DU 06-W-200 VAWT airfoil (du06-w-200-dt) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: DU 06-W-200 VAWT airfoil (du06-w-200-dt) Reynolds number: 100,000 Max Cl/Cd: 36.15 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-du06-w-200-dt-100000-n5.txt Download as CSV file: xf-du06-w-200-dt-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DU 06-W-200 VAWT 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
------ -------- --------- --------- -------- -------- --------
-18.750 -0.9673 0.13320 0.12733 0.0094 1.0000 0.0650
-18.500 -0.9997 0.12281 0.11670 0.0035 1.0000 0.0667
-18.250 -0.9926 0.12048 0.11444 0.0025 1.0000 0.0689
-18.000 -1.0015 0.11504 0.10892 -0.0004 1.0000 0.0710
-17.750 -1.0177 0.10830 0.10202 -0.0038 1.0000 0.0732
-17.500 -1.0356 0.10144 0.09493 -0.0072 1.0000 0.0755
-17.250 -1.0406 0.09709 0.09053 -0.0090 1.0000 0.0777
-17.000 -1.0439 0.09307 0.08649 -0.0107 1.0000 0.0801
-16.750 -1.0503 0.08854 0.08185 -0.0126 1.0000 0.0827
-16.500 -1.0586 0.08374 0.07686 -0.0146 1.0000 0.0856
-16.250 -1.0632 0.07967 0.07268 -0.0160 1.0000 0.0884
-16.000 -1.0641 0.07631 0.06931 -0.0171 1.0000 0.0912
-15.750 -1.0668 0.07266 0.06558 -0.0184 1.0000 0.0944
-15.500 -1.0700 0.06896 0.06170 -0.0195 1.0000 0.0980
-15.250 -1.0697 0.06593 0.05863 -0.0202 1.0000 0.1013
-15.000 -1.0691 0.06301 0.05568 -0.0209 1.0000 0.1048
-14.750 -1.0688 0.06006 0.05262 -0.0216 1.0000 0.1089
-14.500 -1.0674 0.05730 0.04972 -0.0220 1.0000 0.1130
-14.250 -1.0641 0.05497 0.04743 -0.0222 1.0000 0.1169
-14.000 -1.0614 0.05260 0.04495 -0.0224 1.0000 0.1214
-13.750 -1.0578 0.05031 0.04251 -0.0225 1.0000 0.1262
-13.500 -1.0524 0.04840 0.04066 -0.0223 1.0000 0.1305
-13.250 -1.0472 0.04643 0.03861 -0.0222 1.0000 0.1356
-13.000 -1.0404 0.04457 0.03664 -0.0220 1.0000 0.1411
-12.750 -1.0334 0.04289 0.03501 -0.0217 1.0000 0.1460
-12.500 -1.0255 0.04122 0.03325 -0.0214 1.0000 0.1521
-12.250 -1.0164 0.03970 0.03170 -0.0210 1.0000 0.1579
-12.000 -1.0073 0.03822 0.03022 -0.0206 1.0000 0.1639
-11.750 -0.9965 0.03681 0.02871 -0.0203 1.0000 0.1708
-11.500 -0.9859 0.03551 0.02748 -0.0198 1.0000 0.1769
-11.250 -0.9741 0.03424 0.02612 -0.0194 1.0000 0.1844
-11.000 -0.9619 0.03306 0.02501 -0.0189 1.0000 0.1910
-10.750 -0.9493 0.03191 0.02380 -0.0184 1.0000 0.1990
-10.500 -0.9360 0.03083 0.02276 -0.0179 1.0000 0.2065
-10.250 -0.9224 0.02979 0.02170 -0.0174 1.0000 0.2151
-10.000 -0.9086 0.02880 0.02077 -0.0168 1.0000 0.2233
-9.750 -0.8946 0.02785 0.01980 -0.0163 1.0000 0.2327
-9.500 -0.8812 0.02696 0.01899 -0.0155 1.0000 0.2415
-9.250 -0.8679 0.02613 0.01814 -0.0148 1.0000 0.2517
-9.000 -0.8547 0.02534 0.01745 -0.0140 1.0000 0.2615
-8.750 -0.8371 0.02462 0.01679 -0.0138 0.9855 0.2724
