NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 1,000,000 Max Cl/Cd: 76.77 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08n1-il-1000000-n5.txt Download as CSV file: xf-rc08n1-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(N)1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6224 0.09133 0.08948 0.0209 0.8235 0.0027
-8.000 -0.6251 0.08610 0.08422 0.0161 0.8122 0.0027
-7.750 -0.6262 0.08010 0.07817 0.0096 0.8032 0.0026
-7.500 -0.6229 0.07376 0.07175 0.0043 0.7958 0.0025
-7.250 -0.6169 0.06776 0.06564 0.0005 0.7896 0.0025
-7.000 -0.6079 0.06228 0.06003 -0.0021 0.7836 0.0025
-6.750 -0.5960 0.05730 0.05490 -0.0038 0.7782 0.0026
-6.500 -0.5818 0.05270 0.05015 -0.0049 0.7731 0.0027
-6.250 -0.5659 0.04825 0.04551 -0.0054 0.7680 0.0029
-6.000 -0.5490 0.04357 0.04061 -0.0054 0.7635 0.0032
-5.750 -0.5316 0.03796 0.03470 -0.0046 0.7591 0.0036
-5.500 -0.5145 0.03107 0.02735 -0.0027 0.7548 0.0039
-5.250 -0.4944 0.02659 0.02249 -0.0013 0.7507 0.0043
-5.000 -0.4703 0.02531 0.02107 -0.0010 0.7461 0.0045
-4.750 -0.4461 0.02363 0.01922 -0.0005 0.7414 0.0048
-4.500 -0.4220 0.02120 0.01651 0.0004 0.7372 0.0052
-4.250 -0.3982 0.01618 0.01089 0.0026 0.7333 0.0066
-4.000 -0.3723 0.01459 0.00902 0.0032 0.7288 0.0070
-3.750 -0.3462 0.01386 0.00817 0.0034 0.7244 0.0074
-3.500 -0.3193 0.01360 0.00787 0.0033 0.7199 0.0078
-3.250 -0.2924 0.01304 0.00722 0.0035 0.7150 0.0084
-3.000 -0.2654 0.01232 0.00639 0.0037 0.7105 0.0091
-2.750 -0.2383 0.01170 0.00566 0.0040 0.7057 0.0096
-2.500 -0.2111 0.01117 0.00504 0.0042 0.7004 0.0100
-2.250 -0.1841 0.01072 0.00449 0.0044 0.6954 0.0103
-2.000 -0.1563 0.01056 0.00428 0.0044 0.6901 0.0106
-1.750 -0.1288 0.01028 0.00394 0.0045 0.6845 0.0108
-1.500 -0.1023 0.00973 0.00336 0.0047 0.6788 0.0108
-1.250 -0.0765 0.00905 0.00264 0.0051 0.6717 0.0109
-1.000 -0.0504 0.00849 0.00203 0.0054 0.6649 0.0115
-0.750 -0.0232 0.00821 0.00173 0.0055 0.6573 0.0122
-0.500 0.0043 0.00803 0.00152 0.0056 0.6495 0.0130
-0.250 0.0320 0.00789 0.00134 0.0056 0.6409 0.0137
0.000 0.0598 0.00776 0.00118 0.0055 0.6322 0.0143
0.250 0.0877 0.00767 0.00106 0.0055 0.6238 0.0149
0.500 0.1156 0.00760 0.00098 0.0054 0.6152 0.0156
0.750 0.1436 0.00754 0.00089 0.0054 0.6069 0.0157
1.000 0.1716 0.00750 0.00082 0.0053 0.5978 0.0158
1.250 0.1996 0.00748 0.00076 0.0052 0.5864 0.0161
1.500 0.2276 0.00749 0.00072 0.0051 0.5701 0.0174
1.750 0.2554 0.00753 0.00069 0.0051 0.5480 0.0200
2.000 0.2832 0.00758 0.00071 0.0050 0.5228 0.0269
2.250 0.3107 0.00764 0.00075 0.0049 0.4940 0.0602
2.500 0.3304 0.00626 0.00078 0.0057 0.4670 0.6302
2.750 0.3525 0.00592 0.00083 0.0066 0.4275 0.7801
3.000 0.3835 0.00579 0.00106 0.0060 0.3779 0.9629
3.250 0.4385 0.00621 0.00132 -0.0003 0.3282 0.9940
3.500 0.4741 0.00665 0.00151 -0.0025 0.2661 0.9965
