NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 50,000 Max Cl/Cd: 32.18 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc510-il-50000-n5.txt Download as CSV file: xf-rc510-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(5)-10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5178 0.11632 0.10969 0.0039 1.0000 0.1115
-9.250 -0.5277 0.11311 0.10657 -0.0016 1.0000 0.1127
-9.000 -0.5351 0.10924 0.10278 -0.0070 1.0000 0.1131
-8.750 -0.5072 0.10410 0.09765 -0.0016 1.0000 0.1164
-8.000 -0.5079 0.08599 0.07952 -0.0170 1.0000 0.0680
-7.750 -0.5029 0.08164 0.07518 -0.0187 1.0000 0.0654
-7.250 -0.5061 0.07009 0.06322 -0.0265 1.0000 0.0576
-7.000 -0.4964 0.06626 0.05934 -0.0271 1.0000 0.0570
-6.750 -0.4869 0.06244 0.05541 -0.0277 1.0000 0.0563
-6.500 -0.4771 0.05872 0.05155 -0.0280 1.0000 0.0558
-6.250 -0.4675 0.05521 0.04786 -0.0279 1.0000 0.0554
-6.000 -0.4600 0.05209 0.04451 -0.0269 1.0000 0.0552
-5.750 -0.4570 0.04953 0.04172 -0.0248 1.0000 0.0555
-5.500 -0.4554 0.04724 0.03914 -0.0220 1.0000 0.0562
-5.250 -0.4294 0.04353 0.03478 -0.0235 0.9878 0.0576
-5.000 -0.4001 0.04030 0.03114 -0.0253 0.9772 0.0588
-4.750 -0.3683 0.03755 0.02805 -0.0271 0.9683 0.0598
-4.250 -0.3036 0.03303 0.02278 -0.0299 0.9514 0.0640
-4.000 -0.2686 0.03106 0.02027 -0.0311 0.9441 0.0692
-3.750 -0.2366 0.02940 0.01842 -0.0321 0.9361 0.0733
-3.500 -0.2017 0.02806 0.01684 -0.0334 0.9294 0.0803
-3.250 -0.1690 0.02685 0.01551 -0.0342 0.9221 0.0886
-3.000 -0.1342 0.02581 0.01430 -0.0352 0.9155 0.1007
-2.750 -0.1004 0.02491 0.01330 -0.0361 0.9091 0.1156
-2.500 -0.0698 0.02411 0.01247 -0.0367 0.9020 0.1343
-2.250 -0.0406 0.02343 0.01170 -0.0370 0.8955 0.1587
-2.000 -0.0165 0.02274 0.01111 -0.0366 0.8877 0.1878
-1.750 0.0093 0.02181 0.01048 -0.0366 0.8817 0.2399
-1.500 0.1242 0.01932 0.01032 -0.0495 0.8868 0.9909
-1.250 0.1607 0.01938 0.01005 -0.0516 0.8804 1.0000
-1.000 0.1834 0.01958 0.01003 -0.0511 0.8718 1.0000
-0.750 0.2071 0.01979 0.01004 -0.0507 0.8644 1.0000
-0.500 0.2294 0.02004 0.01012 -0.0501 0.8565 1.0000
-0.250 0.2528 0.02030 0.01023 -0.0496 0.8495 1.0000
0.000 0.2746 0.02061 0.01042 -0.0489 0.8417 1.0000
0.250 0.2982 0.02091 0.01059 -0.0483 0.8353 1.0000
0.500 0.3193 0.02128 0.01088 -0.0475 0.8272 1.0000
0.750 0.3429 0.02158 0.01110 -0.0468 0.8202 1.0000
1.000 0.3645 0.02186 0.01131 -0.0456 0.8091 1.0000
1.250 0.3856 0.02201 0.01139 -0.0437 0.7948 1.0000
1.500 0.4066 0.02201 0.01129 -0.0413 0.7780 1.0000
