NREL's S827 Airfoil (s827-nr) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S827 Airfoil (s827-nr) Reynolds number: 500,000 Max Cl/Cd: 84.44 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s827-nr-500000.txt Download as CSV file: xf-s827-nr-500000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S827 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.1586 0.08967 0.08668 -0.1256 0.8297 0.0149
-13.250 -0.2744 0.07876 0.07539 -0.1424 0.8450 0.0136
-13.000 -0.2890 0.07488 0.07141 -0.1427 0.8414 0.0136
-12.750 -0.3101 0.07056 0.06699 -0.1425 0.8381 0.0137
-12.500 -0.3272 0.06698 0.06329 -0.1419 0.8349 0.0138
-12.250 -0.3441 0.06364 0.05981 -0.1409 0.8319 0.0138
-12.000 -0.3570 0.06076 0.05687 -0.1398 0.8290 0.0140
-11.750 -0.3686 0.05816 0.05418 -0.1383 0.8259 0.0141
-11.500 -0.3796 0.05568 0.05159 -0.1366 0.8230 0.0142
-11.250 -0.3900 0.05353 0.04933 -0.1347 0.8203 0.0144
-11.000 -0.3976 0.05149 0.04714 -0.1327 0.8178 0.0145
-10.750 -0.4060 0.04951 0.04509 -0.1304 0.8153 0.0147
-10.500 -0.4121 0.04757 0.04305 -0.1279 0.8127 0.0149
-10.250 -0.4162 0.04571 0.04106 -0.1253 0.8103 0.0151
-10.000 -0.4192 0.04388 0.03910 -0.1225 0.8081 0.0155
-9.750 -0.4203 0.04202 0.03709 -0.1196 0.8061 0.0160
-9.500 -0.4178 0.04025 0.03513 -0.1167 0.8042 0.0167
-9.250 -0.3203 0.02911 0.02277 -0.1219 0.8054 0.0095
-9.000 -0.2345 0.02771 0.02114 -0.1258 0.8056 0.0080
-8.750 -0.1898 0.02795 0.02119 -0.1256 0.8046 0.0079
-8.500 -0.1662 0.02792 0.02109 -0.1245 0.8032 0.0080
-8.250 -0.1486 0.02763 0.02078 -0.1232 0.8016 0.0079
-8.000 -0.1343 0.02717 0.02030 -0.1219 0.7999 0.0081
-7.750 -0.1224 0.02656 0.01966 -0.1204 0.7983 0.0088
-7.500 -0.1115 0.02588 0.01895 -0.1190 0.7968 0.0086
-7.250 -0.1086 0.02487 0.01792 -0.1166 0.7951 0.0101
-7.000 -0.1001 0.02405 0.01709 -0.1150 0.7937 0.0124
-6.750 -0.0972 0.02299 0.01598 -0.1128 0.7923 0.0144
-6.500 -0.0975 0.02179 0.01479 -0.1102 0.7908 0.0225
-6.250 -0.0988 0.02055 0.01359 -0.1076 0.7894 0.0336
-6.000 -0.1016 0.01925 0.01238 -0.1047 0.7880 0.0519
-5.750 -0.1054 0.01793 0.01115 -0.1017 0.7864 0.0748
-5.500 -0.1140 0.01645 0.00979 -0.0980 0.7846 0.1065
-5.250 -0.1270 0.01499 0.00845 -0.0934 0.7826 0.1455
-5.000 -0.1519 0.01396 0.00753 -0.0860 0.7805 0.1924
-4.750 -0.1794 0.01219 0.00617 -0.0782 0.7788 0.3394
-4.500 -0.1785 0.01125 0.00580 -0.0746 0.7777 0.5195
-4.250 -0.1521 0.01158 0.00602 -0.0748 0.7769 0.5421