3.750 0.5051 0.00706 0.00170 -0.0036 0.2095 0.9978
4.000 0.5359 0.00741 0.00190 -0.0047 0.1693 0.9989
4.250 0.5670 0.00775 0.00210 -0.0058 0.1351 0.9999
4.500 0.5927 0.00806 0.00230 -0.0056 0.1081 1.0000
4.750 0.6179 0.00836 0.00250 -0.0054 0.0862 1.0000
5.000 0.6432 0.00862 0.00270 -0.0051 0.0715 1.0000
5.250 0.6684 0.00886 0.00294 -0.0048 0.0603 1.0000
5.500 0.6935 0.00914 0.00317 -0.0044 0.0490 1.0000
5.750 0.7184 0.00944 0.00343 -0.0041 0.0385 1.0000
6.000 0.7433 0.00973 0.00369 -0.0038 0.0311 1.0000
6.250 0.7682 0.01001 0.00398 -0.0034 0.0256 1.0000
6.500 0.7928 0.01036 0.00432 -0.0030 0.0188 1.0000
6.750 0.8176 0.01065 0.00463 -0.0027 0.0169 1.0000
7.000 0.8420 0.01100 0.00498 -0.0023 0.0138 1.0000
7.250 0.8663 0.01140 0.00540 -0.0018 0.0108 1.0000
7.500 0.8908 0.01170 0.00573 -0.0015 0.0093 1.0000
7.750 0.9147 0.01210 0.00615 -0.0011 0.0075 1.0000
8.000 0.9383 0.01256 0.00663 -0.0006 0.0060 1.0000
8.250 0.9623 0.01294 0.00706 -0.0002 0.0055 1.0000
8.500 0.9860 0.01337 0.00753 0.0002 0.0047 1.0000
8.750 1.0091 0.01387 0.00806 0.0007 0.0038 1.0000
9.000 1.0320 0.01445 0.00871 0.0012 0.0031 1.0000
9.250 1.0551 0.01496 0.00930 0.0017 0.0027 1.0000
9.500 1.0778 0.01553 0.00993 0.0021 0.0023 1.0000
9.750 1.0999 0.01618 0.01065 0.0026 0.0020 1.0000
10.000 1.1210 0.01700 0.01155 0.0032 0.0016 1.0000
10.250 1.1403 0.01810 0.01280 0.0040 0.0014 1.0000
10.500 1.1608 0.01893 0.01376 0.0047 0.0014 1.0000
10.750 1.1804 0.01986 0.01482 0.0053 0.0013 1.0000
11.000 1.1989 0.02091 0.01601 0.0061 0.0013 1.0000
11.250 1.2160 0.02209 0.01734 0.0069 0.0012 1.0000
11.500 1.2315 0.02340 0.01882 0.0079 0.0012 1.0000
11.750 1.2451 0.02485 0.02044 0.0089 0.0012 1.0000
12.000 1.2562 0.02647 0.02225 0.0101 0.0011 1.0000
12.250 1.2642 0.02826 0.02424 0.0114 0.0011 1.0000
12.500 1.2655 0.03019 0.02636 0.0133 0.0011 1.0000
12.750 1.2641 0.03252 0.02888 0.0145 0.0011 1.0000
13.000 1.2605 0.03534 0.03190 0.0150 0.0011 1.0000
13.250 1.2552 0.03858 0.03534 0.0149 0.0011 1.0000
13.500 1.2468 0.04241 0.03936 0.0143 0.0011 1.0000
13.750 1.2354 0.04690 0.04405 0.0129 0.0011 1.0000
14.000 1.2217 0.05209 0.04942 0.0108 0.0011 1.0000
14.250 1.2053 0.05815 0.05566 0.0077 0.0011 1.0000
14.500 1.1850 0.06549 0.06317 0.0035 0.0011 1.0000
14.750 1.1626 0.07406 0.07190 -0.0018 0.0011 1.0000
15.000 1.1361 0.08436 0.08233 -0.0080 0.0011 1.0000
15.250 1.1044 0.09649 0.09460 -0.0146 0.0011 1.0000
15.500 1.0689 0.10942 0.10763 -0.0207 0.0011 1.0000
15.750 1.0303 0.12283 0.12112 -0.0268 0.0012 1.0000
16.000 0.9925 0.13656 0.13489 -0.0331 0.0012 1.0000
16.250 0.9554 0.15097 0.14935 -0.0398 0.0013 1.0000
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Polar data table (+)
Polar graphs
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