NREL's S814 Airfoil (s814-nr) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S814 Airfoil (s814-nr) Reynolds number: 100,000 Max Cl/Cd: 40.14 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s814-nr-100000-n5.txt Download as CSV file: xf-s814-nr-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S814 Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1705 0.10748 0.09970 -0.0389 0.9824 0.2912
-8.000 -0.1495 0.10491 0.09710 -0.0422 0.9775 0.2926
-7.750 -0.1317 0.10230 0.09446 -0.0452 0.9713 0.2942
-7.500 -0.1307 0.09889 0.09098 -0.0493 0.9656 0.2989
-7.250 -0.1019 0.09727 0.08939 -0.0507 0.9572 0.2998
-7.000 -0.0723 0.09526 0.08737 -0.0536 0.9516 0.3007
-6.750 -0.0540 0.09340 0.08552 -0.0545 0.9403 0.3014
-6.500 -0.0374 0.09161 0.08375 -0.0551 0.9275 0.3023
-6.250 -0.0216 0.08985 0.08200 -0.0557 0.9133 0.3032
-6.000 -0.0062 0.08799 0.08015 -0.0564 0.8985 0.3042
-5.750 0.0121 0.08613 0.07829 -0.0577 0.8847 0.3058
-5.500 0.0321 0.08398 0.07611 -0.0598 0.8724 0.3080
-5.250 0.0525 0.08142 0.07350 -0.0629 0.8606 0.3095
-5.000 0.0721 0.07870 0.07071 -0.0662 0.8467 0.3105
-4.750 0.0917 0.07565 0.06757 -0.0705 0.8327 0.3117
-4.500 0.1096 0.07219 0.06398 -0.0760 0.8179 0.3139
-4.250 0.1615 0.07052 0.06223 -0.0814 0.8020 0.3149
-4.000 0.2058 0.06885 0.06046 -0.0861 0.7838 0.3159
-3.750 0.2412 0.06723 0.05871 -0.0898 0.7645 0.3169
-3.500 0.2669 0.06558 0.05694 -0.0923 0.7456 0.3176
-3.250 0.2878 0.06407 0.05533 -0.0940 0.7281 0.3183
-3.000 0.3055 0.06266 0.05382 -0.0950 0.7119 0.3189
-2.750 0.3213 0.06131 0.05240 -0.0957 0.6974 0.3197
-2.500 0.3352 0.05997 0.05098 -0.0962 0.6848 0.3206
-2.250 0.3467 0.05866 0.04961 -0.0962 0.6731 0.3216
-2.000 0.3546 0.05736 0.04828 -0.0958 0.6627 0.3232
-1.750 0.3569 0.05571 0.04655 -0.0955 0.6538 0.3253
-1.500 0.3482 0.05358 0.04440 -0.0942 0.6453 0.3268
-1.250 0.3112 0.04991 0.04066 -0.0917 0.6387 0.3288
-1.000 0.3282 0.04936 0.04011 -0.0908 0.6309 0.3294
-0.750 0.3426 0.04876 0.03952 -0.0898 0.6232 0.3300
-0.500 0.3575 0.04805 0.03879 -0.0891 0.6166 0.3307
-0.250 0.3681 0.04731 0.03808 -0.0879 0.6093 0.3316
0.000 0.3804 0.04643 0.03718 -0.0872 0.6032 0.3326
0.250 0.3910 0.04552 0.03629 -0.0865 0.5974 0.3340
0.500 0.3986 0.04443 0.03523 -0.0859 0.5916 0.3360
0.750 0.3663 0.03429 0.02471 -0.0973 0.5878 0.3460
1.250 0.4141 0.03332 0.02376 -0.0984 0.5780 0.3483
1.500 0.4383 0.03263 0.02308 -0.0994 0.5730 0.3500
1.750 0.4665 0.03163 0.02201 -0.1017 0.5686 0.3523
2.000 0.5028 0.02931 0.01943 -0.1083 0.5641 0.3582
2.250 0.5323 0.02847 0.01854 -0.1107 0.5590 0.3610
2.500 0.5584 0.02851 0.01866 -0.1106 0.5547 0.3620
2.750 0.5870 0.02851 0.01871 -0.1110 0.5512 0.3633
3.000 0.6167 0.02848 0.01870 -0.1117 0.5480 0.3647
3.250 0.6412 0.02841 0.01874 -0.1118 0.5437 0.3663
