NREL's S820 Airfoil (s820-nr) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S820 Airfoil (s820-nr) Reynolds number: 1,000,000 Max Cl/Cd: 109.84 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s820-nr-1000000-n5.txt Download as CSV file: xf-s820-nr-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S820 Airfoil
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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.2859 0.08197 0.07937 -0.0962 0.7902 0.0047
-11.250 -0.2955 0.07369 0.07104 -0.1018 0.7870 0.0047
-11.000 -0.3148 0.06681 0.06408 -0.1054 0.7838 0.0047
-10.750 -0.3296 0.06218 0.05937 -0.1074 0.7806 0.0047
-10.500 -0.3486 0.05761 0.05470 -0.1083 0.7771 0.0047
-10.250 -0.3637 0.05411 0.05109 -0.1082 0.7736 0.0047
-10.000 -0.3778 0.05103 0.04789 -0.1073 0.7703 0.0047
-9.750 -0.3933 0.04794 0.04469 -0.1057 0.7673 0.0047
-9.500 -0.4039 0.04557 0.04221 -0.1035 0.7642 0.0047
-9.250 -0.4165 0.04326 0.03977 -0.1003 0.7611 0.0047
-9.000 -0.4255 0.04130 0.03767 -0.0967 0.7583 0.0047
-8.000 -0.4390 0.02609 0.02139 -0.0838 0.7485 0.0027
-7.750 -0.4271 0.02311 0.01810 -0.0819 0.7465 0.0026
-7.500 -0.4093 0.02060 0.01530 -0.0805 0.7446 0.0025
-7.250 -0.3880 0.01841 0.01283 -0.0796 0.7428 0.0024
-7.000 -0.3651 0.01683 0.01105 -0.0790 0.7411 0.0024
-6.750 -0.3420 0.01579 0.00988 -0.0784 0.7395 0.0024
-6.500 -0.3198 0.01486 0.00885 -0.0777 0.7376 0.0024
-6.250 -0.2995 0.01395 0.00784 -0.0767 0.7355 0.0024
-6.000 -0.2788 0.01330 0.00711 -0.0758 0.7334 0.0024
-5.750 -0.2582 0.01270 0.00644 -0.0749 0.7314 0.0024
-5.500 -0.2378 0.01211 0.00576 -0.0739 0.7296 0.0024
-5.250 -0.2157 0.01167 0.00526 -0.0732 0.7278 0.0024
-5.000 -0.1929 0.01124 0.00478 -0.0727 0.7262 0.0025
-4.750 -0.1690 0.01087 0.00437 -0.0723 0.7246 0.0025
-4.500 -0.1446 0.01052 0.00399 -0.0720 0.7228 0.0026
-4.250 -0.1194 0.01022 0.00365 -0.0718 0.7209 0.0027
-4.000 -0.0936 0.00999 0.00338 -0.0718 0.7190 0.0028
-3.750 -0.0681 0.00970 0.00305 -0.0716 0.7170 0.0032
-3.500 -0.0416 0.00952 0.00284 -0.0717 0.7151 0.0039
-3.250 -0.0147 0.00937 0.00266 -0.0718 0.7134 0.0047
-3.000 0.0124 0.00921 0.00250 -0.0719 0.7118 0.0072
-2.750 0.0394 0.00902 0.00236 -0.0721 0.7102 0.0178
-2.500 0.0501 0.00689 0.00165 -0.0707 0.7082 0.4186
-2.250 0.0727 0.00617 0.00156 -0.0704 0.7063 0.6089
-2.000 0.1016 0.00617 0.00157 -0.0708 0.7044 0.6244
-1.750 0.1306 0.00617 0.00154 -0.0712 0.7025 0.6287
-1.500 0.1597 0.00617 0.00149 -0.0717 0.7006 0.6309
-1.250 0.1888 0.00619 0.00146 -0.0722 0.6988 0.6331
-1.000 0.2181 0.00619 0.00144 -0.0727 0.6972 0.6350
-0.750 0.2474 0.00620 0.00143 -0.0733 0.6955 0.6368
-0.500 0.2765 0.00619 0.00142 -0.0738 0.6935 0.6386
-0.250 0.3056 0.00619 0.00142 -0.0742 0.6914 0.6406
0.000 0.3346 0.00619 0.00142 -0.0747 0.6894 0.6428
0.250 0.3636 0.00621 0.00142 -0.0752 0.6875 0.6450
0.500 0.3926 0.00623 0.00143 -0.0757 0.6856 0.6472
0.750 0.4216 0.00626 0.00145 -0.0762 0.6837 0.6493
1.000 0.4507 0.00628 0.00147 -0.0767 0.6817 0.6511
1.250 0.4796 0.00628 0.00150 -0.0772 0.6795 0.6531
1.500 0.5083 0.00629 0.00153 -0.0776 0.6763 0.6551
1.750 0.5364 0.00631 0.00154 -0.0779 0.6711 0.6572
2.000 0.5644 0.00632 0.00156 -0.0782 0.6622 0.6597
2.250 0.5917 0.00635 0.00157 -0.0783 0.6518 0.6621
2.500 0.6189 0.00641 0.00159 -0.0784 0.6424 0.6645
2.750 0.6467 0.00646 0.00165 -0.0786 0.6330 0.6665
3.000 0.6737 0.00651 0.00171 -0.0787 0.6233 0.6687
