NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 100,000 Max Cl/Cd: 49 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10n1-il-100000.txt Download as CSV file: xf-rc10n1-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(N)1 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5843 0.11013 0.10562 0.0045 1.0000 0.0762
-9.250 -0.5955 0.10537 0.10093 -0.0037 1.0000 0.0768
-9.000 -0.6098 0.10070 0.09626 -0.0105 1.0000 0.0770
-8.750 -0.5766 0.09641 0.09203 0.0001 1.0000 0.0811
-8.500 -0.5734 0.09243 0.08808 -0.0016 1.0000 0.0841
-8.250 -0.5774 0.08746 0.08317 -0.0068 1.0000 0.0866
-8.000 -0.5871 0.08262 0.07829 -0.0123 1.0000 0.0891
-7.750 -0.6074 0.07977 0.07505 -0.0183 1.0000 0.0913
-7.500 -0.5871 0.07332 0.06892 -0.0166 1.0000 0.0950
-7.250 -0.5798 0.06992 0.06545 -0.0175 1.0000 0.1023
-7.000 -0.5790 0.06544 0.06080 -0.0195 1.0000 0.1084
-6.500 -0.5565 0.05836 0.05363 -0.0200 1.0000 0.1265
-6.250 -0.5471 0.05488 0.05001 -0.0206 1.0000 0.1386
-6.000 -0.5360 0.05179 0.04684 -0.0205 1.0000 0.1530
-5.750 -0.5268 0.04912 0.04415 -0.0195 1.0000 0.1689
-5.500 -0.5290 0.04729 0.04230 -0.0165 1.0000 0.1836
-5.250 -0.5328 0.04568 0.04069 -0.0129 1.0000 0.1995
-4.500 -0.5071 0.03930 0.03446 -0.0072 0.9934 0.3005
-4.250 -0.4832 0.03706 0.03242 -0.0061 0.9866 0.3602
-4.000 -0.4187 0.03153 0.02511 -0.0157 0.9789 0.2387
-3.750 -0.3463 0.02711 0.01845 -0.0155 0.9744 0.0804
-3.500 -0.3015 0.02519 0.01612 -0.0179 0.9700 0.0746
-3.250 -0.2639 0.02333 0.01405 -0.0194 0.9635 0.0730
-3.000 -0.2206 0.02212 0.01256 -0.0220 0.9584 0.0766
-2.750 -0.1801 0.02036 0.01085 -0.0243 0.9537 0.0796
-2.500 -0.1474 0.01920 0.00981 -0.0253 0.9464 0.0845
-2.250 -0.1116 0.01822 0.00884 -0.0269 0.9408 0.0966
-2.000 -0.0855 0.01754 0.00813 -0.0266 0.9320 0.1166
-1.750 0.0844 0.01490 0.00832 -0.0460 0.9566 1.0000
-1.500 0.1145 0.01476 0.00800 -0.0479 0.9444 1.0000
-1.250 0.1408 0.01471 0.00780 -0.0489 0.9326 1.0000
-1.000 0.1650 0.01473 0.00768 -0.0493 0.9214 1.0000
-0.750 0.1888 0.01481 0.00763 -0.0492 0.9113 1.0000
-0.500 0.2108 0.01495 0.00767 -0.0488 0.9005 1.0000
-0.250 0.2327 0.01514 0.00777 -0.0483 0.8902 1.0000
0.000 0.2548 0.01535 0.00790 -0.0476 0.8808 1.0000
0.250 0.2763 0.01557 0.00805 -0.0467 0.8712 1.0000
0.500 0.2979 0.01585 0.00828 -0.0459 0.8612 1.0000
0.750 0.3192 0.01612 0.00850 -0.0448 0.8519 1.0000
1.000 0.3404 0.01637 0.00873 -0.0436 0.8424 1.0000
1.250 0.3619 0.01670 0.00903 -0.0427 0.8320 1.0000
1.500 0.3831 0.01699 0.00931 -0.0414 0.8226 1.0000
