NREL's S829 Airfoil (s829-nr) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S829 Airfoil (s829-nr) Reynolds number: 50,000 Max Cl/Cd: 20.96 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s829-nr-50000-n5.txt Download as CSV file: xf-s829-nr-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S829 Airfoil
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.500 -0.4857 0.11113 0.10441 -0.0684 1.0000 0.0366
-13.250 -0.4989 0.10585 0.09919 -0.0702 1.0000 0.0364
-13.000 -0.5193 0.10040 0.09378 -0.0717 1.0000 0.0362
-12.750 -0.5379 0.09644 0.08986 -0.0717 1.0000 0.0358
-12.500 -0.5582 0.09304 0.08651 -0.0709 1.0000 0.0352
-12.250 -0.5856 0.08936 0.08282 -0.0700 1.0000 0.0350
-12.000 -0.6113 0.08613 0.07958 -0.0683 1.0000 0.0344
-11.750 -0.6394 0.08314 0.07654 -0.0660 1.0000 0.0341
-11.500 -0.6673 0.08040 0.07374 -0.0631 1.0000 0.0339
-11.250 -0.6945 0.07787 0.07113 -0.0596 1.0000 0.0336
-11.000 -0.7147 0.07418 0.06722 -0.0582 0.9964 0.0334
-10.750 -0.7341 0.07056 0.06333 -0.0564 0.9918 0.0337
-10.500 -0.7482 0.06758 0.06009 -0.0540 0.9871 0.0336
-10.250 -0.7620 0.06458 0.05675 -0.0511 0.9826 0.0340
-10.000 -0.7779 0.06223 0.05405 -0.0467 0.9773 0.0348
-9.750 -0.7891 0.05980 0.05118 -0.0424 0.9726 0.0356
-9.500 -0.7900 0.05715 0.04804 -0.0394 0.9692 0.0363
-9.250 -0.7910 0.05465 0.04532 -0.0361 0.9656 0.0375
-9.000 -0.7836 0.05229 0.04277 -0.0340 0.9624 0.0392
-8.750 -0.7680 0.04989 0.04005 -0.0328 0.9598 0.0410
-8.500 -0.7470 0.04768 0.03743 -0.0323 0.9577 0.0452
-8.250 -0.7162 0.04528 0.03471 -0.0334 0.9563 0.0513
-8.000 -0.6608 0.04271 0.03173 -0.0379 0.9563 0.0611
-7.750 -0.6142 0.04093 0.02983 -0.0410 0.9558 0.0751
-7.500 -0.5869 0.03983 0.02873 -0.0411 0.9540 0.0906
-7.250 -0.5664 0.03882 0.02771 -0.0401 0.9519 0.1082
-7.000 -0.5516 0.03784 0.02684 -0.0386 0.9495 0.1330
-6.750 -0.5383 0.03676 0.02593 -0.0369 0.9470 0.1649
-6.500 -0.5290 0.03556 0.02511 -0.0348 0.9445 0.2166
-6.250 -0.5282 0.03432 0.02439 -0.0312 0.9418 0.2901
-6.000 -0.5367 0.03328 0.02397 -0.0254 0.9384 0.3790
-5.750 -0.4442 0.03730 0.02899 -0.0287 0.9402 0.6658
-5.500 -0.4421 0.03806 0.02955 -0.0236 0.9371 0.7067
-5.250 -0.4230 0.03936 0.03057 -0.0208 0.9346 0.7412
-5.000 -0.3967 0.04070 0.03160 -0.0193 0.9326 0.7709
-4.750 -0.3606 0.04200 0.03257 -0.0198 0.9311 0.7961
-4.500 -0.3309 0.04289 0.03314 -0.0197 0.9296 0.8184
-4.250 -0.2854 0.04365 0.03357 -0.0226 0.9287 0.8351
-4.000 -0.2575 0.04401 0.03371 -0.0228 0.9269 0.8497
-3.750 -0.2312 0.04422 0.03371 -0.0230 0.9248 0.8629
-3.500 -0.2067 0.04432 0.03360 -0.0231 0.9225 0.8749
-3.250 -0.1767 0.04434 0.03341 -0.0244 0.9206 0.8849
-3.000 -0.1457 0.04428 0.03318 -0.0261 0.9190 0.8933
-2.750 -0.1223 0.04430 0.03306 -0.0263 0.9172 0.9026
-2.500 -0.0873 0.04418 0.03276 -0.0289 0.9156 0.9091
-2.250 -0.0729 0.04420 0.03267 -0.0275 0.9128 0.9172
-2.000 -0.0501 0.04412 0.03250 -0.0278 0.9105 0.9237
-1.750 -0.0316 0.04414 0.03243 -0.0272 0.9079 0.9306
-1.500 -0.0027 0.04404 0.03223 -0.0287 0.9056 0.9357
-1.250 0.0191 0.04413 0.03224 -0.0288 0.9034 0.9418
-1.000 0.0505 0.04408 0.03212 -0.0308 0.9016 0.9460
-0.750 0.0657 0.04409 0.03210 -0.0297 0.8982 0.9512
-0.500 0.0803 0.04422 0.03220 -0.0285 0.8949 0.9565
