NASA/LANGLEY RC-10(B)3 AIRFOIL (rc10b3-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-10(B)3 AIRFOIL (rc10b3-il) Reynolds number: 200,000 Max Cl/Cd: 62.91 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10b3-il-200000.txt Download as CSV file: xf-rc10b3-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(B)3 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5184 0.09099 0.08744 -0.0206 1.0000 0.0521
-8.750 -0.5241 0.08581 0.08230 -0.0258 1.0000 0.0533
-8.500 -0.5384 0.08029 0.07676 -0.0315 1.0000 0.0544
-8.250 -0.5620 0.07576 0.07200 -0.0349 1.0000 0.0557
-8.000 -0.5726 0.07284 0.06885 -0.0339 1.0000 0.0559
-7.750 -0.5737 0.06617 0.06227 -0.0333 1.0000 0.0569
-7.500 -0.5677 0.06327 0.05943 -0.0315 1.0000 0.0577
-7.250 -0.5673 0.06108 0.05726 -0.0287 1.0000 0.0585
-7.000 -0.5705 0.05912 0.05528 -0.0253 1.0000 0.0595
-6.750 -0.5754 0.05718 0.05330 -0.0215 1.0000 0.0606
-6.500 -0.5798 0.05512 0.05116 -0.0178 1.0000 0.0622
-6.250 -0.5832 0.05294 0.04882 -0.0141 1.0000 0.0649
-5.750 -0.5872 0.04697 0.04231 -0.0072 0.9994 0.0702
-5.500 -0.5600 0.04413 0.03946 -0.0089 0.9960 0.0727
-5.250 -0.5345 0.04202 0.03661 -0.0093 0.9906 0.0825
-5.000 -0.5074 0.03832 0.03310 -0.0111 0.9873 0.0856
-4.750 -0.4773 0.03635 0.03068 -0.0124 0.9835 0.0977
-4.500 -0.4493 0.02619 0.01898 -0.0077 0.9795 0.0445
-4.250 -0.4176 0.02357 0.01603 -0.0084 0.9759 0.0433
-4.000 -0.3813 0.02164 0.01377 -0.0098 0.9732 0.0430
-3.750 -0.3424 0.02024 0.01212 -0.0118 0.9711 0.0440
-3.500 -0.3125 0.01919 0.01092 -0.0121 0.9656 0.0462
-3.250 -0.2772 0.01793 0.00970 -0.0137 0.9621 0.0488
-3.000 -0.2390 0.01706 0.00882 -0.0158 0.9592 0.0523
-2.750 -0.1996 0.01612 0.00788 -0.0182 0.9570 0.0581
-2.500 -0.1728 0.01567 0.00745 -0.0180 0.9498 0.0684
-2.250 -0.1383 0.01465 0.00666 -0.0195 0.9459 0.1036
-2.000 -0.1095 0.01302 0.00615 -0.0204 0.9425 0.3373
-1.750 0.0064 0.01261 0.00776 -0.0352 0.9590 0.9627
-1.500 0.0969 0.01269 0.00764 -0.0476 0.9649 0.9897
-1.250 0.1628 0.01238 0.00721 -0.0557 0.9645 1.0000
-1.000 0.2027 0.01217 0.00691 -0.0584 0.9577 1.0000
-0.750 0.2329 0.01204 0.00671 -0.0590 0.9466 1.0000
-0.500 0.2598 0.01193 0.00655 -0.0588 0.9351 1.0000
-0.250 0.2834 0.01185 0.00642 -0.0577 0.9239 1.0000
0.000 0.3050 0.01178 0.00629 -0.0562 0.9129 1.0000
0.250 0.3260 0.01172 0.00621 -0.0547 0.9007 1.0000
0.500 0.3469 0.01166 0.00611 -0.0531 0.8890 1.0000
0.750 0.3676 0.01157 0.00598 -0.0513 0.8778 1.0000
1.000 0.3884 0.01146 0.00583 -0.0494 0.8666 1.0000
1.250 0.4105 0.01139 0.00574 -0.0480 0.8540 1.0000
1.500 0.4328 0.01131 0.00566 -0.0467 0.8420 1.0000
1.750 0.4553 0.01123 0.00555 -0.0453 0.8301 1.0000
