GOE 563 AIRFOIL (goe563-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 563 AIRFOIL (goe563-il) Reynolds number: 100,000 Max Cl/Cd: 50.95 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe563-il-100000-n5.txt Download as CSV file: xf-goe563-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 563 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4680 0.10721 0.10177 -0.0274 1.0000 0.0400
-10.250 -0.4725 0.10213 0.09674 -0.0297 1.0000 0.0393
-10.000 -0.4770 0.09716 0.09182 -0.0320 1.0000 0.0389
-9.750 -0.4770 0.09350 0.08820 -0.0331 1.0000 0.0393
-9.500 -0.4749 0.09058 0.08531 -0.0338 1.0000 0.0406
-9.250 -0.4771 0.08690 0.08168 -0.0351 1.0000 0.0415
-9.000 -0.4834 0.08263 0.07748 -0.0368 1.0000 0.0419
-8.750 -0.4937 0.07818 0.07311 -0.0387 1.0000 0.0421
-8.500 -0.5107 0.07347 0.06850 -0.0408 1.0000 0.0419
-8.250 -0.5271 0.06739 0.06245 -0.0440 1.0000 0.0416
-8.000 -0.5420 0.06147 0.05647 -0.0457 1.0000 0.0414
-7.750 -0.5539 0.05607 0.05093 -0.0459 1.0000 0.0416
-7.500 -0.5624 0.05095 0.04559 -0.0452 1.0000 0.0419
-7.250 -0.5657 0.04646 0.04080 -0.0439 1.0000 0.0426
-7.000 -0.5650 0.04235 0.03630 -0.0425 1.0000 0.0445
-6.750 -0.5617 0.03810 0.03150 -0.0408 1.0000 0.0461
-6.500 -0.5536 0.03446 0.02725 -0.0390 1.0000 0.0472
-6.250 -0.5414 0.03171 0.02400 -0.0373 1.0000 0.0484
-6.000 -0.5257 0.03042 0.02267 -0.0360 1.0000 0.0503
-5.750 -0.5089 0.02920 0.02127 -0.0347 1.0000 0.0523
-5.500 -0.4912 0.02756 0.01929 -0.0334 1.0000 0.0537
-5.250 -0.4626 0.02578 0.01712 -0.0340 0.9971 0.0553
-5.000 -0.4310 0.02435 0.01532 -0.0351 0.9935 0.0577
-4.750 -0.3995 0.02314 0.01377 -0.0361 0.9894 0.0603
-4.500 -0.3674 0.02204 0.01258 -0.0374 0.9859 0.0623
-4.250 -0.3364 0.02118 0.01164 -0.0382 0.9815 0.0644
-4.000 -0.3042 0.02040 0.01077 -0.0393 0.9771 0.0670
-3.750 -0.2712 0.01970 0.00993 -0.0404 0.9732 0.0702
-3.500 -0.2414 0.01909 0.00929 -0.0410 0.9677 0.0744
-3.250 -0.2074 0.01868 0.00888 -0.0425 0.9634 0.0819
-3.000 -0.1777 0.01819 0.00838 -0.0430 0.9577 0.0897
-2.750 -0.1461 0.01770 0.00789 -0.0439 0.9523 0.1024
-2.500 -0.1107 0.01723 0.00747 -0.0456 0.9484 0.1221
-2.250 -0.0837 0.01690 0.00717 -0.0455 0.9410 0.1442
-2.000 -0.0501 0.01650 0.00696 -0.0470 0.9360 0.1820
-1.750 -0.0189 0.01590 0.00675 -0.0480 0.9307 0.2571
-1.500 0.0074 0.01491 0.00674 -0.0481 0.9242 0.4886
-1.250 0.0381 0.01418 0.00676 -0.0480 0.9202 0.6833
-1.000 0.0979 0.01381 0.00685 -0.0533 0.9204 0.9028
-0.750 0.1699 0.01379 0.00671 -0.0623 0.9212 1.0000
-0.500 0.1999 0.01379 0.00659 -0.0628 0.9132 1.0000
-0.250 0.2336 0.01376 0.00646 -0.0639 0.9063 1.0000
0.000 0.2627 0.01375 0.00637 -0.0641 0.8973 1.0000
0.250 0.3004 0.01361 0.00615 -0.0658 0.8893 1.0000
0.500 0.3309 0.01345 0.00592 -0.0658 0.8757 1.0000
0.750 0.3611 0.01328 0.00570 -0.0658 0.8613 1.0000
1.000 0.3886 0.01320 0.00558 -0.0653 0.8478 1.0000
1.250 0.4152 0.01317 0.00552 -0.0647 0.8350 1.0000