-8.500 -0.8058 0.02390 0.01605 -0.0160 0.9441 0.2866
-8.250 -0.7779 0.02326 0.01537 -0.0171 0.9115 0.3009
-8.000 -0.7548 0.02271 0.01482 -0.0170 0.8841 0.3146
-7.750 -0.7339 0.02224 0.01432 -0.0163 0.8612 0.3285
-7.500 -0.7135 0.02181 0.01386 -0.0154 0.8412 0.3427
-7.250 -0.6928 0.02141 0.01344 -0.0146 0.8238 0.3578
-7.000 -0.6714 0.02106 0.01305 -0.0138 0.8083 0.3733
-6.750 -0.6492 0.02073 0.01269 -0.0130 0.7937 0.3891
-6.500 -0.6263 0.02044 0.01238 -0.0124 0.7795 0.4054
-6.250 -0.6027 0.02019 0.01210 -0.0118 0.7663 0.4214
-6.000 -0.5787 0.01997 0.01185 -0.0113 0.7545 0.4377
-5.750 -0.5539 0.01980 0.01166 -0.0108 0.7425 0.4539
-5.500 -0.5287 0.01968 0.01154 -0.0103 0.7308 0.4697
-5.250 -0.5031 0.01967 0.01151 -0.0096 0.7204 0.4847
-5.000 -0.4766 0.01974 0.01161 -0.0090 0.7087 0.4988
-4.750 -0.4498 0.01988 0.01174 -0.0084 0.6983 0.5122
-4.500 -0.4229 0.02004 0.01187 -0.0078 0.6886 0.5253
-4.250 -0.3957 0.02017 0.01197 -0.0074 0.6786 0.5385
-4.000 -0.3679 0.02055 0.01232 -0.0066 0.6699 0.5480
-3.750 -0.3402 0.02078 0.01252 -0.0063 0.6597 0.5584
-3.500 -0.3135 0.02083 0.01245 -0.0061 0.6519 0.5707
-3.250 -0.2850 0.02126 0.01287 -0.0055 0.6424 0.5777
-3.000 -0.2581 0.02131 0.01279 -0.0054 0.6352 0.5886
-2.750 -0.2301 0.02161 0.01305 -0.0051 0.6265 0.5963
-2.500 -0.2029 0.02169 0.01303 -0.0050 0.6191 0.6057
-2.250 -0.1752 0.02193 0.01322 -0.0048 0.6117 0.6132
-2.000 -0.1478 0.02200 0.01321 -0.0048 0.6043 0.6218
-1.750 -0.1203 0.02226 0.01338 -0.0043 0.5983 0.6279
-1.500 -0.0929 0.02224 0.01331 -0.0046 0.5906 0.6372
-1.250 -0.0655 0.02252 0.01355 -0.0040 0.5844 0.6419
-1.000 -0.0380 0.02239 0.01331 -0.0046 0.5782 0.6523
-0.750 -0.0108 0.02268 0.01360 -0.0039 0.5712 0.6558
-0.500 0.0164 0.02284 0.01369 -0.0035 0.5657 0.6611
-0.250 0.0440 0.02274 0.01353 -0.0041 0.5592 0.6698
0.000 0.0708 0.02295 0.01374 -0.0035 0.5527 0.6732
0.250 0.0980 0.02305 0.01377 -0.0031 0.5476 0.6782
0.500 0.1259 0.02295 0.01363 -0.0039 0.5408 0.6864
0.750 0.1524 0.02311 0.01380 -0.0033 0.5347 0.6895
1.000 0.1794 0.02320 0.01382 -0.0028 0.5296 0.6936
1.250 0.2067 0.02321 0.01385 -0.0032 0.5225 0.6997
1.500 0.2342 0.02320 0.01381 -0.0033 0.5164 0.7048
1.750 0.2608 0.02329 0.01385 -0.0028 0.5112 0.7081
2.000 0.2872 0.02338 0.01402 -0.0027 0.5041 0.7124
2.250 0.3156 0.02333 0.01390 -0.0032 0.4980 0.7182
2.500 0.3427 0.02337 0.01391 -0.0032 0.4922 0.7219
2.750 0.3683 0.02347 0.01410 -0.0027 0.4850 0.7251
3.000 0.3954 0.02350 0.01409 -0.0026 0.4791 0.7289
3.250 0.4229 0.02355 0.01416 -0.0029 0.4723 0.7336
3.500 0.4508 0.02356 0.01418 -0.0034 0.4652 0.7378
3.750 0.4772 0.02359 0.01417 -0.0029 0.4598 0.7405
4.000 0.5024 0.02372 0.01443 -0.0027 0.4517 0.7437
