NREL's S826 Airfoil (s826-nr) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S826 Airfoil (s826-nr) Reynolds number: 1,000,000 Max Cl/Cd: 126.69 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s826-nr-1000000-n5.txt Download as CSV file: xf-s826-nr-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S826 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3594 0.04326 0.04108 -0.1254 0.9730 0.0031
-10.000 -0.3839 0.03675 0.03432 -0.1293 0.9667 0.0031
-9.750 -0.4009 0.03067 0.02791 -0.1339 0.9607 0.0030
-9.500 -0.3820 0.02505 0.02182 -0.1434 0.9586 0.0030
-9.250 -0.3887 0.02327 0.01986 -0.1400 0.9497 0.0030
-9.000 -0.3649 0.02088 0.01719 -0.1426 0.9473 0.0029
-8.750 -0.3365 0.01887 0.01494 -0.1451 0.9457 0.0029
-8.000 -0.2920 0.01562 0.01129 -0.1409 0.9215 0.0029
-7.750 -0.2614 0.01447 0.01000 -0.1426 0.9146 0.0029
-7.500 -0.2208 0.01340 0.00879 -0.1462 0.9076 0.0029
-7.250 -0.1716 0.01240 0.00765 -0.1518 0.8994 0.0030
-7.000 -0.1253 0.01154 0.00663 -0.1566 0.8828 0.0030
-6.750 -0.0900 0.01101 0.00590 -0.1588 0.8567 0.0030
-6.500 -0.0618 0.01064 0.00533 -0.1594 0.8293 0.0031
-6.000 -0.0088 0.01010 0.00447 -0.1597 0.7832 0.0032
-5.750 0.0180 0.00984 0.00406 -0.1599 0.7632 0.0033
-5.500 0.0450 0.00961 0.00369 -0.1601 0.7450 0.0035
-5.250 0.0726 0.00942 0.00338 -0.1604 0.7286 0.0039
-5.000 0.1003 0.00926 0.00313 -0.1606 0.7138 0.0047
-4.750 0.1284 0.00908 0.00290 -0.1610 0.7006 0.0081
-4.500 0.1567 0.00894 0.00271 -0.1614 0.6885 0.0139
-4.250 0.1851 0.00883 0.00255 -0.1617 0.6775 0.0186
-4.000 0.2137 0.00867 0.00239 -0.1622 0.6672 0.0315
-3.750 0.2425 0.00851 0.00224 -0.1626 0.6572 0.0498
-3.500 0.2720 0.00827 0.00209 -0.1634 0.6480 0.0907
-3.250 0.3018 0.00799 0.00192 -0.1642 0.6391 0.1501
-3.000 0.3354 0.00707 0.00155 -0.1667 0.6302 0.3558
-2.750 0.3639 0.00706 0.00154 -0.1669 0.6225 0.3847
-2.500 0.3922 0.00708 0.00154 -0.1671 0.6151 0.3977
-2.250 0.4205 0.00711 0.00153 -0.1672 0.6088 0.4036
-2.000 0.4488 0.00715 0.00153 -0.1674 0.6024 0.4111
-1.500 0.5053 0.00722 0.00155 -0.1677 0.5911 0.4216
-1.250 0.5335 0.00727 0.00155 -0.1678 0.5854 0.4251
-1.000 0.5614 0.00734 0.00157 -0.1678 0.5801 0.4286
-0.750 0.5898 0.00736 0.00159 -0.1680 0.5751 0.4313
-0.500 0.6179 0.00741 0.00161 -0.1681 0.5702 0.4332
-0.250 0.6459 0.00747 0.00164 -0.1682 0.5660 0.4352
0.000 0.6742 0.00750 0.00168 -0.1684 0.5622 0.4375
0.250 0.7024 0.00754 0.00172 -0.1685 0.5582 0.4398
0.500 0.7303 0.00761 0.00176 -0.1686 0.5541 0.4420
0.750 0.7581 0.00768 0.00181 -0.1687 0.5503 0.4438
1.000 0.7864 0.00771 0.00186 -0.1688 0.5468 0.4460
1.250 0.8144 0.00776 0.00193 -0.1689 0.5433 0.4484
1.500 0.8421 0.00783 0.00200 -0.1690 0.5392 0.4511
1.750 0.8697 0.00790 0.00207 -0.1690 0.5341 0.4539
2.000 0.8970 0.00798 0.00214 -0.1690 0.5263 0.4564
2.250 0.9240 0.00808 0.00222 -0.1689 0.5180 0.4585
2.500 0.9510 0.00818 0.00231 -0.1688 0.5091 0.4611
2.750 0.9783 0.00827 0.00241 -0.1688 0.5015 0.4638
3.000 1.0051 0.00837 0.00251 -0.1687 0.4941 0.4667
3.250 1.0322 0.00846 0.00262 -0.1686 0.4867 0.4699
3.500 1.0585 0.00860 0.00273 -0.1684 0.4772 0.4727
3.750 1.0855 0.00869 0.00285 -0.1683 0.4681 0.4756
