NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Reynolds number: 100,000 Max Cl/Cd: 41.79 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc1264c-il-100000-n5.txt Download as CSV file: xf-rc1264c-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC12-64C 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4847 0.11567 0.11017 -0.0263 1.0000 0.0845
-10.250 -0.4865 0.11181 0.10636 -0.0283 1.0000 0.0873
-10.000 -0.5225 0.10499 0.09967 -0.0379 1.0000 0.0911
-9.750 -0.5081 0.10182 0.09652 -0.0356 1.0000 0.0920
-9.500 -0.5012 0.09849 0.09322 -0.0351 1.0000 0.0931
-9.250 -0.4993 0.09487 0.08963 -0.0357 1.0000 0.0943
-9.000 -0.5020 0.09086 0.08567 -0.0372 1.0000 0.0959
-8.750 -0.5130 0.08615 0.08102 -0.0401 1.0000 0.0979
-8.500 -0.6324 0.08584 0.08044 -0.0375 1.0000 0.1038
-8.250 -0.6042 0.07901 0.07384 -0.0382 1.0000 0.1051
-7.750 -0.6050 0.07278 0.06767 -0.0339 1.0000 0.1069
-7.250 -0.6516 0.05739 0.05140 -0.0256 1.0000 0.0508
-7.000 -0.6723 0.05145 0.04472 -0.0191 1.0000 0.0421
-6.750 -0.6654 0.04919 0.04240 -0.0165 1.0000 0.0407
-6.500 -0.6620 0.04663 0.03965 -0.0132 1.0000 0.0396
-6.250 -0.6638 0.04313 0.03538 -0.0082 1.0000 0.0369
-6.000 -0.6534 0.04103 0.03315 -0.0058 1.0000 0.0365
-5.750 -0.6270 0.03834 0.03018 -0.0065 0.9961 0.0357
-5.500 -0.6003 0.03551 0.02683 -0.0065 0.9922 0.0338
-5.250 -0.5709 0.03330 0.02414 -0.0068 0.9881 0.0323
-5.000 -0.5391 0.03134 0.02191 -0.0077 0.9849 0.0312
-4.750 -0.5092 0.02965 0.01978 -0.0078 0.9804 0.0298
-4.500 -0.4760 0.02828 0.01819 -0.0088 0.9767 0.0291
-4.250 -0.4424 0.02699 0.01677 -0.0100 0.9735 0.0285
-4.000 -0.4120 0.02588 0.01557 -0.0104 0.9686 0.0280
-3.750 -0.3782 0.02488 0.01448 -0.0116 0.9647 0.0275
-3.500 -0.3438 0.02401 0.01355 -0.0129 0.9614 0.0270
-3.250 -0.3162 0.02331 0.01279 -0.0128 0.9551 0.0267
-3.000 -0.2836 0.02263 0.01209 -0.0139 0.9508 0.0265
-2.750 -0.2517 0.02195 0.01142 -0.0149 0.9464 0.0264
-2.500 -0.2259 0.02141 0.01087 -0.0146 0.9393 0.0264
-2.250 -0.1932 0.02081 0.01026 -0.0158 0.9348 0.0270
-2.000 -0.1629 0.02034 0.00974 -0.0164 0.9293 0.0280
-1.750 -0.1348 0.01997 0.00933 -0.0165 0.9223 0.0288
-1.500 -0.0992 0.01961 0.00892 -0.0180 0.9181 0.0292
-1.250 -0.0704 0.01935 0.00862 -0.0182 0.9113 0.0295
-1.000 -0.0400 0.01908 0.00832 -0.0186 0.9047 0.0301
-0.750 -0.0041 0.01880 0.00801 -0.0201 0.9005 0.0312
-0.500 0.0207 0.01866 0.00786 -0.0194 0.8915 0.0326
-0.250 0.0497 0.01761 0.00760 -0.0199 0.8860 0.2190
0.000 0.0678 0.01562 0.00787 -0.0176 0.8804 0.7415
0.250 0.1067 0.01573 0.00822 -0.0187 0.8740 0.8281
0.500 0.1481 0.01582 0.00834 -0.0206 0.8692 0.8685
0.750 0.1804 0.01592 0.00842 -0.0211 0.8608 0.8861
1.000 0.2175 0.01594 0.00843 -0.0225 0.8539 0.9025
1.250 0.2597 0.01612 0.00863 -0.0247 0.8465 0.9277
1.500 0.3062 0.01622 0.00874 -0.0278 0.8383 0.9455
