NREL's S832 Airfoil (s832-nr) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S832 Airfoil (s832-nr) Reynolds number: 500,000 Max Cl/Cd: 126.01 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s832-nr-500000.txt Download as CSV file: xf-s832-nr-500000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S832 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
------ -------- --------- --------- -------- -------- --------
-9.750 0.0188 0.08673 0.08339 -0.1268 0.8019 0.0109
-9.500 0.0236 0.08390 0.08057 -0.1280 0.7985 0.0111
-9.250 0.0282 0.08115 0.07781 -0.1291 0.7953 0.0112
-9.000 0.0320 0.07826 0.07489 -0.1304 0.7923 0.0115
-8.750 0.0355 0.07542 0.07206 -0.1317 0.7893 0.0116
-8.500 0.0382 0.07251 0.06918 -0.1331 0.7859 0.0118
-8.250 0.0391 0.06936 0.06605 -0.1347 0.7824 0.0123
-8.000 0.0390 0.06627 0.06294 -0.1366 0.7792 0.0122
-7.750 0.0343 0.06294 0.05961 -0.1391 0.7758 0.0123
-7.500 0.0251 0.06012 0.05680 -0.1396 0.7720 0.0125
-7.250 0.0237 0.05711 0.05375 -0.1403 0.7687 0.0127
-7.000 0.0260 0.05400 0.05055 -0.1410 0.7658 0.0131
-6.750 0.0328 0.05087 0.04726 -0.1414 0.7632 0.0135
-6.500 0.0435 0.04782 0.04406 -0.1411 0.7605 0.0138
-6.250 0.0524 0.02761 0.02396 -0.1325 0.7511 0.0140
-6.000 0.0570 0.02447 0.02073 -0.1319 0.7489 0.0144
-5.750 0.0679 0.02217 0.01836 -0.1314 0.7466 0.0147
-5.500 0.0802 0.01995 0.01606 -0.1307 0.7440 0.0152
-5.250 0.0947 0.01772 0.01370 -0.1301 0.7415 0.0158
-5.000 0.1108 0.01546 0.01125 -0.1293 0.7392 0.0169
-4.750 0.1353 0.01404 0.00945 -0.1277 0.7371 0.0187
-4.500 0.1457 0.01051 0.00562 -0.1265 0.7352 0.0195
-4.250 0.1648 0.00919 0.00421 -0.1263 0.7331 0.0204
-4.000 0.1855 0.00811 0.00302 -0.1258 0.7308 0.0218
-3.750 0.2128 0.00807 0.00264 -0.1246 0.7284 0.0254
-3.500 0.2439 0.01882 0.01264 -0.1300 0.7319 0.0176
-3.250 0.2687 0.01702 0.01060 -0.1292 0.7296 0.0171
-3.000 0.2950 0.01562 0.00900 -0.1287 0.7274 0.0170
-2.750 0.3209 0.01375 0.00687 -0.1278 0.7254 0.0150
-2.500 0.3462 0.01295 0.00596 -0.1273 0.7234 0.0148
-2.250 0.3713 0.01237 0.00532 -0.1269 0.7216 0.0154
-2.000 0.3971 0.01201 0.00488 -0.1267 0.7196 0.0168
-1.750 0.4206 0.01157 0.00444 -0.1260 0.7174 0.0206
-1.500 0.4457 0.01130 0.00424 -0.1258 0.7151 0.0404
-1.250 0.4721 0.01118 0.00415 -0.1258 0.7129 0.0589
-1.000 0.4977 0.01085 0.00395 -0.1258 0.7110 0.1052
-0.750 0.5046 0.00897 0.00411 -0.1226 0.7090 0.7900
-0.500 0.5289 0.00917 0.00431 -0.1215 0.7073 0.8324
-0.250 0.5535 0.00941 0.00451 -0.1206 0.7056 0.8528
0.000 0.5757 0.00955 0.00467 -0.1193 0.7034 0.8659
0.250 0.5976 0.00969 0.00482 -0.1179 0.7011 0.8783
0.500 0.6210 0.00982 0.00493 -0.1170 0.6989 0.8892
0.750 0.6436 0.00989 0.00499 -0.1158 0.6970 0.8963
1.000 0.6694 0.00997 0.00503 -0.1156 0.6953 0.9038
1.250 0.6961 0.01000 0.00502 -0.1156 0.6937 0.9069
1.500 0.7250 0.01006 0.00503 -0.1161 0.6921 0.9090
1.750 0.7513 0.01014 0.00510 -0.1163 0.6901 0.9116
2.000 0.7765 0.01020 0.00518 -0.1162 0.6878 0.9148
2.250 0.8034 0.01027 0.00525 -0.1165 0.6857 0.9179
2.500 0.8291 0.01030 0.00529 -0.1164 0.6838 0.9203
2.750 0.8559 0.01035 0.00534 -0.1166 0.6820 0.9229
3.000 0.8838 0.01039 0.00538 -0.1170 0.6803 0.9257
3.250 0.9129 0.01046 0.00543 -0.1177 0.6787 0.9286
3.500 0.9420 0.01059 0.00554 -0.1184 0.6770 0.9318