4.250 0.5295 0.02374 0.01443 -0.0027 0.4450 0.7472
4.500 0.5572 0.02381 0.01454 -0.0032 0.4377 0.7512
4.750 0.5847 0.02386 0.01461 -0.0035 0.4300 0.7546
5.000 0.6102 0.02391 0.01467 -0.0031 0.4233 0.7571
5.250 0.6352 0.02404 0.01491 -0.0029 0.4147 0.7599
5.500 0.6621 0.02405 0.01488 -0.0028 0.4079 0.7629
5.750 0.6877 0.02421 0.01516 -0.0030 0.3983 0.7662
6.000 0.7161 0.02424 0.01514 -0.0035 0.3912 0.7694
6.250 0.7407 0.02443 0.01546 -0.0034 0.3814 0.7719
6.500 0.7658 0.02450 0.01552 -0.0031 0.3740 0.7741
6.750 0.7897 0.02472 0.01586 -0.0028 0.3643 0.7763
7.000 0.8149 0.02486 0.01599 -0.0027 0.3562 0.7788
7.250 0.8393 0.02511 0.01631 -0.0027 0.3465 0.7814
7.500 0.8644 0.02534 0.01655 -0.0028 0.3380 0.7837
7.750 0.8893 0.02562 0.01685 -0.0030 0.3288 0.7859
8.000 0.9127 0.02595 0.01722 -0.0029 0.3201 0.7882
8.250 0.9342 0.02627 0.01758 -0.0024 0.3117 0.7901
8.500 0.9554 0.02669 0.01805 -0.0019 0.3033 0.7920
8.750 0.9767 0.02711 0.01847 -0.0015 0.2951 0.7939
9.000 0.9971 0.02763 0.01906 -0.0012 0.2868 0.7958
9.250 1.0172 0.02814 0.01959 -0.0009 0.2790 0.7979
9.500 1.0365 0.02877 0.02027 -0.0005 0.2713 0.8004
9.750 1.0551 0.02943 0.02096 -0.0002 0.2638 0.8030
10.000 1.0723 0.03011 0.02167 0.0003 0.2572 0.8050
10.250 1.0850 0.03088 0.02255 0.0014 0.2503 0.8068
10.500 1.0993 0.03158 0.02321 0.0023 0.2445 0.8088
10.750 1.1112 0.03267 0.02445 0.0029 0.2378 0.8109
11.000 1.1242 0.03375 0.02557 0.0033 0.2316 0.8132
11.250 1.1379 0.03485 0.02669 0.0035 0.2261 0.8156
11.500 1.1476 0.03634 0.02831 0.0035 0.2199 0.8183
11.750 1.1593 0.03764 0.02964 0.0036 0.2146 0.8210
12.000 1.1686 0.03908 0.03114 0.0039 0.2096 0.8235
12.250 1.1745 0.04091 0.03312 0.0039 0.2043 0.8262
12.500 1.1832 0.04251 0.03477 0.0040 0.1995 0.8289
12.750 1.1931 0.04407 0.03635 0.0040 0.1951 0.8316
13.000 1.1951 0.04652 0.03897 0.0035 0.1901 0.8344
13.500 1.2111 0.05001 0.04251 0.0033 0.1816 0.8399
13.750 1.2087 0.05296 0.04566 0.0027 0.1774 0.8429
14.000 1.2073 0.05587 0.04872 0.0021 0.1732 0.8461
14.250 1.2110 0.05818 0.05110 0.0017 0.1692 0.8497
14.750 1.2091 0.06402 0.05717 0.0004 0.1619 0.8573
15.000 1.2023 0.06773 0.06105 -0.0007 0.1581 0.8616
15.250 1.2036 0.07047 0.06386 -0.0014 0.1546 0.8668
15.500 1.2178 0.07139 0.06474 -0.0010 0.1515 0.8730
15.750 1.1944 0.07765 0.07129 -0.0034 0.1481 0.8782
16.000 1.1751 0.08358 0.07744 -0.0058 0.1446 0.8843
16.250 1.1681 0.08765 0.08164 -0.0071 0.1413 0.8929
16.500 1.1777 0.08896 0.08298 -0.0070 0.1385 0.9098
16.750 1.1625 0.09344 0.08762 -0.0083 0.1357 0.9997
17.250 1.0215 0.12500 0.11974 -0.0243 0.1258 1.0000
17.500 1.0708 0.12011 0.11477 -0.0220 0.1246 1.0000
|
Polar data table (+)
Polar graphs
<< Back to DU 06-W-200 VAWT airfoil (du06-w-200-dt)