NREL's S802 Airfoil (s802-nr) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S802 Airfoil (s802-nr) Reynolds number: 500,000 Max Cl/Cd: 113.64 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s802-nr-500000-n5.txt Download as CSV file: xf-s802-nr-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S802 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3110 0.08199 0.07960 -0.0735 0.9804 0.0101
-9.500 -0.3160 0.07562 0.07326 -0.0778 0.9707 0.0102
-9.250 -0.4282 0.03295 0.02977 -0.1187 0.9264 0.0102
-9.000 -0.4041 0.02667 0.02279 -0.1259 0.9168 0.0109
-8.750 -0.3741 0.02324 0.01877 -0.1300 0.9060 0.0116
-8.500 -0.3439 0.02144 0.01655 -0.1322 0.8937 0.0119
-8.250 -0.3185 0.01943 0.01421 -0.1335 0.8806 0.0123
-8.000 -0.2924 0.01837 0.01297 -0.1341 0.8683 0.0126
-7.750 -0.2668 0.01754 0.01197 -0.1344 0.8564 0.0129
-7.250 -0.2169 0.01606 0.01015 -0.1344 0.8347 0.0136
-7.000 -0.1917 0.01546 0.00938 -0.1342 0.8256 0.0141
-6.750 -0.1662 0.01491 0.00869 -0.1341 0.8170 0.0147
-6.500 -0.1405 0.01437 0.00800 -0.1340 0.8094 0.0152
-6.250 -0.1148 0.01384 0.00734 -0.1338 0.8020 0.0155
-6.000 -0.0887 0.01338 0.00678 -0.1337 0.7955 0.0158
-5.750 -0.0635 0.01260 0.00591 -0.1336 0.7887 0.0163
-5.500 -0.0373 0.01214 0.00537 -0.1336 0.7828 0.0168
-5.250 -0.0107 0.01175 0.00494 -0.1336 0.7766 0.0173
-5.000 0.0162 0.01142 0.00454 -0.1336 0.7709 0.0179
-4.750 0.0435 0.01115 0.00422 -0.1337 0.7660 0.0188
-4.500 0.0709 0.01087 0.00388 -0.1337 0.7610 0.0197
-4.250 0.0985 0.01063 0.00357 -0.1338 0.7564 0.0202
-4.000 0.1262 0.01035 0.00323 -0.1340 0.7524 0.0212
-3.750 0.1541 0.01011 0.00297 -0.1341 0.7480 0.0225
-3.500 0.1820 0.00994 0.00275 -0.1342 0.7436 0.0241
-3.250 0.2101 0.00979 0.00256 -0.1344 0.7397 0.0266
-3.000 0.2383 0.00959 0.00240 -0.1346 0.7360 0.0366
-2.750 0.2665 0.00934 0.00225 -0.1349 0.7324 0.0656
-2.500 0.2948 0.00912 0.00214 -0.1352 0.7289 0.1017
-2.250 0.3232 0.00889 0.00205 -0.1356 0.7257 0.1508
-2.000 0.3518 0.00867 0.00197 -0.1360 0.7227 0.2077
-1.750 0.3806 0.00829 0.00192 -0.1367 0.7193 0.3072
-1.500 0.4097 0.00793 0.00188 -0.1374 0.7160 0.4141
-1.250 0.4388 0.00750 0.00187 -0.1381 0.7130 0.5516
-1.000 0.4664 0.00735 0.00204 -0.1381 0.7103 0.6654
-0.500 0.5230 0.00746 0.00214 -0.1381 0.7046 0.7094
-0.250 0.5512 0.00751 0.00220 -0.1382 0.7016 0.7225
0.000 0.5792 0.00757 0.00225 -0.1382 0.6985 0.7323
0.250 0.6074 0.00764 0.00230 -0.1382 0.6954 0.7414
0.500 0.6354 0.00772 0.00236 -0.1382 0.6922 0.7495
0.750 0.6631 0.00776 0.00242 -0.1382 0.6877 0.7569
1.000 0.6905 0.00781 0.00248 -0.1380 0.6828 0.7637
1.250 0.7183 0.00790 0.00252 -0.1380 0.6778 0.7704
1.500 0.7450 0.00794 0.00260 -0.1377 0.6715 0.7760
1.750 0.7719 0.00801 0.00265 -0.1375 0.6643 0.7823
2.000 0.7989 0.00806 0.00271 -0.1373 0.6569 0.7868
2.250 0.8253 0.00812 0.00276 -0.1370 0.6489 0.7901
2.500 0.8521 0.00817 0.00284 -0.1368 0.6407 0.7931
2.750 0.8785 0.00826 0.00289 -0.1365 0.6310 0.7960
3.000 0.9050 0.00832 0.00296 -0.1362 0.6199 0.7989
3.250 0.9311 0.00840 0.00304 -0.1359 0.6077 0.8013
3.500 0.9565 0.00849 0.00313 -0.1354 0.5931 0.8037
3.750 0.9807 0.00863 0.00322 -0.1346 0.5723 0.8063
4.000 1.0024 0.00887 0.00334 -0.1334 0.5341 0.8089
