NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Reynolds number: 200,000 Max Cl/Cd: 58.11 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc1264c-il-200000.txt Download as CSV file: xf-rc1264c-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.3748 0.10185 0.09821 -0.0362 1.0000 0.0629
-10.250 -0.4031 0.09498 0.09138 -0.0419 1.0000 0.0646
-10.000 -0.4338 0.08700 0.08343 -0.0482 1.0000 0.0648
-9.750 -0.3918 0.08895 0.08541 -0.0396 1.0000 0.0675
-9.500 -0.3931 0.08549 0.08197 -0.0398 1.0000 0.0696
-9.250 -0.4080 0.08014 0.07666 -0.0422 1.0000 0.0717
-9.000 -0.4334 0.07327 0.06982 -0.0463 1.0000 0.0723
-8.750 -0.4602 0.06728 0.06384 -0.0487 1.0000 0.0721
-8.500 -0.4874 0.06299 0.05954 -0.0481 1.0000 0.0718
-8.250 -0.5163 0.05996 0.05650 -0.0447 1.0000 0.0715
-8.000 -0.5466 0.05788 0.05443 -0.0390 1.0000 0.0711
-7.750 -0.5750 0.05554 0.05204 -0.0338 1.0000 0.0714
-7.500 -0.6567 0.06178 0.05778 -0.0271 1.0000 0.0677
-7.250 -0.6607 0.05994 0.05588 -0.0236 1.0000 0.0700
-7.000 -0.6944 0.05750 0.05286 -0.0163 1.0000 0.0745
-6.750 -0.6795 0.05478 0.05039 -0.0155 1.0000 0.0761
-6.500 -0.6731 0.05325 0.04888 -0.0129 1.0000 0.0791
-6.250 -0.6836 0.05081 0.04604 -0.0083 1.0000 0.0855
-6.000 -0.6715 0.04900 0.04438 -0.0067 1.0000 0.0882
-5.750 -0.6744 0.04733 0.04231 -0.0027 1.0000 0.0970
-5.500 -0.6484 0.04512 0.04029 -0.0038 0.9980 0.1023
-5.250 -0.6217 0.04262 0.03764 -0.0054 0.9944 0.1143
-5.000 -0.5964 0.04035 0.03522 -0.0065 0.9899 0.1279
-4.750 -0.5678 0.03822 0.03296 -0.0079 0.9861 0.1432
-4.500 -0.5419 0.03625 0.03090 -0.0087 0.9814 0.1595
-4.250 -0.5136 0.03425 0.02883 -0.0097 0.9770 0.1770
-4.000 -0.4623 0.02549 0.01771 -0.0061 0.9756 0.0665
-3.750 -0.4192 0.02345 0.01505 -0.0068 0.9739 0.0505
-3.500 -0.3870 0.02161 0.01317 -0.0072 0.9697 0.0464
-3.250 -0.3505 0.02095 0.01231 -0.0085 0.9653 0.0450
-3.000 -0.3123 0.01977 0.01124 -0.0105 0.9627 0.0442
-2.750 -0.2729 0.01882 0.01034 -0.0126 0.9605 0.0422
-2.500 -0.2475 0.01824 0.00975 -0.0120 0.9527 0.0407
-2.250 -0.2105 0.01770 0.00921 -0.0139 0.9487 0.0398
-2.000 -0.1707 0.01710 0.00864 -0.0164 0.9460 0.0392
-1.750 -0.1464 0.01677 0.00829 -0.0157 0.9374 0.0388
-1.500 -0.1080 0.01640 0.00788 -0.0177 0.9333 0.0386
-1.250 -0.0658 0.01606 0.00751 -0.0206 0.9306 0.0390
-1.000 -0.0392 0.01589 0.00731 -0.0202 0.9223 0.0398
-0.750 -0.0011 0.01558 0.00703 -0.0222 0.9177 0.0432
-0.500 0.0127 0.01274 0.00659 -0.0204 0.9137 0.6124
-0.250 0.0315 0.01234 0.00690 -0.0175 0.9050 0.7826
0.000 0.0700 0.01238 0.00715 -0.0184 0.9008 0.8621
0.250 0.1129 0.01240 0.00718 -0.0205 0.8975 0.8895
0.500 0.1456 0.01252 0.00729 -0.0208 0.8901 0.9084
0.750 0.1854 0.01254 0.00730 -0.0225 0.8838 0.9200
1.000 0.2320 0.01256 0.00729 -0.0253 0.8796 0.9333
