E61 (5.64%) (e61-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: E61 (5.64%) (e61-il) Reynolds number: 1,000,000 Max Cl/Cd: 176.6 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e61-il-1000000-n5.txt Download as CSV file: xf-e61-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: E61 (5.64%)
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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.1650 0.11801 0.11618 -0.0775 0.9815 0.0046
-10.250 -0.1583 0.11527 0.11344 -0.0781 0.9789 0.0046
-10.000 -0.1479 0.11214 0.11031 -0.0801 0.9774 0.0047
-9.750 -0.1367 0.10907 0.10723 -0.0821 0.9763 0.0047
-9.500 -0.1248 0.10590 0.10406 -0.0844 0.9752 0.0047
-8.500 -0.0002 0.07559 0.07387 -0.0974 0.9623 0.0041
-8.250 0.0097 0.07180 0.07008 -0.0995 0.9612 0.0034
-8.000 0.0205 0.06799 0.06628 -0.1018 0.9603 0.0031
-7.500 -0.0364 0.08135 0.07955 -0.1003 0.9615 0.0030
-7.000 0.0352 0.05568 0.05401 -0.1027 0.9411 0.0027
-6.750 0.0419 0.05278 0.05111 -0.1036 0.9367 0.0028
-6.500 0.0452 0.05020 0.04854 -0.1036 0.9299 0.0029
-6.250 0.0585 0.04662 0.04496 -0.1067 0.9250 0.0029
-6.000 0.0799 0.04289 0.04122 -0.1118 0.9225 0.0036
-5.750 0.0973 0.03901 0.03732 -0.1164 0.9158 0.0039
-5.500 0.1257 0.03400 0.03228 -0.1248 0.9116 0.0042
-5.250 0.1648 0.02726 0.02548 -0.1381 0.9075 0.0050
-4.750 0.2017 0.03936 0.03743 -0.1568 0.9113 0.0052
-4.500 0.2529 0.03538 0.03337 -0.1675 0.9093 0.0055
-4.250 0.3006 0.03141 0.02928 -0.1767 0.9043 0.0062
-4.000 0.3717 0.02432 0.02193 -0.1922 0.9019 0.0074
-3.750 0.4520 0.01528 0.01227 -0.2079 0.9017 0.0085
-3.500 0.5027 0.01283 0.00940 -0.2140 0.8991 0.0090
-3.250 0.5535 0.01024 0.00634 -0.2205 0.8964 0.0106
-3.000 0.5883 0.00966 0.00562 -0.2224 0.8893 0.0113
-2.750 0.6253 0.00900 0.00477 -0.2247 0.8824 0.0119
-2.500 0.6579 0.00844 0.00404 -0.2259 0.8733 0.0124
-2.250 0.6896 0.00805 0.00352 -0.2269 0.8639 0.0131
-2.000 0.7212 0.00753 0.00286 -0.2278 0.8537 0.0128
-1.750 0.7513 0.00717 0.00235 -0.2283 0.8419 0.0125
-1.500 0.7802 0.00692 0.00198 -0.2286 0.8290 0.0122
-1.250 0.8083 0.00676 0.00170 -0.2287 0.8156 0.0120
-0.750 0.8629 0.00661 0.00134 -0.2284 0.7896 0.0119
-0.500 0.8895 0.00660 0.00123 -0.2282 0.7766 0.0121
-0.250 0.9157 0.00662 0.00116 -0.2278 0.7633 0.0127
0.000 0.9417 0.00667 0.00113 -0.2274 0.7500 0.0148
0.250 0.9672 0.00675 0.00115 -0.2269 0.7366 0.0162
0.500 0.9926 0.00684 0.00117 -0.2264 0.7225 0.0207
0.750 1.0189 0.00683 0.00128 -0.2261 0.7075 0.0932
1.000 1.0469 0.00670 0.00145 -0.2265 0.6922 0.2753
1.250 1.0743 0.00661 0.00164 -0.2267 0.6759 0.4455
1.500 1.1015 0.00654 0.00186 -0.2269 0.6562 0.6342
2.000 1.1426 0.00647 0.00219 -0.2237 0.6091 1.0000
