.0327 0.9531 0.0319
-4.750 -0.3482 0.02995 0.02149 -0.0329 0.9453 0.0321
-4.500 -0.3157 0.02813 0.01931 -0.0338 0.9394 0.0341
-4.250 -0.2852 0.02650 0.01730 -0.0340 0.9323 0.0355
-4.000 -0.2526 0.02479 0.01530 -0.0347 0.9263 0.0359
-3.750 -0.2197 0.02332 0.01360 -0.0354 0.9205 0.0366
-3.500 -0.1888 0.02215 0.01225 -0.0357 0.9134 0.0375
-3.250 -0.1565 0.02059 0.01074 -0.0369 0.9087 0.0407
-3.000 -0.1302 0.01976 0.00990 -0.0366 0.9003 0.0440
-2.750 -0.1007 0.01884 0.00895 -0.0368 0.8946 0.0465
-2.500 -0.0768 0.01824 0.00827 -0.0360 0.8863 0.0494
-2.250 -0.0512 0.01752 0.00752 -0.0358 0.8800 0.0558
-2.000 -0.0276 0.01709 0.00696 -0.0350 0.8720 0.0618
-1.750 -0.0010 0.01660 0.00639 -0.0347 0.8658 0.0741
-1.500 0.0227 0.01614 0.00595 -0.0340 0.8583 0.0984
-1.250 0.0279 0.01385 0.00597 -0.0297 0.8515 0.6908
-1.000 0.0501 0.01379 0.00614 -0.0274 0.8451 0.7797
-0.750 0.0775 0.01390 0.00637 -0.0259 0.8392 0.8417
-0.500 0.1150 0.01412 0.00657 -0.0263 0.8354 0.8866
-0.250 0.1481 0.01420 0.00657 -0.0274 0.8289 0.8996
0.000 0.1805 0.01422 0.00649 -0.0284 0.8230 0.9092
0.250 0.2165 0.01424 0.00643 -0.0301 0.8178 0.9172
0.500 0.2500 0.01430 0.00645 -0.0315 0.8109 0.9264
0.750 0.2849 0.01432 0.00640 -0.0330 0.8057 0.9358
1.000 0.3212 0.01439 0.00648 -0.0350 0.7983 0.9438
1.250 0.3568 0.01439 0.00644 -0.0366 0.7913 0.9529
1.500 0.3930 0.01437 0.00641 -0.0385 0.7800 0.9611
1.750 0.4281 0.01423 0.00625 -0.0397 0.7647 0.9697
2.000 0.4630 0.01408 0.00606 -0.0410 0.7484 0.9789
2.250 0.4991 0.01402 0.00602 -0.0428 0.7354 0.9881
2.500 0.5363 0.01399 0.00605 -0.0450 0.7243 0.9971
2.750 0.5616 0.01400 0.00608 -0.0447 0.7138 1.0000
3.000 0.5814 0.01402 0.00610 -0.0432 0.7028 1.0000
3.250 0.6004 0.01409 0.00622 -0.0416 0.6899 1.0000
3.500 0.6201 0.01416 0.00636 -0.0400 0.6760 1.0000
3.750 0.6401 0.01422 0.00647 -0.0385 0.6606 1.0000
4.000 0.6606 0.01429 0.00658 -0.0369 0.6435 1.0000
4.250 0.6801 0.01438 0.00674 -0.0352 0.6221 1.0000
4.500 0.7001 0.01445 0.00687 -0.0335 0.5979 1.0000
4.750 0.7195 0.01457 0.00701 -0.0317 0.5663 1.0000
5.000 0.7383 0.01474 0.00713 -0.0298 0.5246 1.0000
5.250 0.7554 0.01504 0.00725 -0.0275 0.4685 1.0000
5.500 0.7693 0.01563 0.00749 -0.0248 0.4030 1.0000
5.750 0.7810 0.01647 0.00797 -0.0221 0.3406 1.0000
6.000 0.7930 0.01736 0.00852 -0.0196 0.2861 1.0000
6.250 0.8071 0.01815 0.00910 -0.0176 0.2438 1.0000
6.500 0.8232 0.01885 0.00970 -0.0159 0.2127 1.0000
6.750 0.8391 0.01958 0.01036 -0.0141 0.1778 1.0000
7.000 0.8542 0.02042 0.01103 -0.0124 0.1293 1.0000
7.250 0.8658 0.02165 0.01193 -0.0102 0.0848 1.0000
7.500 0.8784 0.02287 0.01301 -0.0081 0.0638 1.0000
7.750 0.8915 0.02403 0.01416 -0.0060 0.0545 1.0000
8.000 0.9035 0.02524 0.01539 -0.0039 0.0487 1.0000
8.250 0.9162 0.02638 0.01666 -0.0018 0.0435 1.0000
8.500 0.9280 0.02757 0.01795 0.0002 0.0401 1.0000
8.750 0.9359 0.02915 0.01955 0.0026 0.0373 1.0000
9.000 0.9492 0.03034 0.02092 0.0045 0.0345 1.0000
9.250 0.9613 0.03169 0.02240 0.0064 0.0321 1.0000
9.500 0.9736 0.03324 0.02404 0.0082 0.0305 1.0000
9.750 0.9870 0.03499 0.02590 0.0097 0.0292 1.0000
10.000 1.0029 0.03746 0.02844 0.0105 0.0276 1.0000
10.250 1.0144 0.03914 0.03043 0.0123 0.0263 1.0000
10.500 1.0246 0.04110 0.03267 0.0140 0.0251 1.0000
10.750 1.0318 0.04328 0.03511 0.0158 0.0240 1.0000
11.000 1.0369 0.04572 0.03782 0.0176 0.0235 1.0000
11.250 1.0379 0.04828 0.04065 0.0196 0.0230 1.0000
11.500 1.0353 0.05103 0.04366 0.0214 0.0226 1.0000
11.750 1.0297 0.05399 0.04688 0.0229 0.0224 1.0000
12.000 1.0212 0.05724 0.05042 0.0241 0.0222 1.0000
12.250 1.0100 0.06082 0.05425 0.0248 0.0221 1.0000
12.500 0.9961 0.06484 0.05851 0.0249 0.0220 1.0000
12.750 0.9795 0.06944 0.06335 0.0241 0.0220 1.0000
13.000 0.9603 0.07473 0.06886 0.0224 0.0220 1.0000
13.250 0.9386 0.08090 0.07525 0.0196 0.0222 1.0000
13.500 0.9094 0.08919 0.08375 0.0147 0.0225 1.0000
13.750 0.8754 0.10012 0.09487 0.0073 0.0231 1.0000
14.000 0.8338 0.11582 0.11068 -0.0031 0.0238 1.0000
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Polar data table (+)
Polar graphs
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