4.250 1.0197 0.00939 0.00359 -0.1314 0.4785 0.8120
4.500 1.0356 0.01007 0.00398 -0.1293 0.4192 0.8152
4.750 1.0510 0.01078 0.00442 -0.1271 0.3617 0.8179
5.000 1.0677 0.01143 0.00484 -0.1252 0.3122 0.8207
5.250 1.0851 0.01203 0.00525 -0.1235 0.2712 0.8237
5.500 1.1031 0.01260 0.00566 -0.1219 0.2350 0.8273
5.750 1.1208 0.01317 0.00608 -0.1202 0.2025 0.8309
6.000 1.1386 0.01367 0.00648 -0.1186 0.1768 0.8340
6.250 1.1558 0.01418 0.00690 -0.1168 0.1539 0.8376
6.500 1.1724 0.01466 0.00731 -0.1149 0.1344 0.8417
6.750 1.1888 0.01513 0.00773 -0.1130 0.1187 0.8458
7.000 1.2049 0.01560 0.00816 -0.1110 0.1055 0.8502
7.250 1.2213 0.01607 0.00861 -0.1091 0.0943 0.8556
7.500 1.2375 0.01655 0.00909 -0.1073 0.0852 0.8614
7.750 1.2534 0.01703 0.00958 -0.1053 0.0772 0.8678
8.250 1.2841 0.01804 0.01063 -0.1015 0.0640 0.8839
8.500 1.2992 0.01851 0.01116 -0.0995 0.0593 0.8955
8.750 1.3120 0.01902 0.01173 -0.0971 0.0549 0.9160
9.000 1.3254 0.01952 0.01231 -0.0950 0.0513 1.0000
9.250 1.3415 0.02015 0.01297 -0.0935 0.0479 1.0000
9.500 1.3562 0.02090 0.01372 -0.0919 0.0443 1.0000
9.750 1.3708 0.02165 0.01450 -0.0903 0.0417 1.0000
10.000 1.3859 0.02238 0.01530 -0.0889 0.0396 1.0000
10.250 1.3999 0.02320 0.01615 -0.0873 0.0372 1.0000
10.500 1.4122 0.02417 0.01713 -0.0857 0.0348 1.0000
10.750 1.4257 0.02507 0.01809 -0.0842 0.0333 1.0000
11.000 1.4393 0.02598 0.01907 -0.0828 0.0319 1.0000
11.250 1.4521 0.02696 0.02011 -0.0814 0.0302 1.0000
11.500 1.4637 0.02808 0.02126 -0.0799 0.0284 1.0000
11.750 1.4744 0.02928 0.02251 -0.0785 0.0268 1.0000
12.000 1.4873 0.03033 0.02363 -0.0773 0.0254 1.0000
12.250 1.4992 0.03148 0.02483 -0.0761 0.0238 1.0000
12.500 1.5096 0.03278 0.02618 -0.0748 0.0224 1.0000
12.750 1.5185 0.03426 0.02770 -0.0735 0.0211 1.0000
13.000 1.5291 0.03559 0.02913 -0.0724 0.0203 1.0000
13.250 1.5391 0.03701 0.03065 -0.0714 0.0194 1.0000
13.500 1.5482 0.03854 0.03225 -0.0703 0.0184 1.0000
13.750 1.5564 0.04020 0.03397 -0.0693 0.0174 1.0000
14.000 1.5628 0.04206 0.03588 -0.0683 0.0165 1.0000
14.250 1.5712 0.04376 0.03769 -0.0674 0.0157 1.0000
14.500 1.5787 0.04559 0.03962 -0.0666 0.0148 1.0000
14.750 1.5851 0.04758 0.04168 -0.0659 0.0139 1.0000
15.000 1.5900 0.04976 0.04394 -0.0652 0.0130 1.0000
15.250 1.5948 0.05201 0.04629 -0.0646 0.0123 1.0000
15.500 1.5996 0.05434 0.04873 -0.0641 0.0116 1.0000
15.750 1.6030 0.05686 0.05135 -0.0638 0.0109 1.0000
16.000 1.6052 0.05961 0.05418 -0.0635 0.0103 1.0000
16.250 1.6058 0.06262 0.05729 -0.0633 0.0097 1.0000
16.500 1.6069 0.06567 0.06046 -0.0633 0.0092 1.0000
16.750 1.6068 0.06896 0.06388 -0.0635 0.0087 1.0000
17.000 1.6053 0.07252 0.06756 -0.0638 0.0083 1.0000
17.250 1.6023 0.07642 0.07157 -0.0644 0.0079 1.0000
17.500 1.5972 0.08075 0.07603 -0.0653 0.0075 1.0000
17.750 1.5911 0.08539 0.08081 -0.0665 0.0073 1.0000
18.000 1.5852 0.09015 0.08572 -0.0679 0.0070 1.0000
18.250 1.5777 0.09532 0.09105 -0.0697 0.0068 1.0000
18.500 1.5689 0.10085 0.09672 -0.0718 0.0066 1.0000
18.750 1.5587 0.10683 0.10286 -0.0744 0.0064 1.0000
19.000 1.5471 0.11322 0.10941 -0.0774 0.0062 1.0000
19.250 1.5341 0.12006 0.11640 -0.0808 0.0061 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S802 Airfoil (s802-nr)