NREL's S817 Airfoil (s817-nr) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S817 Airfoil (s817-nr) Reynolds number: 200,000 Max Cl/Cd: 45.28 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s817-nr-200000.txt Download as CSV file: xf-s817-nr-200000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S817 Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.1792 0.10101 0.09698 -0.1049 0.8684 0.0296
-11.000 -0.1799 0.09717 0.09316 -0.1074 0.8648 0.0309
-10.750 -0.1795 0.09355 0.08952 -0.1098 0.8617 0.0313
-10.500 -0.1862 0.08896 0.08493 -0.1144 0.8586 0.0319
-10.250 -0.1937 0.08451 0.08048 -0.1184 0.8553 0.0321
-10.000 -0.1427 0.07195 0.06815 -0.1116 0.8444 0.0339
-9.750 -0.1385 0.06861 0.06483 -0.1119 0.8427 0.0350
-9.500 -0.1530 0.06355 0.05975 -0.1152 0.8408 0.0346
-9.250 -0.1633 0.05962 0.05580 -0.1165 0.8388 0.0349
-9.000 -0.1758 0.05601 0.05215 -0.1170 0.8368 0.0351
-8.750 -0.1901 0.05284 0.04891 -0.1168 0.8347 0.0350
-8.500 -0.2010 0.04998 0.04599 -0.1156 0.8327 0.0357
-8.250 -0.2134 0.04756 0.04348 -0.1133 0.8308 0.0362
-8.000 -0.2224 0.04519 0.04103 -0.1110 0.8288 0.0369
-7.750 -0.2268 0.04265 0.03838 -0.1091 0.8263 0.0380
-7.500 -0.2294 0.04019 0.03578 -0.1069 0.8240 0.0400
-7.250 -0.2342 0.04094 0.03604 -0.1019 0.8214 0.0428
-7.000 -0.2369 0.03724 0.03212 -0.0993 0.8197 0.0434
-6.750 -0.2296 0.03188 0.02686 -0.0995 0.8185 0.0451
-6.500 -0.2197 0.02950 0.02445 -0.0984 0.8172 0.0474
-6.250 -0.2093 0.02766 0.02243 -0.0966 0.8158 0.0516
-6.000 -0.2077 0.02597 0.02039 -0.0932 0.8132 0.0575
-5.750 -0.1984 0.03910 0.03300 -0.0970 0.8154 0.0609
-5.500 -0.1860 0.04181 0.03521 -0.0933 0.8127 0.0694
-5.250 -0.1774 0.03670 0.03017 -0.0924 0.8105 0.0726
-5.000 -0.1628 0.03526 0.02869 -0.0910 0.8087 0.0774
-4.750 -0.1508 0.03422 0.02740 -0.0888 0.8069 0.0875
-4.500 -0.1067 0.02941 0.02158 -0.0859 0.8059 0.0332
-4.250 -0.0785 0.02748 0.01954 -0.0857 0.8048 0.0311
-4.000 -0.0481 0.02584 0.01773 -0.0858 0.8037 0.0307
-3.750 -0.0191 0.02483 0.01655 -0.0855 0.8027 0.0329
-3.500 0.0056 0.02380 0.01559 -0.0852 0.8018 0.0388
-3.250 0.0254 0.02322 0.01495 -0.0837 0.8008 0.0443
-3.000 -0.0085 0.02548 0.01729 -0.0746 0.7950 0.0421
-2.750 -0.0193 0.02604 0.01788 -0.0687 0.7918 0.0450
-2.500 -0.0163 0.02626 0.01813 -0.0650 0.7899 0.0506
-2.250 -0.0018 0.02582 0.01767 -0.0629 0.7885 0.0589
-2.000 0.0197 0.02522 0.01705 -0.0621 0.7874 0.0835
-1.750 -0.0487 0.02846 0.02029 -0.0475 0.7814 0.0623
-1.500 -0.0465 0.02878 0.02061 -0.0441 0.7795 0.0754
-1.250 -0.0425 0.02766 0.02039 -0.0416 0.7781 0.3135
-1.000 -0.0450 0.02784 0.02207 -0.0348 0.7760 0.7538
-0.750 -0.0349 0.02911 0.02336 -0.0303 0.7747 0.7862
-0.500 -0.0247 0.03022 0.02453 -0.0251 0.7735 0.8077
-0.250 0.0002 0.03088 0.02514 -0.0224 0.7722 0.8267
0.000 -0.0183 0.03197 0.02623 -0.0167 0.7687 0.8318
0.250 -0.0191 0.03272 0.02698 -0.0122 0.7659 0.8396
0.500 -0.0110 0.03334 0.02755 -0.0092 0.7642 0.8483
0.750 0.0076 0.03376 0.02790 -0.0075 0.7616 0.8581
1.000 0.0398 0.03381 0.02788 -0.0074 0.7590 0.8637
1.250 0.0874 0.03384 0.02779 -0.0108 0.7574 0.8641
1.500 0.0629 0.03481 0.02879 -0.0049 0.7510 0.8668
1.750 0.0867 0.03516 0.02908 -0.0054 0.7480 0.8678