2.250 1.1659 0.00676 0.00234 -0.2228 0.5802 1.0000
2.500 1.1885 0.00711 0.00253 -0.2218 0.5450 1.0000
2.750 1.2102 0.00756 0.00276 -0.2206 0.5021 1.0000
3.000 1.2316 0.00807 0.00303 -0.2194 0.4544 1.0000
3.250 1.2532 0.00858 0.00335 -0.2182 0.4094 1.0000
3.500 1.2730 0.00926 0.00371 -0.2168 0.3496 1.0000
3.750 1.2886 0.01036 0.00427 -0.2147 0.2515 1.0000
4.000 1.3074 0.01116 0.00474 -0.2131 0.1885 1.0000
4.250 1.3260 0.01199 0.00523 -0.2115 0.1292 1.0000
4.500 1.3443 0.01285 0.00578 -0.2098 0.0743 1.0000
4.750 1.3614 0.01385 0.00646 -0.2079 0.0201 1.0000
5.000 1.3825 0.01442 0.00697 -0.2067 0.0047 1.0000
5.250 1.4054 0.01477 0.00735 -0.2058 0.0035 1.0000
5.500 1.4278 0.01514 0.00776 -0.2048 0.0031 1.0000
5.750 1.4497 0.01557 0.00823 -0.2037 0.0026 1.0000
6.000 1.4707 0.01608 0.00880 -0.2024 0.0021 1.0000
6.250 1.4887 0.01687 0.00973 -0.2004 0.0017 1.0000
6.500 1.5078 0.01744 0.01035 -0.1987 0.0016 1.0000
6.750 1.5257 0.01808 0.01106 -0.1968 0.0016 1.0000
7.000 1.5425 0.01880 0.01190 -0.1947 0.0015 1.0000
7.250 1.5581 0.01964 0.01283 -0.1924 0.0015 1.0000
7.500 1.5726 0.02058 0.01387 -0.1899 0.0015 1.0000
7.750 1.5858 0.02163 0.01503 -0.1872 0.0014 1.0000
8.000 1.5977 0.02279 0.01630 -0.1843 0.0014 1.0000
8.250 1.6087 0.02409 0.01772 -0.1814 0.0014 1.0000
8.500 1.6189 0.02551 0.01927 -0.1783 0.0014 1.0000
8.750 1.6284 0.02707 0.02097 -0.1752 0.0014 1.0000
9.000 1.6393 0.02855 0.02257 -0.1725 0.0014 1.0000
9.250 1.6503 0.02998 0.02413 -0.1699 0.0013 1.0000
9.500 1.6616 0.03136 0.02565 -0.1674 0.0012 1.0000
9.750 1.6717 0.03293 0.02734 -0.1648 0.0012 1.0000
10.000 1.6810 0.03458 0.02913 -0.1622 0.0011 1.0000
10.250 1.6880 0.03669 0.03142 -0.1593 0.0011 1.0000
10.500 1.6947 0.03870 0.03358 -0.1566 0.0010 1.0000
10.750 1.6987 0.04117 0.03625 -0.1536 0.0010 1.0000
11.000 1.7008 0.04388 0.03916 -0.1506 0.0010 1.0000
11.250 1.7017 0.04660 0.04208 -0.1476 0.0010 1.0000
11.500 1.7016 0.04934 0.04500 -0.1449 0.0010 1.0000
11.750 1.6980 0.05262 0.04849 -0.1420 0.0010 1.0000
12.000 1.6923 0.05619 0.05227 -0.1394 0.0009 1.0000
12.250 1.6882 0.05945 0.05569 -0.1372 0.0009 1.0000
12.500 1.6746 0.06441 0.06091 -0.1349 0.0009 1.0000
12.750 1.6653 0.06868 0.06536 -0.1334 0.0009 1.0000
13.000 1.6580 0.07271 0.06957 -0.1325 0.0008 1.0000
13.250 1.6367 0.07956 0.07669 -0.1318 0.0009 1.0000
13.500 1.6202 0.08585 0.08319 -0.1322 0.0009 1.0000
13.750 1.6215 0.08893 0.08631 -0.1328 0.0008 1.0000
14.000 1.5931 0.09832 0.09599 -0.1349 0.0008 1.0000
14.250 1.5709 0.10704 0.10493 -0.1381 0.0009 1.0000
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Polar data table (+)
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