1.750 0.4270 0.02195 0.01114 -0.0388 0.7608 1.0000
2.000 0.4472 0.02198 0.01113 -0.0366 0.7440 1.0000
2.250 0.4690 0.02201 0.01112 -0.0347 0.7307 1.0000
2.500 0.4903 0.02210 0.01121 -0.0331 0.7165 1.0000
2.750 0.5117 0.02215 0.01128 -0.0314 0.7018 1.0000
3.000 0.5333 0.02216 0.01130 -0.0296 0.6864 1.0000
3.250 0.5552 0.02210 0.01125 -0.0278 0.6701 1.0000
3.500 0.5761 0.02209 0.01129 -0.0260 0.6506 1.0000
3.750 0.5975 0.02201 0.01127 -0.0242 0.6294 1.0000
4.000 0.6190 0.02188 0.01117 -0.0223 0.6053 1.0000
4.250 0.6402 0.02178 0.01109 -0.0203 0.5754 1.0000
4.500 0.6612 0.02171 0.01105 -0.0184 0.5376 1.0000
4.750 0.6818 0.02169 0.01094 -0.0164 0.4876 1.0000
5.000 0.7018 0.02183 0.01072 -0.0141 0.4299 1.0000
5.250 0.7204 0.02239 0.01084 -0.0121 0.3838 1.0000
5.500 0.7388 0.02322 0.01138 -0.0106 0.3513 1.0000
5.750 0.7577 0.02412 0.01206 -0.0093 0.3277 1.0000
6.000 0.7775 0.02501 0.01279 -0.0082 0.3094 1.0000
6.250 0.7980 0.02589 0.01359 -0.0072 0.2935 1.0000
6.500 0.8193 0.02676 0.01443 -0.0063 0.2800 1.0000
6.750 0.8413 0.02761 0.01532 -0.0055 0.2675 1.0000
7.000 0.8636 0.02849 0.01625 -0.0048 0.2559 1.0000
7.250 0.8864 0.02939 0.01716 -0.0041 0.2458 1.0000
7.500 0.9094 0.03032 0.01816 -0.0034 0.2363 1.0000
7.750 0.9317 0.03131 0.01929 -0.0027 0.2267 1.0000
8.000 0.9548 0.03230 0.02025 -0.0021 0.2185 1.0000
8.250 0.9762 0.03342 0.02161 -0.0015 0.2098 1.0000
8.500 0.9983 0.03452 0.02280 -0.0008 0.2022 1.0000
8.750 1.0183 0.03572 0.02422 -0.0001 0.1939 1.0000
9.000 1.0393 0.03698 0.02559 0.0005 0.1870 1.0000
9.250 1.0575 0.03833 0.02718 0.0014 0.1792 1.0000
9.500 1.0766 0.03966 0.02863 0.0021 0.1725 1.0000
9.750 1.0923 0.04128 0.03060 0.0031 0.1654 1.0000
10.000 1.1110 0.04257 0.03193 0.0039 0.1592 1.0000
10.250 1.1209 0.04452 0.03430 0.0051 0.1520 1.0000
10.500 1.1403 0.04565 0.03541 0.0058 0.1463 1.0000
10.750 1.1427 0.04823 0.03850 0.0074 0.1401 1.0000
11.000 1.1553 0.04957 0.03997 0.0085 0.1341 1.0000
11.250 1.1581 0.05200 0.04269 0.0099 0.1290 1.0000
11.500 1.1546 0.05472 0.04574 0.0115 0.1241 1.0000
11.750 1.1725 0.05530 0.04623 0.0124 0.1187 1.0000
12.000 1.1506 0.05929 0.05063 0.0144 0.1160 1.0000
12.250 1.1274 0.06380 0.05545 0.0150 0.1137 1.0000
12.500 1.1018 0.06911 0.06101 0.0144 0.1120 1.0000
12.750 1.0651 0.07655 0.06865 0.0119 0.1115 1.0000
13.000 0.9881 0.09233 0.08451 0.0031 0.1140 1.0000
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Polar data table (+)
Polar graphs
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