NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 500,000 Max Cl/Cd: 70.38 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc10n1-il-500000-n5.txt Download as CSV file: xf-rc10n1-il-500000-n5.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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5066 0.08600 0.08376 0.0009 0.8452 0.0084
-9.250 -0.5157 0.08029 0.07803 -0.0012 0.8361 0.0084
-8.750 -0.5589 0.06368 0.06137 -0.0127 0.8227 0.0081
-8.250 -0.6766 0.05989 0.05714 -0.0108 0.8262 0.0077
-8.000 -0.6754 0.05509 0.05212 -0.0110 0.8171 0.0078
-7.750 -0.6682 0.05125 0.04808 -0.0109 0.8093 0.0079
-7.500 -0.6553 0.04877 0.04542 -0.0106 0.8022 0.0082
-7.250 -0.6407 0.04628 0.04275 -0.0103 0.7958 0.0085
-7.000 -0.6265 0.04304 0.03929 -0.0097 0.7899 0.0090
-6.750 -0.6134 0.03853 0.03444 -0.0085 0.7847 0.0094
-6.500 -0.5999 0.03336 0.02884 -0.0067 0.7795 0.0097
-6.250 -0.5852 0.02779 0.02271 -0.0045 0.7749 0.0102
-6.000 -0.5672 0.02316 0.01748 -0.0025 0.7706 0.0106
-5.750 -0.5447 0.02081 0.01472 -0.0015 0.7656 0.0111
-5.500 -0.5216 0.01916 0.01285 -0.0009 0.7610 0.0116
-5.250 -0.4966 0.01803 0.01154 -0.0005 0.7567 0.0119
-5.000 -0.4710 0.01703 0.01040 -0.0001 0.7520 0.0123
-4.750 -0.4452 0.01614 0.00934 0.0003 0.7474 0.0127
-4.500 -0.4193 0.01530 0.00837 0.0006 0.7434 0.0131
-4.250 -0.3932 0.01451 0.00748 0.0010 0.7388 0.0135
-4.000 -0.3673 0.01380 0.00667 0.0014 0.7343 0.0139
-3.750 -0.3414 0.01321 0.00599 0.0018 0.7303 0.0143
-3.500 -0.3149 0.01276 0.00547 0.0020 0.7258 0.0149
-3.250 -0.2888 0.01227 0.00491 0.0023 0.7213 0.0153
-3.000 -0.2629 0.01178 0.00435 0.0027 0.7171 0.0155
-2.750 -0.2371 0.01125 0.00377 0.0031 0.7128 0.0159
-2.500 -0.2110 0.01077 0.00323 0.0034 0.7082 0.0169
-2.250 -0.1842 0.01049 0.00291 0.0036 0.7038 0.0182
-2.000 -0.1569 0.01026 0.00265 0.0037 0.6995 0.0198
-1.750 -0.1294 0.01006 0.00242 0.0037 0.6948 0.0219
-1.500 -0.1021 0.00983 0.00217 0.0039 0.6901 0.0269
-1.250 -0.0745 0.00968 0.00201 0.0039 0.6856 0.0345
-1.000 -0.0470 0.00948 0.00187 0.0039 0.6802 0.0543
-0.750 -0.0309 0.00755 0.00153 0.0052 0.6752 0.5159
-0.500 -0.0074 0.00710 0.00149 0.0061 0.6703 0.6416
-0.250 0.0178 0.00684 0.00143 0.0066 0.6644 0.7022
0.250 0.0667 0.00620 0.00150 0.0089 0.6521 0.9014
0.500 0.1017 0.00626 0.00156 0.0076 0.6455 0.9366
0.750 0.1336 0.00630 0.00156 0.0068 0.6385 0.9463
1.000 0.1672 0.00635 0.00156 0.0055 0.6312 0.9532
1.250 0.1990 0.00640 0.00158 0.0046 0.6235 0.9601
1.750 0.2645 0.00652 0.00163 0.0025 0.6085 0.9704
2.000 0.2971 0.00658 0.00166 0.0014 0.6004 0.9744
2.250 0.3308 0.00664 0.00170 0.0000 0.5908 0.9778
2.500 0.3625 0.00671 0.00174 -0.0009 0.5794 0.9818
2.750 0.3941 0.00682 0.00177 -0.0018 0.5565 0.9854
3.000 0.4278 0.00696 0.00181 -0.0033 0.5304 0.9882
3.250 0.4608 0.00709 0.00188 -0.0046 0.5056 0.9913
3.500 0.4926 0.00730 0.00199 -0.0056 0.4727 0.9943
3.750 0.5251 0.00758 0.00210 -0.0070 0.4270 0.9960
4.000 0.5574 0.00792 0.00227 -0.0084 0.3829 0.9979
