NREL's S810 Airfoil (s810-nr) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S810 Airfoil (s810-nr) Reynolds number: 500,000 Max Cl/Cd: 83.24 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s810-nr-500000.txt Download as CSV file: xf-s810-nr-500000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S810 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.4508 0.08900 0.08556 -0.0774 0.7767 0.0121
-12.000 -0.4612 0.08503 0.08151 -0.0785 0.7751 0.0121
-11.750 -0.4762 0.08113 0.07752 -0.0792 0.7736 0.0121
-11.500 -0.4921 0.07749 0.07377 -0.0793 0.7721 0.0121
-11.250 -0.5119 0.07422 0.07038 -0.0785 0.7707 0.0121
-11.000 -0.5217 0.07126 0.06735 -0.0780 0.7691 0.0121
-10.750 -0.5382 0.06852 0.06449 -0.0764 0.7675 0.0121
-10.500 -0.5524 0.06602 0.06187 -0.0744 0.7659 0.0122
-10.250 -0.5613 0.06368 0.05943 -0.0725 0.7643 0.0122
-10.000 -0.5795 0.06163 0.05723 -0.0686 0.7626 0.0122
-9.750 -0.5895 0.05970 0.05515 -0.0652 0.7612 0.0122
-9.500 -0.5944 0.05746 0.05274 -0.0624 0.7599 0.0122
-9.250 -0.6177 0.05009 0.04512 -0.0587 0.7587 0.0127
-9.000 -0.6140 0.04699 0.04193 -0.0574 0.7577 0.0131
-8.750 -0.6066 0.04473 0.03955 -0.0559 0.7566 0.0133
-8.500 -0.5981 0.04262 0.03735 -0.0546 0.7555 0.0137
-8.250 -0.5869 0.04062 0.03521 -0.0533 0.7542 0.0141
-8.000 -0.5742 0.03871 0.03316 -0.0520 0.7529 0.0150
-7.750 -0.5581 0.03692 0.03120 -0.0507 0.7517 0.0161
-7.500 -0.5378 0.03581 0.02988 -0.0494 0.7506 0.0177
-7.250 -0.5168 0.03673 0.03059 -0.0478 0.7495 0.0188
-7.000 -0.5003 0.03589 0.02951 -0.0462 0.7485 0.0190
-6.750 -0.4803 0.02944 0.02279 -0.0459 0.7478 0.0205
-6.500 -0.4532 0.02750 0.02082 -0.0464 0.7470 0.0213
-6.250 -0.4267 0.02599 0.01922 -0.0467 0.7462 0.0224
-6.000 -0.3990 0.02470 0.01784 -0.0468 0.7455 0.0238
-5.750 -0.3705 0.02372 0.01675 -0.0470 0.7447 0.0258
-5.500 -0.3532 0.02438 0.01726 -0.0454 0.7438 0.0282
-5.250 -0.2341 0.01998 0.01284 -0.0558 0.7441 0.0178
-5.000 -0.2084 0.01943 0.01230 -0.0550 0.7436 0.0178
-4.750 -0.1920 0.01879 0.01167 -0.0533 0.7428 0.0178
-4.250 -0.1688 0.01734 0.01020 -0.0496 0.7408 0.0198
-4.000 -0.1672 0.01640 0.00923 -0.0465 0.7395 0.0201
-3.750 -0.1672 0.01569 0.00846 -0.0430 0.7384 0.0202
-3.500 -0.1596 0.01495 0.00767 -0.0405 0.7374 0.0212
-3.250 -0.1443 0.01442 0.00708 -0.0391 0.7365 0.0227
-3.000 -0.1251 0.01401 0.00661 -0.0382 0.7357 0.0261
-2.750 -0.1036 0.01367 0.00619 -0.0377 0.7349 0.0312
-2.500 -0.0959 0.01221 0.00534 -0.0357 0.7339 0.2514
-2.250 -0.0902 0.01046 0.00490 -0.0335 0.7328 0.6101
-2.000 -0.0613 0.01118 0.00562 -0.0336 0.7321 0.6329
-1.750 -0.0322 0.01170 0.00618 -0.0338 0.7315 0.6423
-1.500 -0.0034 0.01206 0.00645 -0.0343 0.7308 0.6491
-1.250 0.0253 0.01220 0.00662 -0.0347 0.7302 0.6516
-1.000 0.0538 0.01239 0.00683 -0.0352 0.7294 0.6542
-0.750 0.0824 0.01271 0.00719 -0.0355 0.7284 0.6593
-0.500 0.1107 0.01282 0.00727 -0.0361 0.7276 0.6623
-0.250 0.1390 0.01283 0.00726 -0.0368 0.7265 0.6635
0.000 0.1671 0.01282 0.00727 -0.0373 0.7253 0.6646
0.250 0.1953 0.01286 0.00733 -0.0379 0.7243 0.6656
0.500 0.2236 0.01290 0.00739 -0.0385 0.7232 0.6666
0.750 0.2521 0.01294 0.00744 -0.0391 0.7220 0.6676
1.000 0.2806 0.01298 0.00750 -0.0397 0.7209 0.6687
1.250 0.3093 0.01302 0.00754 -0.0403 0.7197 0.6698
1.500 0.3381 0.01305 0.00758 -0.0410 0.7186 0.6710
1.750 0.3670 0.01310 0.00763 -0.0416 0.7174 0.6723