1.250 0.2810 0.01272 0.00746 -0.0287 0.8707 0.9474
1.500 0.3534 0.01271 0.00743 -0.0364 0.8632 0.9626
1.750 0.4289 0.01252 0.00722 -0.0449 0.8473 0.9810
2.000 0.5126 0.01194 0.00659 -0.0554 0.8248 1.0000
2.250 0.5332 0.01177 0.00637 -0.0537 0.8018 1.0000
2.500 0.5547 0.01162 0.00616 -0.0522 0.7790 1.0000
2.750 0.5767 0.01153 0.00598 -0.0509 0.7565 1.0000
3.000 0.5985 0.01148 0.00589 -0.0497 0.7318 1.0000
3.250 0.6205 0.01146 0.00579 -0.0484 0.7054 1.0000
3.500 0.6420 0.01148 0.00574 -0.0471 0.6736 1.0000
3.750 0.6630 0.01157 0.00568 -0.0456 0.6349 1.0000
4.000 0.6828 0.01175 0.00567 -0.0439 0.5856 1.0000
4.250 0.7010 0.01207 0.00572 -0.0421 0.5214 1.0000
4.500 0.7159 0.01265 0.00588 -0.0397 0.4329 1.0000
4.750 0.7267 0.01360 0.00621 -0.0369 0.3200 1.0000
5.000 0.7363 0.01473 0.00672 -0.0340 0.2159 1.0000
5.250 0.7477 0.01570 0.00729 -0.0314 0.1590 1.0000
5.500 0.7617 0.01639 0.00781 -0.0290 0.1353 1.0000
5.750 0.7760 0.01699 0.00833 -0.0266 0.1225 1.0000
6.000 0.7900 0.01757 0.00883 -0.0242 0.1136 1.0000
6.250 0.8051 0.01808 0.00934 -0.0219 0.1069 1.0000
6.500 0.8188 0.01866 0.00987 -0.0195 0.1010 1.0000
6.750 0.8330 0.01929 0.01052 -0.0170 0.0960 1.0000
7.000 0.8478 0.01981 0.01104 -0.0147 0.0911 1.0000
7.250 0.8612 0.02074 0.01187 -0.0123 0.0864 1.0000
7.500 0.8769 0.02121 0.01244 -0.0100 0.0822 1.0000
7.750 0.8919 0.02184 0.01304 -0.0079 0.0780 1.0000
8.000 0.9085 0.02284 0.01403 -0.0060 0.0741 1.0000
8.250 0.9244 0.02344 0.01471 -0.0039 0.0704 1.0000
8.500 0.9409 0.02420 0.01541 -0.0021 0.0670 1.0000
8.750 0.9589 0.02526 0.01654 -0.0005 0.0638 1.0000
9.000 0.9759 0.02607 0.01746 0.0013 0.0610 1.0000
9.250 0.9939 0.02697 0.01833 0.0028 0.0586 1.0000
9.500 1.0143 0.02852 0.01994 0.0037 0.0562 1.0000
9.750 1.0292 0.02948 0.02109 0.0058 0.0539 1.0000
10.000 1.0450 0.03044 0.02214 0.0076 0.0519 1.0000
10.250 1.0639 0.03169 0.02338 0.0087 0.0503 1.0000
10.500 1.0822 0.03393 0.02577 0.0097 0.0490 1.0000
10.750 1.0915 0.03550 0.02767 0.0124 0.0480 1.0000
11.000 1.1004 0.03741 0.02988 0.0150 0.0471 1.0000
11.250 1.1076 0.03939 0.03214 0.0177 0.0462 1.0000
11.500 1.1142 0.04126 0.03422 0.0202 0.0453 1.0000
11.750 1.1230 0.04285 0.03593 0.0223 0.0445 1.0000
12.000 1.1355 0.04435 0.03746 0.0237 0.0436 1.0000
12.250 1.1441 0.04698 0.04016 0.0251 0.0429 1.0000
12.500 1.1357 0.04994 0.04337 0.0288 0.0426 1.0000
12.750 1.1208 0.05246 0.04618 0.0331 0.0425 1.0000
13.000 1.1048 0.05541 0.04939 0.0367 0.0424 1.0000
13.250 1.0870 0.05875 0.05298 0.0394 0.0425 1.0000
13.500 1.0679 0.06252 0.05698 0.0413 0.0425 1.0000
13.750 1.0483 0.06675 0.06142 0.0423 0.0426 1.0000
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Polar data table (+)
Polar graphs
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