1.750 0.4042 0.01723 0.00953 -0.0399 0.8133 1.0000
2.000 0.4256 0.01756 0.00990 -0.0389 0.8023 1.0000
2.250 0.4468 0.01787 0.01022 -0.0376 0.7918 1.0000
2.500 0.4680 0.01805 0.01040 -0.0359 0.7829 1.0000
2.750 0.4895 0.01832 0.01071 -0.0346 0.7716 1.0000
3.000 0.5109 0.01861 0.01105 -0.0334 0.7598 1.0000
3.250 0.5323 0.01882 0.01134 -0.0319 0.7482 1.0000
3.500 0.5537 0.01892 0.01149 -0.0301 0.7366 1.0000
3.750 0.5746 0.01883 0.01144 -0.0278 0.7223 1.0000
4.000 0.5952 0.01855 0.01120 -0.0253 0.7056 1.0000
4.250 0.6164 0.01817 0.01089 -0.0227 0.6888 1.0000
4.500 0.6384 0.01781 0.01057 -0.0204 0.6725 1.0000
4.750 0.6610 0.01739 0.01019 -0.0182 0.6557 1.0000
5.000 0.6834 0.01704 0.00994 -0.0162 0.6349 1.0000
5.250 0.7063 0.01656 0.00953 -0.0141 0.6131 1.0000
5.500 0.7288 0.01618 0.00923 -0.0121 0.5852 1.0000
5.750 0.7509 0.01589 0.00900 -0.0102 0.5475 1.0000
6.000 0.7718 0.01575 0.00882 -0.0082 0.4919 1.0000
6.250 0.7879 0.01624 0.00883 -0.0056 0.3845 1.0000
6.500 0.7959 0.01822 0.00980 -0.0030 0.2440 1.0000
6.750 0.8074 0.02010 0.01110 -0.0010 0.1796 1.0000
7.000 0.8224 0.02169 0.01240 0.0007 0.1483 1.0000
7.250 0.8393 0.02322 0.01376 0.0022 0.1270 1.0000
7.500 0.8588 0.02479 0.01526 0.0035 0.1120 1.0000
7.750 0.8790 0.02637 0.01682 0.0045 0.0994 1.0000
8.000 0.9020 0.02850 0.01878 0.0052 0.0901 1.0000
8.250 0.9228 0.02988 0.02049 0.0064 0.0821 1.0000
8.500 0.9453 0.03237 0.02297 0.0070 0.0760 1.0000
8.750 0.9656 0.03459 0.02562 0.0083 0.0717 1.0000
9.000 0.9847 0.03655 0.02767 0.0091 0.0665 1.0000
9.250 1.0004 0.04014 0.03159 0.0102 0.0635 1.0000
9.500 1.0117 0.04336 0.03538 0.0121 0.0623 1.0000
9.750 1.0188 0.04697 0.03953 0.0140 0.0612 1.0000
10.000 1.0210 0.05055 0.04359 0.0159 0.0597 1.0000
10.250 1.0197 0.05415 0.04758 0.0177 0.0583 1.0000
10.500 1.0124 0.05825 0.05206 0.0195 0.0581 1.0000
10.750 0.9976 0.06278 0.05693 0.0212 0.0589 1.0000
11.000 0.9775 0.06697 0.06135 0.0228 0.0597 1.0000
11.250 0.9559 0.07155 0.06610 0.0229 0.0606 1.0000
11.500 0.9335 0.07673 0.07144 0.0216 0.0614 1.0000
11.750 0.9126 0.08252 0.07733 0.0195 0.0623 1.0000
12.000 0.8961 0.08873 0.08360 0.0170 0.0632 1.0000
12.250 0.8290 0.10700 0.10199 0.0037 0.0758 1.0000
12.500 0.8415 0.11061 0.10566 0.0054 0.0775 1.0000
12.750 0.7547 0.14308 0.13786 -0.0212 0.1339 1.0000
13.000 0.6171 0.13728 0.13233 -0.0030 0.1440 1.0000
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Polar data table (+)
Polar graphs
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