1.500 0.4419 0.01314 0.00548 -0.0641 0.8218 1.0000
1.750 0.4687 0.01312 0.00544 -0.0635 0.8079 1.0000
2.000 0.4954 0.01310 0.00542 -0.0629 0.7933 1.0000
2.250 0.5219 0.01310 0.00542 -0.0622 0.7776 1.0000
2.500 0.5485 0.01310 0.00542 -0.0615 0.7608 1.0000
2.750 0.5753 0.01310 0.00541 -0.0608 0.7423 1.0000
3.000 0.6000 0.01314 0.00545 -0.0597 0.7188 1.0000
3.250 0.6248 0.01319 0.00544 -0.0585 0.6890 1.0000
3.500 0.6489 0.01328 0.00541 -0.0571 0.6513 1.0000
3.750 0.6716 0.01346 0.00547 -0.0556 0.6075 1.0000
4.000 0.6940 0.01371 0.00559 -0.0542 0.5650 1.0000
4.250 0.7158 0.01405 0.00577 -0.0527 0.5218 1.0000
4.500 0.7368 0.01446 0.00604 -0.0512 0.4787 1.0000
4.750 0.7570 0.01493 0.00634 -0.0496 0.4349 1.0000
5.000 0.7768 0.01547 0.00670 -0.0480 0.3940 1.0000
5.250 0.7966 0.01603 0.00712 -0.0465 0.3585 1.0000
5.500 0.8162 0.01662 0.00760 -0.0450 0.3263 1.0000
5.750 0.8352 0.01727 0.00810 -0.0435 0.2918 1.0000
6.000 0.8544 0.01790 0.00861 -0.0421 0.2556 1.0000
6.250 0.8735 0.01854 0.00913 -0.0407 0.2215 1.0000
6.500 0.8931 0.01917 0.00969 -0.0394 0.1931 1.0000
6.750 0.9127 0.01983 0.01033 -0.0381 0.1689 1.0000
7.000 0.9314 0.02058 0.01099 -0.0367 0.1450 1.0000
7.250 0.9495 0.02142 0.01175 -0.0353 0.1220 1.0000
7.500 0.9666 0.02237 0.01264 -0.0338 0.1027 1.0000
7.750 0.9840 0.02333 0.01363 -0.0322 0.0838 1.0000
8.000 1.0006 0.02433 0.01464 -0.0306 0.0701 1.0000
8.250 1.0163 0.02540 0.01570 -0.0288 0.0618 1.0000
8.500 1.0315 0.02646 0.01679 -0.0271 0.0550 1.0000
8.750 1.0454 0.02767 0.01808 -0.0252 0.0502 1.0000
9.000 1.0593 0.02885 0.01935 -0.0233 0.0459 1.0000
9.250 1.0695 0.03036 0.02085 -0.0211 0.0425 1.0000
9.500 1.0834 0.03172 0.02245 -0.0192 0.0397 1.0000
9.750 1.0954 0.03318 0.02405 -0.0172 0.0372 1.0000
10.000 1.1054 0.03467 0.02561 -0.0151 0.0350 1.0000
10.250 1.1145 0.03665 0.02763 -0.0132 0.0329 1.0000
10.500 1.1259 0.03844 0.02970 -0.0112 0.0313 1.0000
10.750 1.1356 0.04049 0.03200 -0.0093 0.0299 1.0000
11.000 1.1424 0.04248 0.03420 -0.0073 0.0286 1.0000
11.250 1.1468 0.04446 0.03634 -0.0053 0.0274 1.0000
11.500 1.1498 0.04656 0.03855 -0.0034 0.0263 1.0000
11.750 1.1510 0.04941 0.04149 -0.0018 0.0253 1.0000
12.000 1.1461 0.05211 0.04454 0.0002 0.0247 1.0000
12.250 1.1393 0.05529 0.04804 0.0017 0.0244 1.0000
12.500 1.1293 0.05881 0.05186 0.0028 0.0241 1.0000
12.750 1.1157 0.06289 0.05624 0.0031 0.0239 1.0000
13.000 1.0997 0.06754 0.06116 0.0026 0.0237 1.0000
13.250 1.0829 0.07267 0.06653 0.0012 0.0237 1.0000
13.500 1.0627 0.07879 0.07290 -0.0015 0.0237 1.0000
13.750 1.0415 0.08572 0.08003 -0.0053 0.0238 1.0000
14.000 1.0177 0.09397 0.08847 -0.0105 0.0240 1.0000
14.250 0.9932 0.10333 0.09800 -0.0167 0.0243 1.0000
14.500 0.9654 0.11459 0.10935 -0.0241 0.0246 1.0000
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Polar data table (+)
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