-4.000 -0.1243 0.01190 0.00632 -0.0751 0.7762 0.5518
-3.750 -0.0974 0.01221 0.00660 -0.0754 0.7754 0.5597
-3.500 -0.0710 0.01265 0.00702 -0.0754 0.7747 0.5680
-3.250 -0.0450 0.01282 0.00717 -0.0755 0.7739 0.5723
-3.000 -0.0175 0.01315 0.00754 -0.0758 0.7730 0.5766
-2.750 0.0087 0.01361 0.00797 -0.0758 0.7721 0.5830
-2.500 0.0344 0.01366 0.00796 -0.0760 0.7711 0.5847
-2.250 0.0604 0.01359 0.00786 -0.0763 0.7700 0.5853
-2.000 0.0866 0.01358 0.00783 -0.0767 0.7691 0.5859
-1.750 0.1129 0.01359 0.00782 -0.0770 0.7682 0.5866
-1.500 0.1394 0.01362 0.00783 -0.0774 0.7674 0.5872
-1.250 0.1660 0.01365 0.00785 -0.0778 0.7667 0.5878
-1.000 0.1927 0.01369 0.00788 -0.0782 0.7659 0.5885
-0.750 0.2196 0.01375 0.00792 -0.0787 0.7653 0.5891
-0.500 0.2466 0.01381 0.00797 -0.0792 0.7646 0.5899
-0.250 0.2744 0.01385 0.00800 -0.0798 0.7639 0.5907
0.000 0.3025 0.01390 0.00805 -0.0805 0.7632 0.5915
0.250 0.3301 0.01400 0.00814 -0.0810 0.7625 0.5924
0.500 0.3572 0.01416 0.00829 -0.0816 0.7619 0.5933
0.750 0.3759 0.01444 0.00861 -0.0805 0.7607 0.5942
1.000 0.3868 0.01485 0.00908 -0.0780 0.7582 0.5952
1.250 0.4012 0.01523 0.00949 -0.0761 0.7560 0.5961
1.500 0.4222 0.01545 0.00971 -0.0755 0.7543 0.5970
1.750 0.4469 0.01559 0.00986 -0.0755 0.7530 0.5979
2.000 0.4738 0.01567 0.00995 -0.0760 0.7520 0.5988
2.250 0.5010 0.01570 0.01003 -0.0765 0.7510 0.5999
2.500 0.5292 0.01575 0.01013 -0.0771 0.7502 0.6011
2.750 0.5584 0.01581 0.01024 -0.0780 0.7494 0.6023
3.000 0.5891 0.01587 0.01035 -0.0791 0.7487 0.6034
3.250 0.6205 0.01595 0.01048 -0.0804 0.7480 0.6046
3.500 0.4888 0.01978 0.01435 -0.0519 0.7288 0.6053
3.750 0.5314 0.01928 0.01392 -0.0549 0.7294 0.6066
4.000 0.5750 0.01879 0.01347 -0.0582 0.7299 0.6079
4.250 0.6210 0.01824 0.01299 -0.0618 0.7302 0.6093
4.500 0.6711 0.01760 0.01240 -0.0663 0.7299 0.6107
4.750 0.7889 0.01485 0.00967 -0.0829 0.7296 0.6123
5.000 0.8485 0.01324 0.00804 -0.0887 0.7222 0.6138
5.250 0.8805 0.01282 0.00769 -0.0897 0.7172 0.6152
5.500 0.8993 0.01260 0.00758 -0.0881 0.7112 0.6166
5.750 0.9386 0.01200 0.00700 -0.0904 0.7046 0.6182
6.000 0.9459 0.01193 0.00704 -0.0866 0.6974 0.6197
6.250 0.9650 0.01160 0.00675 -0.0849 0.6886 0.6215
6.500 0.9677 0.01146 0.00667 -0.0799 0.6798 0.6233
6.750 0.9666 0.01162 0.00691 -0.0745 0.6702 0.6250
7.000 0.9685 0.01184 0.00720 -0.0699 0.6575 0.6266
7.250 0.9675 0.01212 0.00752 -0.0648 0.6361 0.6282
7.500 0.9581 0.01264 0.00782 -0.0582 0.5783 0.6299