1.750 0.3590 0.01625 0.00879 -0.0324 0.8287 0.9608
2.000 0.4401 0.01608 0.00866 -0.0423 0.8161 0.9836
2.250 0.5067 0.01562 0.00818 -0.0495 0.7888 0.9985
2.500 0.5329 0.01540 0.00789 -0.0488 0.7602 1.0000
2.750 0.5543 0.01525 0.00765 -0.0470 0.7290 1.0000
3.000 0.5753 0.01517 0.00745 -0.0452 0.6931 1.0000
3.250 0.5963 0.01517 0.00731 -0.0435 0.6555 1.0000
3.500 0.6169 0.01524 0.00725 -0.0419 0.6164 1.0000
3.750 0.6365 0.01540 0.00723 -0.0400 0.5684 1.0000
4.000 0.6548 0.01567 0.00726 -0.0380 0.5103 1.0000
4.250 0.6710 0.01611 0.00736 -0.0358 0.4425 1.0000
4.500 0.6846 0.01676 0.00759 -0.0332 0.3647 1.0000
4.750 0.6970 0.01753 0.00795 -0.0306 0.2922 1.0000
5.000 0.7100 0.01828 0.00837 -0.0282 0.2345 1.0000
5.250 0.7234 0.01898 0.00883 -0.0258 0.1924 1.0000
5.500 0.7369 0.01967 0.00933 -0.0234 0.1625 1.0000
5.750 0.7508 0.02030 0.00987 -0.0211 0.1413 1.0000
6.000 0.7645 0.02093 0.01041 -0.0186 0.1267 1.0000
6.250 0.7777 0.02157 0.01097 -0.0161 0.1163 1.0000
6.500 0.7914 0.02216 0.01155 -0.0137 0.1081 1.0000
6.750 0.8045 0.02281 0.01218 -0.0112 0.1016 1.0000
7.000 0.8183 0.02343 0.01281 -0.0088 0.0956 1.0000
7.250 0.8301 0.02420 0.01350 -0.0062 0.0910 1.0000
7.500 0.8450 0.02482 0.01422 -0.0039 0.0861 1.0000
7.750 0.8589 0.02552 0.01494 -0.0017 0.0818 1.0000
8.000 0.8714 0.02645 0.01580 0.0007 0.0783 1.0000
8.250 0.8873 0.02716 0.01665 0.0027 0.0743 1.0000
8.500 0.9020 0.02791 0.01745 0.0047 0.0702 1.0000
8.750 0.9159 0.02880 0.01831 0.0067 0.0671 1.0000
9.000 0.9323 0.02979 0.01941 0.0085 0.0640 1.0000
9.250 0.9477 0.03067 0.02040 0.0103 0.0607 1.0000
9.500 0.9614 0.03154 0.02127 0.0121 0.0580 1.0000
9.750 0.9765 0.03269 0.02249 0.0138 0.0554 1.0000
10.000 0.9927 0.03396 0.02394 0.0153 0.0531 1.0000
10.250 1.0076 0.03521 0.02530 0.0169 0.0511 1.0000
10.500 1.0211 0.03639 0.02650 0.0186 0.0494 1.0000
10.750 1.0348 0.03789 0.02804 0.0201 0.0477 1.0000
11.000 1.0427 0.03940 0.02988 0.0226 0.0457 1.0000
11.250 1.0496 0.04089 0.03160 0.0250 0.0439 1.0000
11.500 1.0563 0.04235 0.03321 0.0272 0.0426 1.0000
11.750 1.0631 0.04391 0.03490 0.0292 0.0416 1.0000
12.000 1.0701 0.04559 0.03665 0.0310 0.0408 1.0000
12.250 1.0727 0.04779 0.03904 0.0330 0.0402 1.0000
12.500 1.0666 0.05056 0.04218 0.0355 0.0395 1.0000
12.750 1.0575 0.05361 0.04557 0.0377 0.0390 1.0000
13.000 1.0452 0.05702 0.04928 0.0394 0.0386 1.0000
13.250 1.0296 0.06087 0.05343 0.0405 0.0383 1.0000
13.500 1.0103 0.06535 0.05819 0.0408 0.0381 1.0000
13.750 0.9868 0.07070 0.06382 0.0399 0.0380 1.0000
14.000 0.9580 0.07738 0.07076 0.0374 0.0380 1.0000
14.250 0.9212 0.08629 0.07992 0.0326 0.0382 1.0000
14.500 0.8653 0.10097 0.09486 0.0230 0.0387 1.0000
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Polar data table (+)
Polar graphs
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