NREL's S815 Airfoil (s815-nr) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S815 Airfoil (s815-nr) Reynolds number: 200,000 Max Cl/Cd: 54.38 at α=10° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s815-nr-200000.txt Download as CSV file: xf-s815-nr-200000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S815 Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.0147 0.11256 0.10719 -0.0479 0.9707 0.3131
-7.250 0.0180 0.11008 0.10471 -0.0519 0.9672 0.3147
-7.000 -0.0417 0.11124 0.10577 -0.0561 0.9566 0.3248
-6.750 0.0021 0.10689 0.10143 -0.0595 0.9529 0.3252
-6.500 0.0474 0.10300 0.09755 -0.0636 0.9502 0.3256
-6.250 0.0782 0.10012 0.09471 -0.0645 0.9393 0.3260
-6.000 0.1177 0.09712 0.09174 -0.0681 0.9346 0.3267
-5.750 0.1420 0.09495 0.08960 -0.0684 0.9194 0.3274
-5.500 0.1636 0.09302 0.08769 -0.0685 0.9007 0.3284
-5.250 0.1938 0.09069 0.08537 -0.0708 0.8857 0.3297
-5.000 0.2314 0.08800 0.08264 -0.0754 0.8704 0.3315
-4.750 0.2010 0.08756 0.08202 -0.0837 0.8543 0.3414
-4.500 0.2899 0.08194 0.07628 -0.0968 0.8370 0.3419
-4.250 0.3785 0.07710 0.07119 -0.1107 0.8047 0.3424
-4.000 0.4278 0.07456 0.06835 -0.1164 0.7660 0.3430
-3.750 0.4562 0.07314 0.06669 -0.1176 0.7357 0.3437
-3.500 0.4773 0.07210 0.06549 -0.1174 0.7122 0.3446
-3.250 0.4959 0.07118 0.06444 -0.1169 0.6937 0.3457
-3.000 0.5129 0.07034 0.06351 -0.1164 0.6776 0.3471
-2.750 0.5280 0.06953 0.06265 -0.1157 0.6640 0.3488
-2.500 0.4568 0.07147 0.06449 -0.1125 0.6576 0.3583
-2.250 0.4835 0.06945 0.06245 -0.1124 0.6463 0.3587
-2.000 0.5104 0.06782 0.06073 -0.1125 0.6367 0.3591
-1.750 0.5350 0.06640 0.05933 -0.1123 0.6272 0.3595
-1.500 0.5593 0.06516 0.05803 -0.1122 0.6189 0.3601
-1.250 0.5823 0.06407 0.05693 -0.1120 0.6115 0.3608
-1.000 0.6026 0.06310 0.05598 -0.1115 0.6041 0.3617
-0.750 0.6236 0.06223 0.05506 -0.1114 0.5982 0.3629
-0.500 0.6420 0.06144 0.05429 -0.1110 0.5927 0.3643
-0.250 0.6574 0.06069 0.05358 -0.1104 0.5872 0.3661
0.000 0.6707 0.05999 0.05289 -0.1097 0.5825 0.3688
0.250 0.6123 0.05972 0.05261 -0.1061 0.5801 0.3763
0.500 0.6444 0.05845 0.05130 -0.1068 0.5757 0.3768
0.750 0.6689 0.05741 0.05035 -0.1063 0.5712 0.3775
1.000 0.6923 0.05649 0.04948 -0.1058 0.5668 0.3782
1.250 0.7138 0.05565 0.04866 -0.1054 0.5624 0.3791
1.500 0.7359 0.05488 0.04789 -0.1053 0.5588 0.3802
1.750 0.7565 0.05424 0.04726 -0.1051 0.5556 0.3816
2.000 0.7699 0.05364 0.04676 -0.1038 0.5524 0.3835
2.250 0.7805 0.05308 0.04629 -0.1025 0.5491 0.3863
2.500 0.5613 0.02800 0.02044 -0.1150 0.5512 0.3619
2.750 0.5866 0.02779 0.02028 -0.1152 0.5477 0.3630
3.000 0.6141 0.02748 0.02000 -0.1159 0.5445 0.3643
3.250 0.6446 0.02699 0.01949 -0.1176 0.5415 0.3660
3.500 0.6786 0.02654 0.01895 -0.1200 0.5385 0.3685
3.750 0.7112 0.02591 0.01824 -0.1226 0.5353 0.3716