-0.250 0.1078 0.04419 0.03214 -0.0298 0.8921 0.9601
0.000 0.1354 0.04427 0.03219 -0.0311 0.8895 0.9639
0.250 0.1585 0.04444 0.03236 -0.0314 0.8868 0.9679
0.500 0.1735 0.04452 0.03246 -0.0304 0.8823 0.9717
0.750 0.1971 0.04462 0.03258 -0.0310 0.8787 0.9750
1.000 0.2247 0.04478 0.03277 -0.0322 0.8755 0.9781
1.250 0.2453 0.04502 0.03306 -0.0321 0.8718 0.9813
1.500 0.2631 0.04517 0.03326 -0.0316 0.8667 0.9843
1.750 0.2901 0.04534 0.03350 -0.0327 0.8626 0.9868
2.000 0.3189 0.04557 0.03381 -0.0341 0.8589 0.9892
2.250 0.3319 0.04583 0.03417 -0.0327 0.8524 0.9922
2.500 0.3606 0.04607 0.03451 -0.0341 0.8479 0.9944
2.750 0.3812 0.04632 0.03486 -0.0340 0.8417 0.9967
3.000 0.4059 0.04657 0.03524 -0.0346 0.8355 0.9990
3.250 0.4294 0.04684 0.03567 -0.0348 0.8294 1.0000
3.500 0.4430 0.04711 0.03606 -0.0332 0.8214 1.0000
3.750 0.4583 0.04736 0.03643 -0.0317 0.8132 1.0000
4.000 0.4826 0.04752 0.03676 -0.0317 0.8056 1.0000
4.250 0.4938 0.04772 0.03713 -0.0295 0.7951 1.0000
4.500 0.5133 0.04779 0.03736 -0.0284 0.7851 1.0000
4.750 0.5491 0.04754 0.03735 -0.0298 0.7763 1.0000
5.000 0.5678 0.04732 0.03732 -0.0283 0.7632 1.0000
5.250 0.5926 0.04678 0.03700 -0.0273 0.7493 1.0000
5.500 0.6127 0.04611 0.03654 -0.0254 0.7330 1.0000
5.750 0.6246 0.04547 0.03615 -0.0222 0.7142 1.0000
6.000 0.6513 0.04413 0.03507 -0.0206 0.6953 1.0000
6.250 0.6555 0.04353 0.03465 -0.0161 0.6727 1.0000
6.500 0.6656 0.04262 0.03395 -0.0123 0.6487 1.0000
6.750 0.6680 0.04209 0.03359 -0.0076 0.6208 1.0000
7.000 0.6701 0.04160 0.03329 -0.0029 0.5882 1.0000
7.250 0.6796 0.04069 0.03256 0.0012 0.5478 1.0000
7.500 0.7509 0.03582 0.02692 0.0020 0.3690 1.0000
7.750 0.7442 0.03697 0.02744 0.0072 0.2924 1.0000
8.000 0.7363 0.03833 0.02839 0.0119 0.2424 1.0000
8.250 0.7312 0.03963 0.02939 0.0161 0.2041 1.0000
8.500 0.7295 0.04085 0.03039 0.0198 0.1757 1.0000
8.750 0.7311 0.04198 0.03140 0.0232 0.1528 1.0000
9.000 0.7338 0.04309 0.03237 0.0263 0.1343 1.0000
9.250 0.7395 0.04413 0.03332 0.0291 0.1187 1.0000
9.500 0.7507 0.04508 0.03426 0.0314 0.1061 1.0000
9.750 0.7625 0.04599 0.03522 0.0336 0.0937 1.0000
10.000 0.7842 0.04686 0.03630 0.0349 0.0827 1.0000
10.250 0.8178 0.04802 0.03764 0.0350 0.0724 1.0000
10.500 0.8330 0.04930 0.03889 0.0365 0.0653 1.0000
10.750 0.8587 0.05111 0.04110 0.0373 0.0589 1.0000
11.000 0.8699 0.05270 0.04272 0.0392 0.0549 1.0000
11.250 0.8720 0.05467 0.04501 0.0422 0.0510 1.0000
11.500 0.8708 0.05662 0.04728 0.0456 0.0481 1.0000
11.750 0.8685 0.05874 0.04965 0.0489 0.0463 1.0000
12.000 0.8629 0.06089 0.05201 0.0523 0.0450 1.0000
12.250 0.8550 0.06303 0.05430 0.0557 0.0439 1.0000
12.500 0.8463 0.06521 0.05659 0.0589 0.0429 1.0000
12.750 0.8345 0.06783 0.05928 0.0620 0.0418 1.0000
13.000 0.8132 0.07058 0.06231 0.0654 0.0414 1.0000
13.250 0.7929 0.07366 0.06561 0.0682 0.0412 1.0000
13.500 0.7692 0.07735 0.06953 0.0702 0.0409 1.0000
13.750 0.7465 0.08132 0.07368 0.0714 0.0408 1.0000
14.000 0.7263 0.08570 0.07821 0.0717 0.0411 1.0000
14.250 0.7020 0.09087 0.08352 0.0709 0.0411 1.0000
14.500 0.6800 0.09655 0.08931 0.0693 0.0414 1.0000
14.750 0.6604 0.10265 0.09548 0.0670 0.0418 1.0000
15.000 0.6380 0.11003 0.10291 0.0635 0.0420 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S829 Airfoil (s829-nr)