NREL's S826 Airfoil (s826-nr) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S826 Airfoil (s826-nr) Reynolds number: 500,000 Max Cl/Cd: 113.97 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s826-nr-500000.txt Download as CSV file: xf-s826-nr-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3176 0.08270 0.08057 -0.0679 0.9870 0.0180
-8.750 -0.3161 0.07273 0.07056 -0.0792 0.9845 0.0182
-8.500 -0.3252 0.06502 0.06276 -0.0878 0.9776 0.0183
-8.250 -0.3274 0.05806 0.05561 -0.0994 0.9719 0.0185
-8.000 -0.3218 0.05208 0.04936 -0.1099 0.9613 0.0187
-7.750 -0.2984 0.04653 0.04338 -0.1202 0.9583 0.0188
-7.500 -0.2989 0.03956 0.03629 -0.1225 0.9491 0.0196
-7.250 -0.2695 0.03649 0.03310 -0.1268 0.9469 0.0202
-7.000 -0.2360 0.03348 0.02989 -0.1314 0.9455 0.0211
-6.750 -0.2182 0.03109 0.02731 -0.1314 0.9368 0.0222
-6.500 -0.1823 0.02080 0.01570 -0.1337 0.9347 0.0105
-6.250 -0.1474 0.01867 0.01344 -0.1357 0.9331 0.0101
-6.000 -0.1243 0.01705 0.01168 -0.1351 0.9259 0.0099
-5.750 -0.0910 0.01562 0.01014 -0.1365 0.9223 0.0099
-5.500 -0.0539 0.01439 0.00883 -0.1389 0.9200 0.0100
-5.250 -0.0266 0.01350 0.00785 -0.1392 0.9116 0.0104
-5.000 0.0125 0.01218 0.00648 -0.1425 0.9075 0.0116
-4.750 0.0480 0.01133 0.00556 -0.1446 0.9001 0.0151
-4.500 0.0899 0.01043 0.00465 -0.1480 0.8944 0.0311
-4.250 0.1329 0.00970 0.00403 -0.1517 0.8872 0.0664
-4.000 0.1896 0.00778 0.00306 -0.1601 0.8808 0.3702
-3.750 0.2353 0.00769 0.00290 -0.1638 0.8687 0.4129
-3.500 0.2769 0.00771 0.00277 -0.1666 0.8523 0.4302
-3.250 0.3125 0.00787 0.00276 -0.1681 0.8324 0.4443
-3.000 0.3426 0.00804 0.00282 -0.1684 0.8112 0.4536
-2.750 0.3716 0.00820 0.00278 -0.1686 0.7917 0.4594
-2.500 0.3988 0.00834 0.00284 -0.1684 0.7736 0.4647
-2.250 0.4259 0.00852 0.00288 -0.1681 0.7566 0.4705
-2.000 0.4530 0.00864 0.00286 -0.1680 0.7412 0.4748
-1.750 0.4798 0.00873 0.00289 -0.1678 0.7274 0.4783
-1.500 0.5068 0.00886 0.00292 -0.1676 0.7151 0.4819
-1.250 0.5342 0.00893 0.00291 -0.1676 0.7035 0.4846
-1.000 0.5618 0.00900 0.00289 -0.1676 0.6932 0.4868
-0.750 0.5895 0.00909 0.00287 -0.1677 0.6840 0.4889
-0.500 0.6170 0.00911 0.00287 -0.1677 0.6748 0.4910
-0.250 0.6445 0.00917 0.00290 -0.1677 0.6667 0.4931
0.000 0.6721 0.00925 0.00295 -0.1677 0.6590 0.4957
0.250 0.7000 0.00934 0.00300 -0.1678 0.6522 0.4985
0.500 0.7279 0.00942 0.00304 -0.1679 0.6456 0.5012
0.750 0.7562 0.00954 0.00309 -0.1681 0.6399 0.5036
1.000 0.7839 0.00955 0.00314 -0.1682 0.6338 0.5062
1.250 0.8117 0.00964 0.00323 -0.1683 0.6283 0.5088
1.500 0.8399 0.00974 0.00333 -0.1684 0.6234 0.5119
1.750 0.8678 0.00981 0.00342 -0.1685 0.6184 0.5152
2.000 0.8959 0.00994 0.00351 -0.1687 0.6135 0.5182
2.250 0.9236 0.01000 0.00362 -0.1688 0.6081 0.5211
2.500 0.9506 0.01006 0.00372 -0.1687 0.6018 0.5241
2.750 0.9780 0.01020 0.00384 -0.1687 0.5954 0.5275
3.000 1.0041 0.01025 0.00394 -0.1684 0.5878 0.5314
3.250 1.0310 0.01038 0.00405 -0.1683 0.5806 0.5351
3.500 1.0571 0.01041 0.00417 -0.1680 0.5735 0.5385
3.750 1.0838 0.01055 0.00431 -0.1679 0.5671 0.5423
4.000 1.1099 0.01062 0.00445 -0.1676 0.5600 0.5465
4.250 1.1361 0.01073 0.00457 -0.1673 0.5530 0.5504
