NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 100,000 Max Cl/Cd: 41.9 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc1064c-il-100000-n5.txt Download as CSV file: xf-rc1064c-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC10-64C 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5948 0.09214 0.08684 -0.0261 1.0000 0.0355
-10.000 -0.6057 0.08515 0.07990 -0.0308 1.0000 0.0350
-9.750 -0.6259 0.07658 0.07133 -0.0376 1.0000 0.0343
-9.500 -0.6514 0.07012 0.06481 -0.0396 1.0000 0.0337
-9.250 -0.6748 0.06480 0.05937 -0.0385 1.0000 0.0333
-9.000 -0.6937 0.05929 0.05364 -0.0368 1.0000 0.0330
-8.750 -0.7078 0.05434 0.04842 -0.0343 1.0000 0.0328
-8.500 -0.7163 0.04999 0.04374 -0.0313 1.0000 0.0328
-8.250 -0.7194 0.04620 0.03961 -0.0283 1.0000 0.0330
-8.000 -0.7183 0.04285 0.03590 -0.0253 1.0000 0.0332
-7.750 -0.7137 0.03985 0.03253 -0.0224 1.0000 0.0336
-7.500 -0.7057 0.03731 0.02964 -0.0198 1.0000 0.0344
-7.250 -0.6949 0.03519 0.02717 -0.0173 1.0000 0.0358
-7.000 -0.6827 0.03302 0.02458 -0.0149 1.0000 0.0373
-6.750 -0.6681 0.03092 0.02203 -0.0127 1.0000 0.0382
-6.500 -0.6512 0.02904 0.01977 -0.0108 1.0000 0.0389
-6.250 -0.6329 0.02725 0.01774 -0.0092 1.0000 0.0397
-6.000 -0.6141 0.02581 0.01622 -0.0078 1.0000 0.0409
-5.750 -0.5947 0.02470 0.01501 -0.0064 1.0000 0.0424
-5.500 -0.5749 0.02376 0.01398 -0.0050 1.0000 0.0445
-5.250 -0.5547 0.02293 0.01298 -0.0037 1.0000 0.0478
-5.000 -0.5344 0.02198 0.01193 -0.0024 1.0000 0.0505
-4.750 -0.5150 0.02112 0.01108 -0.0010 1.0000 0.0534
-4.500 -0.4951 0.02040 0.01033 0.0003 1.0000 0.0574
-4.250 -0.4752 0.01974 0.00956 0.0017 1.0000 0.0626
-4.000 -0.4555 0.01915 0.00902 0.0030 1.0000 0.0713
-3.750 -0.4304 0.01854 0.00846 0.0031 0.9984 0.0835
-3.500 -0.3969 0.01798 0.00793 0.0015 0.9942 0.1054
-3.250 -0.3647 0.01739 0.00754 0.0001 0.9890 0.1372
-3.000 -0.3293 0.01718 0.00736 -0.0017 0.9839 0.1841
-2.750 -0.2995 0.01641 0.00677 -0.0027 0.9780 0.2182
-2.500 -0.2698 0.01532 0.00647 -0.0039 0.9733 0.3624
-2.250 -0.2454 0.01406 0.00653 -0.0030 0.9692 0.6610
-2.000 -0.2116 0.01403 0.00686 -0.0030 0.9653 0.7942
-1.750 -0.1702 0.01433 0.00721 -0.0045 0.9628 0.8727
-1.500 -0.1231 0.01472 0.00752 -0.0075 0.9614 0.9216
-1.250 -0.0696 0.01505 0.00774 -0.0122 0.9607 0.9496
-1.000 -0.0151 0.01518 0.00775 -0.0177 0.9596 0.9634
-0.750 0.0305 0.01518 0.00766 -0.0216 0.9554 0.9712
-0.500 0.0804 0.01509 0.00750 -0.0264 0.9517 0.9747
-0.250 0.1298 0.01499 0.00734 -0.0311 0.9479 0.9783
0.000 0.1752 0.01493 0.00726 -0.0350 0.9429 0.9832
0.250 0.2244 0.01483 0.00714 -0.0397 0.9393 0.9862
0.500 0.2688 0.01479 0.00711 -0.0434 0.9331 0.9900