3.500 0.6688 0.02827 0.01865 -0.1125 0.5395 0.3683
3.750 0.7003 0.02802 0.01839 -0.1141 0.5357 0.3708
4.000 0.7362 0.02766 0.01794 -0.1166 0.5326 0.3739
4.250 0.7687 0.02743 0.01767 -0.1186 0.5293 0.3767
4.500 0.7906 0.02766 0.01809 -0.1177 0.5253 0.3781
4.750 0.8151 0.02786 0.01843 -0.1173 0.5214 0.3796
5.000 0.8428 0.02803 0.01870 -0.1174 0.5179 0.3816
5.250 0.8743 0.02815 0.01886 -0.1183 0.5148 0.3841
5.500 0.8974 0.02833 0.01916 -0.1180 0.5103 0.3869
5.750 0.9228 0.02843 0.01934 -0.1182 0.5055 0.3902
6.000 0.9537 0.02843 0.01933 -0.1193 0.5012 0.3934
6.250 0.9840 0.02855 0.01953 -0.1196 0.4975 0.3952
6.500 1.0019 0.02889 0.02008 -0.1182 0.4923 0.3969
6.750 1.0225 0.02912 0.02047 -0.1171 0.4867 0.3991
7.000 1.0529 0.02913 0.02051 -0.1176 0.4814 0.4019
7.250 1.0734 0.02932 0.02081 -0.1166 0.4751 0.4048
7.500 1.0957 0.02940 0.02095 -0.1160 0.4678 0.4085
7.750 1.1248 0.02933 0.02091 -0.1161 0.4612 0.4114
8.000 1.1338 0.02968 0.02150 -0.1131 0.4531 0.4134
8.250 1.1617 0.02956 0.02139 -0.1128 0.4456 0.4166
8.500 1.1676 0.02995 0.02200 -0.1095 0.4369 0.4193
8.750 1.1921 0.02982 0.02186 -0.1088 0.4286 0.4233
9.000 1.1950 0.03023 0.02241 -0.1052 0.4197 0.4264
9.250 1.2134 0.03023 0.02248 -0.1035 0.4110 0.4288
9.500 1.2140 0.03093 0.02342 -0.0997 0.4008 0.4308
9.750 1.2240 0.03134 0.02395 -0.0972 0.3904 0.4334
10.000 1.2315 0.03188 0.02459 -0.0946 0.3786 0.4364
10.250 1.2326 0.03283 0.02569 -0.0915 0.3651 0.4394
10.500 1.2343 0.03387 0.02683 -0.0889 0.3500 0.4428
10.750 1.2340 0.03508 0.02814 -0.0861 0.3327 0.4453
11.000 1.2324 0.03647 0.02957 -0.0835 0.3125 0.4476
11.250 1.2282 0.03826 0.03133 -0.0811 0.2899 0.4501
11.500 1.2210 0.04051 0.03347 -0.0789 0.2681 0.4527
11.750 1.2111 0.04331 0.03618 -0.0772 0.2482 0.4553
12.000 1.2005 0.04654 0.03936 -0.0760 0.2294 0.4579
12.250 1.1899 0.05009 0.04284 -0.0753 0.2122 0.4605
12.500 1.1792 0.05380 0.04656 -0.0749 0.1968 0.4622
12.750 1.1693 0.05764 0.05041 -0.0747 0.1827 0.4641
13.000 1.1604 0.06155 0.05431 -0.0748 0.1696 0.4663
13.250 1.1542 0.06534 0.05811 -0.0751 0.1570 0.4688
13.500 1.1499 0.06904 0.06184 -0.0756 0.1454 0.4718
13.750 1.1467 0.07274 0.06552 -0.0762 0.1351 0.4751
14.000 1.1433 0.07652 0.06926 -0.0769 0.1257 0.4783
14.250 1.1433 0.07992 0.07275 -0.0775 0.1164 0.4811
14.500 1.1425 0.08346 0.07630 -0.0782 0.1089 0.4841
14.750 1.1437 0.08686 0.07976 -0.0790 0.1013 0.4878
15.000 1.1451 0.09029 0.08318 -0.0799 0.0950 0.4918
15.250 1.1477 0.09362 0.08656 -0.0808 0.0888 0.4957
15.500 1.1489 0.09706 0.09001 -0.0817 0.0839 0.4989
15.750 1.1534 0.10018 0.09326 -0.0826 0.0785 0.5032
16.000 1.1551 0.10370 0.09678 -0.0838 0.0744 0.5077
16.250 1.1600 0.10678 0.09992 -0.0847 0.0704 0.5127
|
Polar data table (+)
Polar graphs
<< Back to NREL's S814 Airfoil (s814-nr)