4.500 1.1622 0.01080 0.00474 -0.1671 0.5463 0.5543
4.750 1.1878 0.01092 0.00490 -0.1667 0.5386 0.5589
5.250 1.2379 0.01116 0.00524 -0.1657 0.5212 0.5685
5.500 1.2623 0.01124 0.00542 -0.1651 0.5105 0.5736
6.000 1.3099 0.01151 0.00581 -0.1636 0.4857 0.5837
6.250 1.3323 0.01169 0.00602 -0.1626 0.4679 0.5894
6.500 1.3530 0.01193 0.00625 -0.1613 0.4431 0.5951
6.750 1.3695 0.01237 0.00658 -0.1592 0.3991 0.6006
7.000 1.3771 0.01337 0.00721 -0.1557 0.3317 0.6065
7.250 1.3843 0.01442 0.00797 -0.1522 0.2739 0.6122
7.500 1.3934 0.01541 0.00873 -0.1491 0.2269 0.6189
7.750 1.4033 0.01635 0.00950 -0.1461 0.1885 0.6260
8.000 1.4134 0.01727 0.01028 -0.1433 0.1553 0.6338
8.250 1.4234 0.01819 0.01110 -0.1405 0.1276 0.6419
8.500 1.4335 0.01911 0.01193 -0.1378 0.1051 0.6512
8.750 1.4449 0.01996 0.01275 -0.1354 0.0883 0.6611
9.000 1.4561 0.02082 0.01362 -0.1330 0.0746 0.6726
9.250 1.4667 0.02174 0.01454 -0.1305 0.0625 0.6858
9.500 1.4766 0.02271 0.01553 -0.1281 0.0522 0.7015
9.750 1.4858 0.02375 0.01660 -0.1257 0.0430 0.7211
10.000 1.4950 0.02481 0.01773 -0.1233 0.0354 0.7472
10.250 1.5047 0.02584 0.01889 -0.1211 0.0291 0.7870
10.500 1.5061 0.02664 0.01998 -0.1173 0.0245 1.0000
10.750 1.5122 0.02814 0.02146 -0.1150 0.0194 1.0000
11.000 1.5174 0.02976 0.02309 -0.1128 0.0151 1.0000
11.250 1.5219 0.03150 0.02487 -0.1106 0.0118 1.0000
11.500 1.5249 0.03348 0.02692 -0.1085 0.0094 1.0000
11.750 1.5244 0.03589 0.02940 -0.1063 0.0075 1.0000
12.000 1.5281 0.03801 0.03161 -0.1047 0.0062 1.0000
12.250 1.5228 0.04114 0.03486 -0.1027 0.0054 1.0000
12.500 1.5260 0.04352 0.03736 -0.1015 0.0049 1.0000
12.750 1.5275 0.04618 0.04013 -0.1004 0.0045 1.0000
13.000 1.5283 0.04901 0.04307 -0.0995 0.0042 1.0000
13.250 1.5274 0.05214 0.04631 -0.0987 0.0040 1.0000
13.500 1.5224 0.05588 0.05017 -0.0981 0.0038 1.0000
13.750 1.5124 0.06044 0.05488 -0.0977 0.0037 1.0000
14.000 1.5058 0.06468 0.05925 -0.0975 0.0036 1.0000
14.250 1.5042 0.06837 0.06307 -0.0975 0.0035 1.0000
14.500 1.5017 0.07230 0.06713 -0.0976 0.0035 1.0000
14.750 1.4984 0.07642 0.07140 -0.0980 0.0034 1.0000
15.000 1.4947 0.08070 0.07581 -0.0985 0.0034 1.0000
15.250 1.4902 0.08521 0.08046 -0.0992 0.0033 1.0000
15.500 1.4855 0.08985 0.08525 -0.1001 0.0033 1.0000
15.750 1.4805 0.09464 0.09018 -0.1011 0.0032 1.0000
16.000 1.4749 0.09960 0.09531 -0.1024 0.0032 1.0000
16.250 1.4688 0.10477 0.10064 -0.1039 0.0032 1.0000
16.500 1.4624 0.11007 0.10609 -0.1056 0.0032 1.0000
16.750 1.4553 0.11562 0.11180 -0.1077 0.0031 1.0000
17.000 1.4480 0.12129 0.11764 -0.1100 0.0032 1.0000
17.250 1.4400 0.12723 0.12374 -0.1127 0.0032 1.0000
17.500 1.4310 0.13353 0.13022 -0.1157 0.0032 1.0000
17.750 1.4225 0.13987 0.13673 -0.1191 0.0032 1.0000
18.000 1.4124 0.14675 0.14378 -0.1230 0.0032 1.0000
18.250 1.4025 0.15371 0.15090 -0.1272 0.0032 1.0000
18.500 1.3913 0.16121 0.15858 -0.1320 0.0032 1.0000
18.750 1.3805 0.16885 0.16638 -0.1371 0.0032 1.0000
19.000 1.3701 0.17661 0.17429 -0.1424 0.0033 1.0000
19.250 1.3582 0.18507 0.18290 -0.1483 0.0033 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S826 Airfoil (s826-nr)