3.750 0.9633 0.01066 0.00569 -0.1175 0.6749 0.9354
4.000 0.9862 0.01074 0.00585 -0.1169 0.6726 0.9392
4.250 1.0111 0.01082 0.00596 -0.1168 0.6703 0.9432
4.500 1.0375 0.01088 0.00605 -0.1170 0.6683 0.9470
4.750 1.0640 0.01092 0.00611 -0.1171 0.6665 0.9508
5.000 1.0927 0.01097 0.00618 -0.1176 0.6646 0.9548
5.250 1.1154 0.01098 0.00625 -0.1170 0.6606 0.9603
5.500 1.1390 0.01068 0.00598 -0.1163 0.6530 0.9656
5.750 1.1627 0.01049 0.00580 -0.1156 0.6451 0.9722
6.000 1.1923 0.01044 0.00580 -0.1165 0.6395 0.9769
6.250 1.2263 0.01045 0.00582 -0.1183 0.6358 0.9808
6.500 1.2556 0.01054 0.00603 -0.1193 0.6308 0.9870
6.750 1.2880 0.01057 0.00612 -0.1210 0.6250 0.9930
7.000 1.3167 0.01061 0.00620 -0.1218 0.6191 1.0000
7.250 1.3289 0.01067 0.00634 -0.1193 0.6128 1.0000
7.500 1.3508 0.01072 0.00641 -0.1187 0.6062 1.0000
7.750 1.3610 0.01085 0.00664 -0.1157 0.5980 1.0000
8.000 1.3749 0.01101 0.00683 -0.1135 0.5891 1.0000
8.250 1.3880 0.01122 0.00705 -0.1112 0.5789 1.0000
8.500 1.3958 0.01156 0.00742 -0.1079 0.5653 1.0000
8.750 1.4022 0.01201 0.00792 -0.1046 0.5490 1.0000
9.000 1.4055 0.01265 0.00854 -0.1008 0.5271 1.0000
9.250 1.3990 0.01372 0.00949 -0.0957 0.4938 1.0000
9.500 1.3818 0.01549 0.01103 -0.0893 0.4493 1.0000
9.750 1.3611 0.01775 0.01306 -0.0830 0.4059 1.0000
10.000 1.3396 0.02035 0.01541 -0.0771 0.3621 1.0000
10.250 1.3253 0.02279 0.01766 -0.0726 0.3249 1.0000
10.500 1.3130 0.02527 0.01996 -0.0686 0.2874 1.0000
10.750 1.3052 0.02761 0.02213 -0.0653 0.2547 1.0000
11.000 1.2994 0.02990 0.02426 -0.0625 0.2234 1.0000
11.250 1.2963 0.03210 0.02631 -0.0601 0.1943 1.0000
11.750 1.2984 0.03605 0.03006 -0.0563 0.1476 1.0000
12.000 1.2976 0.03825 0.03210 -0.0544 0.1214 1.0000
12.250 1.3042 0.03991 0.03373 -0.0532 0.1080 1.0000
12.500 1.3094 0.04172 0.03549 -0.0520 0.0937 1.0000
12.750 1.3130 0.04368 0.03735 -0.0507 0.0763 1.0000
13.000 1.3200 0.04543 0.03910 -0.0497 0.0674 1.0000
13.250 1.3245 0.04739 0.04099 -0.0486 0.0528 1.0000
13.500 1.3288 0.04942 0.04295 -0.0476 0.0410 1.0000
13.750 1.3345 0.05136 0.04486 -0.0468 0.0327 1.0000
14.000 1.3395 0.05339 0.04688 -0.0459 0.0254 1.0000
14.250 1.3448 0.05546 0.04895 -0.0452 0.0204 1.0000
14.500 1.3526 0.05731 0.05087 -0.0446 0.0178 1.0000
14.750 1.3580 0.05942 0.05299 -0.0440 0.0139 1.0000
15.000 1.3614 0.06179 0.05538 -0.0434 0.0105 1.0000
15.250 1.3605 0.06467 0.05825 -0.0426 0.0053 1.0000
15.500 1.3563 0.06808 0.06175 -0.0417 0.0028 1.0000
15.750 1.3607 0.07054 0.06432 -0.0414 0.0028 1.0000
16.000 1.3578 0.07392 0.06781 -0.0410 0.0020 1.0000
16.250 1.3573 0.07708 0.07110 -0.0407 0.0019 1.0000
16.500 1.3571 0.08027 0.07444 -0.0406 0.0017 1.0000
16.750 1.3572 0.08350 0.07778 -0.0406 0.0017 1.0000
17.000 1.3566 0.08689 0.08130 -0.0408 0.0016 1.0000
17.250 1.3534 0.09065 0.08520 -0.0410 0.0014 1.0000
17.500 1.3515 0.09434 0.08901 -0.0415 0.0015 1.0000
17.750 1.3477 0.09836 0.09316 -0.0421 0.0014 1.0000
18.000 1.3426 0.10265 0.09758 -0.0429 0.0013 1.0000
18.250 1.3381 0.10693 0.10199 -0.0438 0.0013 1.0000
18.500 1.3310 0.11170 0.10689 -0.0450 0.0012 1.0000
18.750 1.3270 0.11604 0.11136 -0.0463 0.0012 1.0000
19.000 1.3212 0.12073 0.11618 -0.0479 0.0012 1.0000
19.250 1.3147 0.12564 0.12123 -0.0496 0.0012 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S832 Airfoil (s832-nr)