NREL's S805A Airfoil (s805a-nr) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S805A Airfoil (s805a-nr) Reynolds number: 100,000 Max Cl/Cd: 50.61 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s805a-nr-100000-n5.txt Download as CSV file: xf-s805a-nr-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S805A Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4682 0.07892 0.07392 -0.0598 1.0000 0.0282
-10.000 -0.4816 0.07441 0.06941 -0.0613 1.0000 0.0280
-9.750 -0.4964 0.07062 0.06561 -0.0618 1.0000 0.0277
-9.500 -0.5153 0.06725 0.06222 -0.0609 1.0000 0.0276
-9.250 -0.5360 0.06459 0.05955 -0.0584 1.0000 0.0274
-9.000 -0.5576 0.06246 0.05740 -0.0545 1.0000 0.0273
-8.750 -0.5758 0.05997 0.05485 -0.0507 1.0000 0.0271
-8.500 -0.5923 0.05763 0.05244 -0.0466 1.0000 0.0270
-8.250 -0.5877 0.05305 0.04753 -0.0477 0.9905 0.0268
-8.000 -0.5739 0.04828 0.04233 -0.0496 0.9782 0.0266
-7.750 -0.5579 0.04395 0.03752 -0.0506 0.9654 0.0267
-7.500 -0.5388 0.04017 0.03320 -0.0513 0.9527 0.0272
-7.250 -0.5173 0.03713 0.02946 -0.0514 0.9394 0.0283
-7.000 -0.4937 0.03418 0.02620 -0.0520 0.9269 0.0293
-6.750 -0.4667 0.03174 0.02343 -0.0526 0.9149 0.0297
-6.500 -0.4385 0.02964 0.02103 -0.0531 0.9024 0.0303
-6.250 -0.4082 0.02777 0.01889 -0.0538 0.8909 0.0310
-6.000 -0.3760 0.02608 0.01695 -0.0547 0.8804 0.0320
-5.750 -0.3428 0.02472 0.01537 -0.0557 0.8705 0.0343
-5.500 -0.3123 0.02363 0.01404 -0.0561 0.8588 0.0365
-5.250 -0.2840 0.02227 0.01264 -0.0563 0.8479 0.0381
-5.000 -0.2575 0.02127 0.01159 -0.0562 0.8378 0.0399
-4.750 -0.2326 0.02045 0.01065 -0.0558 0.8276 0.0425
-4.500 -0.2090 0.01980 0.00985 -0.0551 0.8172 0.0460
-4.250 -0.1855 0.01911 0.00910 -0.0545 0.8083 0.0524
-4.000 -0.1625 0.01850 0.00842 -0.0537 0.7993 0.0621
-3.750 -0.1405 0.01777 0.00776 -0.0527 0.7913 0.0893
-3.500 -0.1247 0.01650 0.00724 -0.0513 0.7838 0.2340
-3.250 -0.1201 0.01489 0.00718 -0.0473 0.7765 0.5466
-3.000 -0.1017 0.01548 0.00822 -0.0432 0.7698 0.7229
-2.750 -0.0794 0.01638 0.00905 -0.0399 0.7641 0.7752
-2.500 -0.0579 0.01732 0.00989 -0.0363 0.7576 0.8126
-2.250 -0.0265 0.01789 0.01029 -0.0353 0.7528 0.8309
-2.000 0.0010 0.01792 0.01015 -0.0352 0.7474 0.8373
-1.750 0.0234 0.01784 0.00990 -0.0344 0.7416 0.8441
-1.500 0.0536 0.01783 0.00972 -0.0348 0.7371 0.8482
-1.250 0.0766 0.01782 0.00960 -0.0341 0.7317 0.8556
-1.000 0.1043 0.01785 0.00951 -0.0341 0.7268 0.8612
-0.750 0.1326 0.01786 0.00940 -0.0342 0.7228 0.8666
-0.500 0.1540 0.01782 0.00927 -0.0333 0.7183 0.8729
-0.250 0.1821 0.01783 0.00921 -0.0336 0.7133 0.8761
0.000 0.2094 0.01782 0.00912 -0.0337 0.7091 0.8800
0.250 0.2338 0.01779 0.00900 -0.0334 0.7059 0.8850
0.500 0.2576 0.01783 0.00904 -0.0330 0.7009 0.8890
0.750 0.2853 0.01787 0.00904 -0.0333 0.6966 0.8924
1.000 0.3118 0.01789 0.00901 -0.0333 0.6930 0.8966
1.250 0.3338 0.01792 0.00901 -0.0326 0.6895 0.9017
1.500 0.3611 0.01801 0.00914 -0.0329 0.6848 0.9048
1.750 0.3886 0.01808 0.00921 -0.0331 0.6807 0.9085
2.000 0.4146 0.01813 0.00924 -0.0331 0.6774 0.9128
2.250 0.4384 0.01823 0.00938 -0.0327 0.6734 0.9171
2.500 0.4658 0.01837 0.00957 -0.0331 0.6687 0.9206