3.250 0.7003 0.00659 0.00178 -0.0788 0.6131 0.6709
3.500 0.7266 0.00669 0.00186 -0.0787 0.6004 0.6733
3.750 0.7513 0.00684 0.00197 -0.0784 0.5789 0.6757
4.000 0.7725 0.00712 0.00212 -0.0774 0.5428 0.6781
4.250 0.7893 0.00759 0.00237 -0.0755 0.4923 0.6803
4.500 0.8020 0.00820 0.00274 -0.0730 0.4355 0.6826
4.750 0.8139 0.00881 0.00313 -0.0703 0.3820 0.6852
5.000 0.8284 0.00927 0.00344 -0.0681 0.3437 0.6881
5.250 0.8382 0.00980 0.00379 -0.0651 0.3012 0.6910
5.500 0.8482 0.01022 0.00409 -0.0620 0.2714 0.6938
5.750 0.8592 0.01071 0.00447 -0.0592 0.2428 0.6962
6.000 0.8726 0.01115 0.00484 -0.0569 0.2205 0.6989
6.250 0.8847 0.01165 0.00526 -0.0544 0.1963 0.7017
6.500 0.8981 0.01214 0.00568 -0.0523 0.1750 0.7046
6.750 0.9122 0.01263 0.00611 -0.0504 0.1574 0.7075
7.000 0.9258 0.01316 0.00658 -0.0484 0.1400 0.7102
7.250 0.9403 0.01367 0.00705 -0.0467 0.1252 0.7131
7.500 0.9549 0.01419 0.00754 -0.0450 0.1124 0.7164
7.750 0.9688 0.01477 0.00808 -0.0433 0.0982 0.7199
8.000 0.9835 0.01535 0.00863 -0.0418 0.0872 0.7233
8.250 0.9989 0.01591 0.00918 -0.0404 0.0780 0.7263
8.500 1.0122 0.01658 0.00982 -0.0387 0.0659 0.7297
8.750 1.0282 0.01714 0.01039 -0.0375 0.0599 0.7333
9.000 1.0417 0.01785 0.01106 -0.0360 0.0502 0.7370
9.250 1.0556 0.01857 0.01175 -0.0346 0.0419 0.7405
9.500 1.0709 0.01920 0.01241 -0.0335 0.0364 0.7443
9.750 1.0837 0.02001 0.01320 -0.0320 0.0283 0.7488
10.000 1.0958 0.02087 0.01403 -0.0305 0.0207 0.7533
10.250 1.1092 0.02168 0.01484 -0.0293 0.0164 0.7579
10.500 1.1220 0.02253 0.01571 -0.0280 0.0124 0.7630
10.750 1.1354 0.02337 0.01657 -0.0268 0.0094 0.7681
11.000 1.1498 0.02415 0.01740 -0.0257 0.0082 0.7734
11.250 1.1618 0.02511 0.01838 -0.0245 0.0054 0.7799
11.500 1.1759 0.02593 0.01928 -0.0235 0.0049 0.7863
11.750 1.1887 0.02685 0.02025 -0.0224 0.0039 0.7938
12.000 1.2021 0.02775 0.02121 -0.0214 0.0031 0.8012
12.250 1.2144 0.02874 0.02227 -0.0204 0.0026 0.8100
12.500 1.2272 0.02970 0.02331 -0.0194 0.0019 0.8197
12.750 1.2391 0.03072 0.02442 -0.0184 0.0016 0.8321
13.000 1.2487 0.03191 0.02570 -0.0172 0.0008 0.8478
13.250 1.2591 0.03299 0.02691 -0.0160 0.0007 0.8708
13.500 1.2704 0.03397 0.02818 -0.0150 0.0007 0.9519
13.750 1.2831 0.03525 0.02953 -0.0147 0.0005 1.0000
14.000 1.2936 0.03658 0.03092 -0.0140 0.0005 1.0000
14.250 1.3036 0.03799 0.03239 -0.0133 0.0004 1.0000
14.500 1.3113 0.03962 0.03410 -0.0125 0.0003 1.0000
14.750 1.3212 0.04109 0.03563 -0.0119 0.0003 1.0000
15.000 1.3290 0.04276 0.03737 -0.0112 0.0003 1.0000
15.250 1.3326 0.04492 0.03964 -0.0104 0.0002 1.0000
15.500 1.3399 0.04675 0.04154 -0.0099 0.0001 1.0000
15.750 1.3464 0.04870 0.04357 -0.0095 0.0001 1.0000
16.000 1.3503 0.05096 0.04594 -0.0091 0.0001 1.0000
16.250 1.3561 0.05306 0.04813 -0.0089 0.0001 1.0000
16.500 1.3589 0.05556 0.05072 -0.0086 0.0001 1.0000
16.750 1.3604 0.05830 0.05357 -0.0085 0.0001 1.0000
17.000 1.3638 0.06085 0.05621 -0.0086 0.0001 1.0000
17.250 1.3652 0.06371 0.05917 -0.0088 0.0001 1.0000
17.500 1.3651 0.06683 0.06239 -0.0091 0.0001 1.0000
17.750 1.3624 0.07042 0.06610 -0.0096 0.0001 1.0000
18.000 1.3577 0.07438 0.07020 -0.0104 0.0001 1.0000
18.250 1.3508 0.07884 0.07479 -0.0115 0.0001 1.0000
18.500 1.3501 0.08243 0.07847 -0.0126 0.0001 1.0000
18.750 1.3397 0.08769 0.08388 -0.0143 0.0001 1.0000
19.000 1.3321 0.09269 0.08900 -0.0162 0.0001 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S820 Airfoil (s820-nr)