4.000 0.7423 0.02528 0.01756 -0.1248 0.5321 0.3742
4.250 0.7727 0.02489 0.01714 -0.1263 0.5288 0.3760
4.500 0.7998 0.02475 0.01710 -0.1265 0.5254 0.3776
4.750 0.8290 0.02478 0.01719 -0.1268 0.5221 0.3791
5.000 0.8613 0.02490 0.01731 -0.1278 0.5190 0.3807
5.250 0.8889 0.02508 0.01757 -0.1280 0.5157 0.3826
5.500 0.9119 0.02512 0.01774 -0.1274 0.5119 0.3846
5.750 0.9384 0.02513 0.01781 -0.1275 0.5079 0.3870
6.000 0.9697 0.02503 0.01770 -0.1285 0.5038 0.3898
6.250 1.0074 0.02496 0.01749 -0.1306 0.4996 0.3927
6.500 1.0327 0.02495 0.01758 -0.1304 0.4950 0.3950
6.750 1.0519 0.02499 0.01781 -0.1289 0.4895 0.3970
7.000 1.0782 0.02497 0.01785 -0.1284 0.4839 0.3995
7.250 1.1149 0.02504 0.01783 -0.1298 0.4784 0.4028
7.500 1.1289 0.02504 0.01802 -0.1274 0.4723 0.4057
7.750 1.1547 0.02488 0.01786 -0.1271 0.4656 0.4092
8.000 1.1929 0.02471 0.01753 -0.1289 0.4595 0.4127
8.250 1.2037 0.02472 0.01780 -0.1259 0.4533 0.4149
8.500 1.2243 0.02465 0.01783 -0.1244 0.4465 0.4176
8.750 1.2560 0.02459 0.01771 -0.1248 0.4400 0.4212
9.000 1.2646 0.02463 0.01794 -0.1215 0.4328 0.4244
9.250 1.2861 0.02447 0.01775 -0.1203 0.4252 0.4284
9.500 1.3029 0.02441 0.01774 -0.1183 0.4176 0.4317
9.750 1.3080 0.02435 0.01784 -0.1143 0.4096 0.4342
10.000 1.3236 0.02434 0.01784 -0.1119 0.4014 0.4374
10.250 1.3259 0.02462 0.01830 -0.1078 0.3921 0.4406
10.500 1.3382 0.02478 0.01844 -0.1053 0.3824 0.4449
10.750 1.3433 0.02518 0.01893 -0.1019 0.3705 0.4486
11.000 1.3490 0.02556 0.01944 -0.0988 0.3573 0.4518
11.250 1.3522 0.02614 0.02012 -0.0954 0.3415 0.4548
11.500 1.3526 0.02698 0.02099 -0.0919 0.3212 0.4579
12.000 1.3383 0.03005 0.02384 -0.0847 0.2720 0.4640
12.250 1.3257 0.03239 0.02603 -0.0815 0.2494 0.4667
12.750 1.2993 0.03808 0.03158 -0.0768 0.2100 0.4712
13.000 1.2861 0.04138 0.03486 -0.0751 0.1929 0.4733
13.250 1.2729 0.04497 0.03844 -0.0740 0.1777 0.4755
13.750 1.2503 0.05259 0.04601 -0.0731 0.1509 0.4808
14.000 1.2419 0.05642 0.04983 -0.0731 0.1393 0.4839
14.250 1.2339 0.06038 0.05374 -0.0734 0.1292 0.4869
14.500 1.2245 0.06455 0.05789 -0.0738 0.1205 0.4895
14.750 1.2215 0.06810 0.06151 -0.0742 0.1117 0.4924
15.000 1.2166 0.07193 0.06536 -0.0747 0.1044 0.4956
15.250 1.2127 0.07575 0.06917 -0.0754 0.0976 0.4992
15.500 1.2123 0.07926 0.07270 -0.0760 0.0911 0.5033
15.750 1.2100 0.08308 0.07643 -0.0769 0.0858 0.5071
16.000 1.2121 0.08634 0.07981 -0.0777 0.0803 0.5108
16.250 1.2114 0.08990 0.08336 -0.0785 0.0758 0.5144
16.500 1.2149 0.09306 0.08659 -0.0792 0.0713 0.5189
16.750 1.2169 0.09646 0.08996 -0.0802 0.0676 0.5239
17.000 1.2215 0.09950 0.09304 -0.0811 0.0639 0.5285
17.250 1.2253 0.10262 0.09624 -0.0820 0.0608 0.5330
|
Polar data table (+)
Polar graphs
<< Back to NREL's S815 Airfoil (s815-nr)