4.250 0.5893 0.00839 0.00252 -0.0098 0.3262 0.9998
4.500 0.6147 0.00894 0.00280 -0.0099 0.2641 1.0000
4.750 0.6388 0.00942 0.00307 -0.0096 0.2144 1.0000
5.000 0.6626 0.00990 0.00336 -0.0093 0.1707 1.0000
5.250 0.6859 0.01039 0.00369 -0.0088 0.1334 1.0000
5.500 0.7094 0.01082 0.00401 -0.0084 0.1084 1.0000
5.750 0.7330 0.01117 0.00431 -0.0078 0.0917 1.0000
6.000 0.7564 0.01155 0.00463 -0.0073 0.0764 1.0000
6.250 0.7798 0.01191 0.00496 -0.0068 0.0645 1.0000
6.500 0.8029 0.01229 0.00532 -0.0062 0.0547 1.0000
6.750 0.8258 0.01270 0.00569 -0.0055 0.0458 1.0000
7.000 0.8490 0.01304 0.00606 -0.0049 0.0407 1.0000
7.250 0.8715 0.01349 0.00649 -0.0043 0.0349 1.0000
7.500 0.8944 0.01385 0.00690 -0.0036 0.0316 1.0000
7.750 0.9163 0.01433 0.00739 -0.0029 0.0274 1.0000
8.000 0.9383 0.01480 0.00790 -0.0021 0.0248 1.0000
8.250 0.9603 0.01523 0.00839 -0.0014 0.0226 1.0000
8.500 0.9816 0.01574 0.00892 -0.0007 0.0201 1.0000
8.750 1.0017 0.01639 0.00960 0.0003 0.0178 1.0000
9.000 1.0230 0.01690 0.01020 0.0010 0.0168 1.0000
9.250 1.0437 0.01746 0.01084 0.0018 0.0157 1.0000
9.500 1.0640 0.01806 0.01150 0.0027 0.0145 1.0000
9.750 1.0833 0.01876 0.01223 0.0035 0.0132 1.0000
10.000 1.1004 0.01968 0.01322 0.0046 0.0121 1.0000
10.250 1.1198 0.02032 0.01396 0.0055 0.0115 1.0000
10.500 1.1378 0.02107 0.01483 0.0064 0.0108 1.0000
10.750 1.1550 0.02188 0.01572 0.0074 0.0102 1.0000
11.000 1.1715 0.02270 0.01663 0.0084 0.0096 1.0000
11.250 1.1863 0.02363 0.01762 0.0095 0.0090 1.0000
11.500 1.1944 0.02494 0.01902 0.0114 0.0083 1.0000
11.750 1.2040 0.02606 0.02026 0.0129 0.0080 1.0000
12.000 1.2150 0.02718 0.02149 0.0140 0.0077 1.0000
12.250 1.2246 0.02850 0.02294 0.0149 0.0073 1.0000
12.500 1.2330 0.03001 0.02457 0.0157 0.0070 1.0000
12.750 1.2408 0.03162 0.02630 0.0164 0.0067 1.0000
13.000 1.2480 0.03335 0.02817 0.0169 0.0064 1.0000
13.250 1.2540 0.03522 0.03016 0.0172 0.0062 1.0000
13.500 1.2587 0.03729 0.03234 0.0175 0.0060 1.0000
13.750 1.2608 0.03968 0.03485 0.0176 0.0058 1.0000
14.000 1.2591 0.04258 0.03787 0.0175 0.0056 1.0000
14.250 1.2522 0.04619 0.04163 0.0171 0.0054 1.0000
14.500 1.2486 0.04953 0.04513 0.0166 0.0053 1.0000
14.750 1.2447 0.05305 0.04881 0.0157 0.0052 1.0000
15.000 1.2389 0.05696 0.05287 0.0146 0.0051 1.0000
15.250 1.2307 0.06137 0.05745 0.0130 0.0050 1.0000
15.500 1.2207 0.06626 0.06250 0.0111 0.0049 1.0000
15.750 1.2082 0.07191 0.06831 0.0085 0.0049 1.0000
16.000 1.1938 0.07825 0.07481 0.0054 0.0048 1.0000
16.250 1.1760 0.08552 0.08224 0.0017 0.0048 1.0000
16.500 1.1555 0.09365 0.09053 -0.0026 0.0048 1.0000
16.750 1.1315 0.10278 0.09982 -0.0074 0.0049 1.0000
17.000 1.1045 0.11284 0.11003 -0.0126 0.0050 1.0000
17.250 1.0755 0.12372 0.12104 -0.0181 0.0051 1.0000
17.500 1.0451 0.13522 0.13266 -0.0239 0.0052 1.0000
17.750 1.0148 0.14732 0.14486 -0.0301 0.0053 1.0000
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Polar data table (+)
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