2.000 0.3957 0.01320 0.00773 -0.0422 0.7162 0.6738
2.250 0.4229 0.01328 0.00787 -0.0427 0.7143 0.6753
2.500 0.4502 0.01333 0.00796 -0.0432 0.7118 0.6766
2.750 0.4785 0.01330 0.00796 -0.0437 0.7091 0.6776
3.000 0.5078 0.01312 0.00781 -0.0443 0.7062 0.6787
3.250 0.5387 0.01290 0.00761 -0.0450 0.7033 0.6797
3.500 0.5664 0.01276 0.00753 -0.0452 0.6984 0.6807
3.750 0.5958 0.01245 0.00725 -0.0456 0.6926 0.6818
4.000 0.6275 0.01214 0.00691 -0.0463 0.6883 0.6831
4.250 0.6542 0.01196 0.00683 -0.0464 0.6818 0.6845
4.500 0.6847 0.01157 0.00644 -0.0469 0.6753 0.6859
4.750 0.7120 0.01133 0.00627 -0.0470 0.6678 0.6873
5.000 0.7419 0.01095 0.00587 -0.0474 0.6596 0.6887
5.250 0.7675 0.01076 0.00581 -0.0473 0.6502 0.6900
5.500 0.7945 0.01058 0.00571 -0.0473 0.6405 0.6912
5.750 0.8202 0.01035 0.00558 -0.0471 0.6244 0.6924
6.000 0.8441 0.01014 0.00540 -0.0465 0.5895 0.6936
6.250 0.8461 0.01080 0.00550 -0.0418 0.4828 0.6949
6.500 0.8377 0.01203 0.00627 -0.0358 0.3887 0.6964
6.750 0.8306 0.01297 0.00691 -0.0301 0.3261 0.6979
7.000 0.8169 0.01408 0.00771 -0.0235 0.2647 0.6995
7.250 0.8110 0.01530 0.00866 -0.0188 0.2116 0.7010
7.500 0.8114 0.01644 0.00960 -0.0155 0.1683 0.7025
7.750 0.8153 0.01755 0.01053 -0.0129 0.1321 0.7040
8.000 0.8211 0.01865 0.01146 -0.0106 0.1008 0.7053
8.500 0.8393 0.02055 0.01320 -0.0071 0.0597 0.7082
8.750 0.8511 0.02141 0.01403 -0.0058 0.0487 0.7097
9.000 0.8630 0.02230 0.01489 -0.0045 0.0411 0.7114
9.250 0.8790 0.02296 0.01563 -0.0037 0.0378 0.7130
9.500 0.8918 0.02383 0.01648 -0.0026 0.0317 0.7146
9.750 0.9081 0.02450 0.01723 -0.0019 0.0273 0.7163
10.000 0.9199 0.02549 0.01821 -0.0007 0.0202 0.7179
10.250 0.9289 0.02672 0.01937 0.0007 0.0120 0.7193
10.500 0.9408 0.02772 0.02043 0.0018 0.0110 0.7209
10.750 0.9486 0.02905 0.02185 0.0033 0.0098 0.7224
11.000 0.9596 0.03017 0.02308 0.0044 0.0091 0.7241
11.250 0.9696 0.03142 0.02442 0.0056 0.0086 0.7259
11.500 0.9793 0.03270 0.02580 0.0067 0.0085 0.7279
11.750 0.9883 0.03408 0.02726 0.0078 0.0079 0.7299
12.000 0.9975 0.03549 0.02874 0.0087 0.0076 0.7318
12.250 1.0051 0.03707 0.03042 0.0097 0.0072 0.7336
12.500 1.0092 0.03900 0.03247 0.0110 0.0069 0.7353
12.750 1.0124 0.04119 0.03482 0.0123 0.0067 0.7369
13.000 1.0210 0.04290 0.03667 0.0131 0.0066 0.7388
13.250 1.0290 0.04475 0.03867 0.0139 0.0065 0.7408
13.500 1.0371 0.04655 0.04061 0.0145 0.0064 0.7429
13.750 1.0427 0.04887 0.04312 0.0153 0.0064 0.7450
14.000 1.0475 0.05119 0.04561 0.0158 0.0063 0.7470
14.250 1.0505 0.05370 0.04832 0.0163 0.0061 0.7490
14.500 1.0508 0.05661 0.05144 0.0168 0.0061 0.7510
14.750 1.0480 0.05999 0.05505 0.0171 0.0061 0.7531
15.000 1.0425 0.06375 0.05903 0.0172 0.0061 0.7553
15.250 1.0354 0.06780 0.06329 0.0169 0.0060 0.7577
15.500 1.0260 0.07233 0.06803 0.0162 0.0060 0.7599
15.750 1.0117 0.07781 0.07374 0.0150 0.0060 0.7618
16.000 0.9966 0.08364 0.07980 0.0133 0.0060 0.7637
16.250 0.9804 0.09004 0.08642 0.0108 0.0061 0.7654
16.500 0.9608 0.09748 0.09409 0.0074 0.0060 0.7671
16.750 0.9434 0.10494 0.10172 0.0035 0.0062 0.7689
17.000 0.9221 0.11379 0.11077 -0.0014 0.0062 0.7705
17.250 0.8997 0.12358 0.12072 -0.0073 0.0063 0.7720
17.500 0.8775 0.13410 0.13140 -0.0138 0.0064 0.7732
17.750 0.8541 0.14584 0.14328 -0.0210 0.0064 0.7741
18.000 0.8181 0.16199 0.15954 -0.0298 0.0067 0.7736
|
Polar data table (+)
Polar graphs
<< Back to NREL's S810 Airfoil (s810-nr)