2.000 0.1260 0.03521 0.02905 -0.0078 0.7452 0.8682
2.250 0.1724 0.03522 0.02899 -0.0111 0.7434 0.8684
2.500 0.1621 0.03608 0.02985 -0.0072 0.7351 0.8712
2.750 0.1958 0.03622 0.02995 -0.0089 0.7318 0.8724
3.000 0.2357 0.03632 0.03002 -0.0113 0.7298 0.8733
3.250 0.2785 0.03637 0.03006 -0.0139 0.7285 0.8739
3.500 0.2672 0.03710 0.03082 -0.0094 0.7179 0.8763
3.750 0.3042 0.03716 0.03088 -0.0112 0.7158 0.8774
4.000 0.3456 0.03716 0.03091 -0.0135 0.7144 0.8785
4.250 0.3411 0.03794 0.03172 -0.0102 0.7037 0.8814
4.500 0.3782 0.03795 0.03175 -0.0120 0.7015 0.8829
4.750 0.4189 0.03790 0.03174 -0.0142 0.7001 0.8844
5.000 0.4206 0.03866 0.03252 -0.0119 0.6891 0.8874
5.250 0.4569 0.03853 0.03247 -0.0133 0.6871 0.8892
5.500 0.4965 0.03829 0.03232 -0.0150 0.6857 0.8911
5.750 0.4971 0.03899 0.03308 -0.0122 0.6744 0.8948
6.000 0.5370 0.03863 0.03282 -0.0139 0.6725 0.8971
6.250 0.5772 0.03823 0.03250 -0.0157 0.6704 0.8993
6.500 0.5879 0.03852 0.03287 -0.0141 0.6595 0.9029
6.750 0.6195 0.03748 0.03193 -0.0140 0.6504 0.9057
7.000 0.7011 0.03302 0.02764 -0.0183 0.6461 0.9065
7.250 0.7675 0.02982 0.02464 -0.0217 0.6446 0.9080
7.500 0.8326 0.02622 0.02123 -0.0246 0.6415 0.9095
7.750 0.8591 0.02501 0.02016 -0.0238 0.6313 0.9139
8.000 0.8761 0.02428 0.01959 -0.0218 0.6181 0.9190
8.250 0.8868 0.02402 0.01947 -0.0193 0.6005 0.9258
8.500 0.8886 0.02465 0.02021 -0.0162 0.5726 0.9343
8.750 0.9337 0.02062 0.01484 -0.0138 0.3661 0.9385
9.000 0.9093 0.02328 0.01701 -0.0084 0.2823 0.9525
9.250 0.8943 0.02632 0.01949 -0.0057 0.1905 0.9698
9.500 0.8900 0.02877 0.02149 -0.0043 0.1309 1.0000
9.750 0.8937 0.03073 0.02325 -0.0032 0.1050 1.0000
10.000 0.9022 0.03240 0.02485 -0.0025 0.0902 1.0000
10.250 0.9118 0.03401 0.02638 -0.0019 0.0802 1.0000
10.500 0.9268 0.03525 0.02767 -0.0016 0.0722 1.0000
10.750 0.9396 0.03665 0.02901 -0.0008 0.0659 1.0000
11.000 0.9577 0.03769 0.03014 -0.0006 0.0611 1.0000
11.250 0.9758 0.03882 0.03122 -0.0002 0.0559 1.0000
11.500 0.9985 0.03974 0.03224 0.0000 0.0520 1.0000
11.750 1.0194 0.04076 0.03334 0.0002 0.0481 1.0000
12.000 1.0539 0.04186 0.03441 -0.0003 0.0433 1.0000
12.250 1.0739 0.04312 0.03590 0.0000 0.0407 1.0000
12.500 1.0875 0.04448 0.03738 0.0005 0.0374 1.0000
12.750 1.1266 0.04686 0.03987 -0.0006 0.0333 1.0000
13.000 1.1302 0.04847 0.04176 0.0008 0.0315 1.0000
13.250 1.1375 0.05026 0.04374 0.0017 0.0290 1.0000
13.500 1.1422 0.05182 0.04530 0.0024 0.0266 1.0000
13.750 1.1460 0.05501 0.04875 0.0035 0.0241 1.0000
14.000 1.1448 0.05760 0.05164 0.0048 0.0221 1.0000
14.250 1.1451 0.05936 0.05347 0.0054 0.0204 1.0000
14.500 1.1456 0.06248 0.05659 0.0058 0.0183 1.0000
14.750 1.1343 0.06593 0.06042 0.0071 0.0173 1.0000
15.000 1.1243 0.06981 0.06462 0.0079 0.0161 1.0000
15.250 1.1174 0.07348 0.06849 0.0082 0.0152 1.0000
15.500 1.1132 0.07658 0.07169 0.0081 0.0145 1.0000
16.000 1.0902 0.08575 0.08115 0.0071 0.0132 1.0000
16.250 1.0704 0.09228 0.08794 0.0061 0.0133 1.0000
16.500 1.0517 0.09815 0.09409 0.0041 0.0130 1.0000
16.750 1.0313 0.10510 0.10131 0.0013 0.0127 1.0000
17.000 1.0129 0.11244 0.10876 -0.0013 0.0133 1.0000
17.250 0.9915 0.12058 0.11712 -0.0058 0.0132 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S817 Airfoil (s817-nr)