4.000 1.1119 0.00881 0.00298 -0.1681 0.4583 0.4787
4.250 1.1377 0.00898 0.00313 -0.1678 0.4434 0.4818
4.500 1.1620 0.00924 0.00330 -0.1673 0.4175 0.4849
4.750 1.1826 0.00979 0.00362 -0.1661 0.3656 0.4877
5.000 1.2014 0.01051 0.00407 -0.1646 0.3115 0.4904
5.250 1.2218 0.01108 0.00447 -0.1634 0.2716 0.4937
5.500 1.2415 0.01168 0.00489 -0.1622 0.2323 0.4973
5.750 1.2612 0.01225 0.00531 -0.1608 0.1988 0.5009
6.250 1.3017 0.01321 0.00609 -0.1584 0.1501 0.5078
6.500 1.3210 0.01371 0.00650 -0.1570 0.1293 0.5114
6.750 1.3389 0.01419 0.00692 -0.1553 0.1101 0.5151
7.000 1.3557 0.01467 0.00733 -0.1534 0.0944 0.5183
7.250 1.3712 0.01524 0.00783 -0.1513 0.0760 0.5219
7.500 1.3856 0.01587 0.00838 -0.1490 0.0580 0.5258
7.750 1.3999 0.01650 0.00894 -0.1468 0.0445 0.5298
8.000 1.4154 0.01708 0.00948 -0.1448 0.0351 0.5336
8.250 1.4320 0.01759 0.01002 -0.1430 0.0295 0.5378
8.500 1.4472 0.01818 0.01061 -0.1410 0.0238 0.5426
8.750 1.4619 0.01881 0.01124 -0.1390 0.0188 0.5476
9.000 1.4758 0.01948 0.01192 -0.1369 0.0146 0.5526
9.250 1.4897 0.02017 0.01264 -0.1349 0.0112 0.5584
9.500 1.5034 0.02088 0.01338 -0.1328 0.0087 0.5638
9.750 1.5162 0.02166 0.01421 -0.1307 0.0067 0.5694
10.000 1.5286 0.02248 0.01508 -0.1286 0.0051 0.5755
10.250 1.5405 0.02334 0.01600 -0.1265 0.0040 0.5815
10.500 1.5524 0.02424 0.01696 -0.1245 0.0032 0.5886
10.750 1.5632 0.02523 0.01801 -0.1224 0.0025 0.5955
11.000 1.5738 0.02625 0.01912 -0.1204 0.0021 0.6036
11.250 1.5838 0.02738 0.02032 -0.1185 0.0016 0.6122
11.500 1.5926 0.02862 0.02166 -0.1165 0.0013 0.6224
11.750 1.6013 0.02992 0.02306 -0.1147 0.0011 0.6336
12.000 1.6100 0.03129 0.02453 -0.1129 0.0010 0.6457
12.250 1.6180 0.03276 0.02611 -0.1113 0.0009 0.6591
12.500 1.6252 0.03436 0.02782 -0.1097 0.0008 0.6739
12.750 1.6321 0.03606 0.02963 -0.1083 0.0007 0.6913
13.250 1.6435 0.03991 0.03377 -0.1057 0.0006 0.7419
13.500 1.6485 0.04203 0.03608 -0.1046 0.0006 0.7848
13.750 1.6485 0.04397 0.03839 -0.1027 0.0006 1.0000
14.000 1.6510 0.04650 0.04103 -0.1018 0.0005 1.0000
14.250 1.6520 0.04929 0.04393 -0.1009 0.0005 1.0000
14.500 1.6525 0.05223 0.04698 -0.1002 0.0005 1.0000
14.750 1.6541 0.05512 0.04997 -0.0998 0.0005 1.0000
15.000 1.6552 0.05815 0.05310 -0.0994 0.0005 1.0000
15.250 1.6554 0.06137 0.05643 -0.0993 0.0005 1.0000
15.500 1.6548 0.06479 0.05996 -0.0992 0.0005 1.0000
15.750 1.6533 0.06842 0.06370 -0.0993 0.0005 1.0000
16.000 1.6512 0.07224 0.06763 -0.0996 0.0005 1.0000
16.250 1.6479 0.07634 0.07185 -0.1001 0.0005 1.0000
16.500 1.6437 0.08064 0.07627 -0.1008 0.0005 1.0000
16.750 1.6385 0.08523 0.08098 -0.1017 0.0005 1.0000
17.000 1.6331 0.08996 0.08583 -0.1028 0.0005 1.0000
17.250 1.6265 0.09498 0.09098 -0.1042 0.0005 1.0000
17.500 1.6184 0.10036 0.09650 -0.1058 0.0005 1.0000
17.750 1.6109 0.10574 0.10201 -0.1076 0.0005 1.0000
18.000 1.6003 0.11181 0.10822 -0.1098 0.0005 1.0000
18.250 1.5907 0.11783 0.11437 -0.1123 0.0005 1.0000
18.500 1.5789 0.12438 0.12106 -0.1152 0.0005 1.0000
18.750 1.5678 0.13095 0.12777 -0.1184 0.0005 1.0000
19.000 1.5568 0.13760 0.13455 -0.1219 0.0005 1.0000
19.250 1.5431 0.14501 0.14210 -0.1260 0.0005 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S826 Airfoil (s826-nr)