7.750 0.9339 0.01410 0.00894 -0.0497 0.5211 0.6314
8.000 0.9084 0.01598 0.01053 -0.0418 0.4670 0.6329
8.250 0.8874 0.01795 0.01225 -0.0350 0.4143 0.6346
8.500 0.8697 0.01999 0.01400 -0.0292 0.3583 0.6363
8.750 0.8578 0.02192 0.01562 -0.0245 0.3013 0.6381
9.000 0.8510 0.02369 0.01708 -0.0207 0.2443 0.6400
9.250 0.8484 0.02533 0.01841 -0.0175 0.1892 0.6417
9.500 0.8513 0.02668 0.01954 -0.0151 0.1434 0.6438
9.750 0.8568 0.02796 0.02065 -0.0131 0.1068 0.6460
10.000 0.8652 0.02911 0.02167 -0.0115 0.0792 0.6483
10.250 0.8743 0.03026 0.02272 -0.0100 0.0576 0.6508
10.500 0.8854 0.03134 0.02375 -0.0088 0.0439 0.6536
10.750 0.8963 0.03246 0.02483 -0.0076 0.0347 0.6563
11.000 0.9086 0.03347 0.02587 -0.0066 0.0282 0.6592
11.250 0.9216 0.03447 0.02694 -0.0056 0.0234 0.6620
11.500 0.9306 0.03579 0.02827 -0.0044 0.0193 0.6651
11.750 0.9452 0.03673 0.02927 -0.0037 0.0163 0.6684
12.000 0.9518 0.03830 0.03086 -0.0023 0.0130 0.6714
12.250 0.9661 0.03929 0.03197 -0.0017 0.0110 0.6748
12.500 0.9764 0.04061 0.03336 -0.0007 0.0097 0.6783
12.750 0.9877 0.04188 0.03469 0.0000 0.0076 0.6822
13.000 0.9983 0.04325 0.03614 0.0008 0.0073 0.6860
13.250 1.0080 0.04468 0.03766 0.0016 0.0064 0.6897
13.500 1.0084 0.04702 0.04012 0.0033 0.0056 0.6934
13.750 1.0217 0.04823 0.04142 0.0036 0.0053 0.6981
14.000 1.0295 0.04996 0.04328 0.0044 0.0049 0.7022
14.250 1.0370 0.05175 0.04519 0.0051 0.0046 0.7066
14.500 1.0458 0.05347 0.04700 0.0056 0.0043 0.7119
14.750 1.0508 0.05561 0.04929 0.0063 0.0043 0.7165
15.000 1.0559 0.05775 0.05156 0.0069 0.0041 0.7214
15.250 1.0562 0.06051 0.05446 0.0077 0.0039 0.7264
15.500 1.0554 0.06353 0.05767 0.0085 0.0037 0.7309
15.750 1.0606 0.06587 0.06020 0.0087 0.0036 0.7366
16.000 1.0648 0.06841 0.06291 0.0088 0.0033 0.7427
16.250 1.0656 0.07137 0.06607 0.0090 0.0033 0.7481
16.500 1.0649 0.07463 0.06953 0.0090 0.0031 0.7540
16.750 1.0637 0.07802 0.07309 0.0087 0.0029 0.7599
17.000 1.0560 0.08246 0.07779 0.0084 0.0030 0.7653
17.500 1.0405 0.09169 0.08746 0.0067 0.0030 0.7771
17.750 1.0295 0.09705 0.09305 0.0052 0.0030 0.7832
18.000 1.0191 0.10250 0.09870 0.0033 0.0030 0.7895
18.250 1.0111 0.10772 0.10409 0.0010 0.0028 0.7972
18.500 1.0002 0.11362 0.11018 -0.0017 0.0027 0.8046
18.750 0.9751 0.12248 0.11934 -0.0057 0.0029 0.8092
19.000 0.9641 0.12895 0.12598 -0.0094 0.0027 0.8176
|
Polar data table (+)
Polar graphs
<< Back to NREL's S827 Airfoil (s827-nr)