2.750 0.4935 0.01846 0.00969 -0.0334 0.6647 0.9246
3.000 0.5197 0.01851 0.00976 -0.0334 0.6614 0.9290
3.250 0.5448 0.01871 0.01007 -0.0334 0.6559 0.9330
3.500 0.5734 0.01883 0.01026 -0.0339 0.6509 0.9369
3.750 0.6025 0.01888 0.01033 -0.0344 0.6469 0.9411
4.000 0.6257 0.01907 0.01067 -0.0340 0.6405 0.9459
4.250 0.6567 0.01914 0.01081 -0.0350 0.6346 0.9492
4.500 0.6860 0.01920 0.01095 -0.0355 0.6283 0.9536
4.750 0.7121 0.01924 0.01109 -0.0354 0.6200 0.9588
5.000 0.7435 0.01923 0.01120 -0.0364 0.6107 0.9623
5.250 0.7767 0.01903 0.01104 -0.0373 0.6014 0.9657
5.500 0.8018 0.01904 0.01120 -0.0371 0.5894 0.9716
5.750 0.8336 0.01896 0.01128 -0.0382 0.5763 0.9756
6.000 0.8634 0.01888 0.01133 -0.0388 0.5623 0.9808
6.250 0.8932 0.01878 0.01138 -0.0394 0.5462 0.9860
6.500 0.9225 0.01868 0.01142 -0.0399 0.5279 0.9919
6.750 0.9470 0.01874 0.01166 -0.0398 0.5036 1.0000
7.000 0.9500 0.01877 0.01175 -0.0354 0.4816 1.0000
7.250 0.9519 0.01885 0.01182 -0.0308 0.4518 1.0000
7.500 0.9570 0.01900 0.01179 -0.0268 0.4048 1.0000
7.750 0.9595 0.01958 0.01199 -0.0228 0.3478 1.0000
8.000 0.9592 0.02057 0.01266 -0.0188 0.2967 1.0000
8.250 0.9573 0.02168 0.01352 -0.0148 0.2539 1.0000
8.500 0.9573 0.02290 0.01453 -0.0114 0.2175 1.0000
8.750 0.9594 0.02418 0.01565 -0.0086 0.1871 1.0000
9.000 0.9628 0.02551 0.01687 -0.0063 0.1616 1.0000
9.250 0.9672 0.02690 0.01816 -0.0042 0.1409 1.0000
9.500 0.9726 0.02833 0.01952 -0.0025 0.1246 1.0000
9.750 0.9788 0.02979 0.02095 -0.0010 0.1107 1.0000
10.000 0.9852 0.03131 0.02247 0.0004 0.0993 1.0000
10.250 0.9933 0.03277 0.02399 0.0016 0.0897 1.0000
10.500 0.9997 0.03442 0.02563 0.0028 0.0829 1.0000
10.750 1.0080 0.03598 0.02727 0.0038 0.0760 1.0000
11.000 1.0151 0.03768 0.02897 0.0047 0.0708 1.0000
11.250 1.0245 0.03925 0.03066 0.0056 0.0655 1.0000
11.500 1.0310 0.04109 0.03249 0.0064 0.0614 1.0000
11.750 1.0414 0.04270 0.03427 0.0072 0.0569 1.0000
12.000 1.0498 0.04445 0.03611 0.0078 0.0527 1.0000
12.250 1.0557 0.04648 0.03812 0.0084 0.0491 1.0000
12.500 1.0646 0.04831 0.04018 0.0088 0.0449 1.0000
12.750 1.0692 0.05044 0.04239 0.0091 0.0413 1.0000
13.000 1.0738 0.05273 0.04475 0.0094 0.0383 1.0000
13.250 1.0795 0.05501 0.04725 0.0097 0.0350 1.0000
13.500 1.0817 0.05754 0.04987 0.0096 0.0325 1.0000
13.750 1.0832 0.06029 0.05263 0.0096 0.0306 1.0000
14.000 1.0870 0.06312 0.05580 0.0098 0.0285 1.0000
14.250 1.0883 0.06620 0.05909 0.0097 0.0269 1.0000
14.500 1.0872 0.06945 0.06248 0.0092 0.0254 1.0000
14.750 1.0858 0.07281 0.06594 0.0086 0.0244 1.0000
15.000 1.0830 0.07646 0.06965 0.0078 0.0234 1.0000
15.250 1.0765 0.08098 0.07449 0.0066 0.0226 1.0000
15.500 1.0678 0.08593 0.07972 0.0050 0.0218 1.0000
15.750 1.0571 0.09137 0.08541 0.0029 0.0212 1.0000
16.000 1.0451 0.09728 0.09156 0.0003 0.0209 1.0000
16.250 1.0307 0.10385 0.09836 -0.0029 0.0206 1.0000
16.500 1.0135 0.11137 0.10610 -0.0070 0.0203 1.0000
16.750 0.9931 0.12013 0.11508 -0.0121 0.0204 1.0000
17.000 0.9664 0.13112 0.12629 -0.0188 0.0208 1.0000
17.250 0.9313 0.14556 0.14091 -0.0276 0.0218 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S805A Airfoil (s805a-nr)