0.750 0.3101 0.01480 0.00716 -0.0464 0.9254 0.9937
1.000 0.3439 0.01486 0.00727 -0.0479 0.9116 0.9983
1.250 0.3492 0.01494 0.00741 -0.0438 0.8840 1.0000
1.500 0.3782 0.01470 0.00721 -0.0438 0.8699 1.0000
1.750 0.4100 0.01437 0.00692 -0.0440 0.8568 1.0000
2.000 0.4395 0.01410 0.00665 -0.0437 0.8386 1.0000
2.250 0.4651 0.01393 0.00648 -0.0427 0.8148 1.0000
2.500 0.4875 0.01383 0.00642 -0.0412 0.7870 1.0000
2.750 0.5106 0.01374 0.00631 -0.0397 0.7568 1.0000
3.000 0.5332 0.01371 0.00623 -0.0381 0.7228 1.0000
3.250 0.5548 0.01375 0.00621 -0.0364 0.6827 1.0000
3.500 0.5758 0.01386 0.00616 -0.0345 0.6263 1.0000
3.750 0.5942 0.01418 0.00613 -0.0321 0.5442 1.0000
4.000 0.6095 0.01478 0.00623 -0.0295 0.4426 1.0000
4.250 0.6226 0.01564 0.00652 -0.0268 0.3340 1.0000
4.500 0.6363 0.01657 0.00696 -0.0245 0.2434 1.0000
4.750 0.6511 0.01745 0.00750 -0.0224 0.1779 1.0000
5.000 0.6668 0.01827 0.00806 -0.0204 0.1358 1.0000
5.250 0.6833 0.01900 0.00867 -0.0185 0.1113 1.0000
5.500 0.6999 0.01972 0.00932 -0.0166 0.0968 1.0000
5.750 0.7166 0.02043 0.01002 -0.0146 0.0866 1.0000
6.000 0.7322 0.02124 0.01078 -0.0126 0.0789 1.0000
6.250 0.7495 0.02193 0.01156 -0.0107 0.0732 1.0000
6.500 0.7661 0.02273 0.01239 -0.0088 0.0691 1.0000
6.750 0.7820 0.02374 0.01336 -0.0069 0.0658 1.0000
7.000 0.8009 0.02455 0.01430 -0.0053 0.0622 1.0000
7.250 0.8195 0.02541 0.01525 -0.0038 0.0586 1.0000
7.500 0.8382 0.02642 0.01628 -0.0024 0.0561 1.0000
7.750 0.8575 0.02780 0.01765 -0.0013 0.0540 1.0000
8.000 0.8784 0.02913 0.01914 -0.0002 0.0523 1.0000
8.250 0.8985 0.03038 0.02061 0.0010 0.0500 1.0000
8.500 0.9174 0.03163 0.02207 0.0022 0.0475 1.0000
8.750 0.9360 0.03306 0.02366 0.0035 0.0458 1.0000
9.000 0.9538 0.03466 0.02540 0.0047 0.0445 1.0000
9.250 0.9708 0.03656 0.02743 0.0059 0.0434 1.0000
9.500 0.9850 0.03903 0.03013 0.0073 0.0422 1.0000
9.750 0.9945 0.04108 0.03263 0.0097 0.0410 1.0000
10.000 1.0012 0.04337 0.03537 0.0121 0.0395 1.0000
10.250 1.0050 0.04594 0.03832 0.0146 0.0384 1.0000
10.500 1.0049 0.04878 0.04151 0.0173 0.0377 1.0000
10.750 1.0006 0.05174 0.04479 0.0202 0.0371 1.0000
11.000 0.9908 0.05470 0.04804 0.0234 0.0367 1.0000
11.250 0.9767 0.05776 0.05134 0.0265 0.0364 1.0000
11.500 0.9602 0.06116 0.05498 0.0287 0.0361 1.0000
11.750 0.9394 0.06529 0.05933 0.0297 0.0359 1.0000
12.000 0.9112 0.07087 0.06515 0.0289 0.0361 1.0000
12.250 0.8695 0.07974 0.07428 0.0244 0.0367 1.0000
12.500 0.8141 0.09524 0.08993 0.0131 0.0379 1.0000
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Polar data table (+)
Polar graphs
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