2.000 0.4778 0.01112 0.00542 -0.0438 0.8172 1.0000
2.250 0.5005 0.01102 0.00530 -0.0424 0.8022 1.0000
2.500 0.5238 0.01094 0.00520 -0.0411 0.7867 1.0000
2.750 0.5475 0.01088 0.00513 -0.0400 0.7706 1.0000
3.000 0.5713 0.01083 0.00508 -0.0390 0.7523 1.0000
3.250 0.5948 0.01077 0.00501 -0.0378 0.7309 1.0000
3.500 0.6185 0.01075 0.00494 -0.0366 0.7057 1.0000
3.750 0.6420 0.01076 0.00489 -0.0355 0.6764 1.0000
4.000 0.6652 0.01083 0.00485 -0.0342 0.6388 1.0000
4.250 0.6880 0.01100 0.00487 -0.0330 0.5928 1.0000
4.500 0.7103 0.01129 0.00494 -0.0319 0.5408 1.0000
4.750 0.7318 0.01170 0.00510 -0.0307 0.4839 1.0000
5.000 0.7527 0.01220 0.00536 -0.0295 0.4274 1.0000
5.250 0.7729 0.01275 0.00567 -0.0283 0.3735 1.0000
5.500 0.7926 0.01334 0.00603 -0.0270 0.3216 1.0000
5.750 0.8113 0.01400 0.00643 -0.0257 0.2685 1.0000
6.000 0.8290 0.01476 0.00693 -0.0242 0.2119 1.0000
6.250 0.8448 0.01572 0.00755 -0.0225 0.1542 1.0000
6.500 0.8601 0.01675 0.00832 -0.0205 0.1161 1.0000
6.750 0.8754 0.01769 0.00910 -0.0185 0.0977 1.0000
7.000 0.8919 0.01849 0.00989 -0.0166 0.0872 1.0000
7.250 0.9070 0.01943 0.01082 -0.0145 0.0801 1.0000
7.500 0.9237 0.02015 0.01156 -0.0126 0.0742 1.0000
7.750 0.9380 0.02132 0.01268 -0.0105 0.0698 1.0000
8.000 0.9556 0.02210 0.01355 -0.0087 0.0664 1.0000
8.250 0.9726 0.02293 0.01443 -0.0069 0.0630 1.0000
8.500 0.9894 0.02437 0.01579 -0.0053 0.0595 1.0000
8.750 1.0078 0.02541 0.01697 -0.0038 0.0574 1.0000
9.000 1.0264 0.02649 0.01817 -0.0022 0.0553 1.0000
9.250 1.0442 0.02759 0.01940 -0.0007 0.0532 1.0000
9.500 1.0615 0.02870 0.02054 0.0008 0.0510 1.0000
9.750 1.0819 0.03154 0.02343 0.0013 0.0486 1.0000
10.000 1.0957 0.03269 0.02484 0.0036 0.0477 1.0000
10.250 1.1088 0.03427 0.02669 0.0058 0.0466 1.0000
10.500 1.1197 0.03595 0.02865 0.0082 0.0453 1.0000
10.750 1.1294 0.03756 0.03047 0.0106 0.0439 1.0000
11.000 1.1395 0.03909 0.03214 0.0129 0.0426 1.0000
11.250 1.1468 0.04096 0.03418 0.0153 0.0417 1.0000
11.500 1.1520 0.04316 0.03654 0.0178 0.0409 1.0000
11.750 1.1538 0.04607 0.03963 0.0204 0.0403 1.0000
12.000 1.1466 0.04998 0.04380 0.0237 0.0398 1.0000
12.500 1.1113 0.05592 0.05024 0.0325 0.0394 1.0000
12.750 1.0921 0.05849 0.05305 0.0359 0.0392 1.0000
13.000 1.0719 0.06178 0.05657 0.0381 0.0392 1.0000
13.250 1.0501 0.06575 0.06075 0.0392 0.0391 1.0000
13.500 1.0271 0.07047 0.06568 0.0392 0.0392 1.0000
13.750 1.0033 0.07610 0.07147 0.0380 0.0393 1.0000
14.000 0.8402 0.11063 0.10672 0.0139 0.0470 1.0000
14.250 0.8257 0.11843 0.11449 0.0111 0.0482 1.0000
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Polar data table (+)
Polar graphs
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