FX 61-168 AIRFOIL (fx61168-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: FX 61-168 AIRFOIL (fx61168-il) Reynolds number: 50,000 Max Cl/Cd: 15.36 at α=12° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-fx61168-il-50000-n5.txt Download as CSV file: xf-fx61168-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: FX 61-168 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.3741 0.11123 0.10422 -0.0574 1.0000 0.0424
-11.500 -0.3741 0.10701 0.10006 -0.0583 1.0000 0.0421
-11.250 -0.3780 0.10208 0.09520 -0.0600 1.0000 0.0417
-11.000 -0.3850 0.09698 0.09016 -0.0619 1.0000 0.0413
-10.750 -0.3968 0.09155 0.08479 -0.0642 1.0000 0.0408
-10.500 -0.4130 0.08623 0.07951 -0.0664 1.0000 0.0404
-10.250 -0.4323 0.08144 0.07476 -0.0679 1.0000 0.0400
-10.000 -0.4546 0.07722 0.07056 -0.0685 1.0000 0.0396
-9.750 -0.4781 0.07385 0.06721 -0.0679 1.0000 0.0393
-9.500 -0.5045 0.07126 0.06467 -0.0659 1.0000 0.0391
-9.250 -0.5332 0.06980 0.06328 -0.0622 1.0000 0.0388
-9.000 -0.5581 0.06803 0.06152 -0.0598 0.9968 0.0386
-8.750 -0.5516 0.06280 0.05591 -0.0651 0.9787 0.0385
-8.500 -0.5410 0.05820 0.05088 -0.0687 0.9621 0.0385
-8.250 -0.5274 0.05427 0.04651 -0.0710 0.9467 0.0388
-8.000 -0.5083 0.05061 0.04234 -0.0732 0.9340 0.0401
-7.750 -0.4876 0.04753 0.03869 -0.0746 0.9212 0.0416
-7.500 -0.4632 0.04463 0.03536 -0.0756 0.9100 0.0432
-7.250 -0.4298 0.04190 0.03240 -0.0773 0.9031 0.0448
-7.000 -0.3981 0.03983 0.03012 -0.0780 0.8941 0.0465
-6.750 -0.3564 0.03795 0.02798 -0.0796 0.8893 0.0507
-6.500 -0.3206 0.03670 0.02652 -0.0794 0.8807 0.0549
-6.250 -0.2777 0.03543 0.02523 -0.0803 0.8760 0.0604
-6.000 -0.2483 0.03445 0.02407 -0.0801 0.8670 0.0679
-5.750 -0.2198 0.03312 0.02271 -0.0810 0.8599 0.0773
-5.500 -0.2044 0.03199 0.02157 -0.0802 0.8491 0.0889
-5.250 -0.1877 0.03051 0.02018 -0.0801 0.8403 0.1107
-5.000 -0.1805 0.02834 0.01862 -0.0796 0.8308 0.1750
-4.750 -0.1794 0.02844 0.02083 -0.0725 0.8219 0.4932
-4.500 -0.1533 0.03064 0.02273 -0.0694 0.8151 0.5910
-4.250 -0.1243 0.03473 0.02663 -0.0619 0.8085 0.6464
-4.000 -0.1014 0.03521 0.02681 -0.0603 0.8016 0.6674
-3.750 -0.0677 0.03511 0.02636 -0.0607 0.7964 0.6744
-3.500 -0.0550 0.03493 0.02598 -0.0592 0.7881 0.6846
-3.250 -0.0237 0.03466 0.02542 -0.0599 0.7832 0.6928
-2.750 0.0128 0.03435 0.02471 -0.0580 0.7696 0.7099
-2.500 0.0343 0.03425 0.02443 -0.0573 0.7638 0.7167
-2.250 0.0488 0.03417 0.02421 -0.0560 0.7568 0.7254
-2.000 0.0796 0.03395 0.02378 -0.0565 0.7525 0.7319
-1.750 0.0871 0.03405 0.02379 -0.0541 0.7448 0.7406
-1.250 0.1373 0.03379 0.02323 -0.0539 0.7349 0.7539
-1.000 0.1440 0.03398 0.02337 -0.0512 0.7272 0.7605
-0.750 0.1719 0.03383 0.02308 -0.0515 0.7228 0.7661
-0.500 0.1836 0.03396 0.02313 -0.0501 0.7161 0.7732
-0.250 0.2013 0.03405 0.02316 -0.0487 0.7102 0.7780
0.000 0.2320 0.03388 0.02285 -0.0497 0.7063 0.7837
0.250 0.2361 0.03429 0.02325 -0.0470 0.6983 0.7895
0.500 0.2597 0.03431 0.02320 -0.0467 0.6933 0.7943
0.750 0.2945 0.03411 0.02288 -0.0483 0.6899 0.7993
1.000 0.2919 0.03477 0.02358 -0.0447 0.6806 0.8048
1.250 0.3211 0.03469 0.02342 -0.0451 0.6764 0.8092
1.500 0.3288 0.03524 0.02397 -0.0432 0.6686 0.8147
1.750 0.3520 0.03540 0.02409 -0.0432 0.6631 0.8193
2.000 0.3857 0.03524 0.02389 -0.0442 0.6596 0.8232
2.250 0.3830 0.03621 0.02490 -0.0411 0.6499 0.8288
2.500 0.4157 0.03621 0.02487 -0.0424 0.6458 0.8331
2.750 0.4171 0.03705 0.02576 -0.0394 0.6373 0.8379
3.000 0.4424 0.03727 0.02597 -0.0397 0.6319 0.8424
3.250 0.4818 0.03715 0.02583 -0.0418 0.6288 0.8466
3.500 0.4711 0.03845 0.02722 -0.0377 0.6179 0.8517
3.750 0.5070 0.03832 0.02710 -0.0390 0.6143 0.8558
4.000 0.5044 0.03979 0.02863 -0.0367 0.6036 0.8613
4.250 0.5365 0.03965 0.02851 -0.0373 0.5995 0.8653
4.500 0.5346 0.04120 0.03016 -0.0351 0.5891 0.8708
4.750 0.5670 0.04122 0.03021 -0.0360 0.5846 0.8752
5.250 0.5951 0.04295 0.03209 -0.0346 0.5695 0.8852
5.750 0.6253 0.04485 0.03417 -0.0338 0.5542 0.8956
6.250 0.6567 0.04680 0.03630 -0.0334 0.5387 0.9071
6.750 0.6892 0.04866 0.03839 -0.0330 0.5231 0.9197
7.250 0.7220 0.05054 0.04051 -0.0327 0.5076 0.9348
7.500 0.7204 0.05311 0.04320 -0.0325 0.4954 0.9446
7.750 0.7590 0.05245 0.04271 -0.0334 0.4920 0.9553
8.250 0.7670 0.05736 0.04786 -0.0348 0.4678 1.0000
8.500 0.8069 0.05674 0.04741 -0.0359 0.4632 1.0000
9.000 0.8558 0.05825 0.04920 -0.0378 0.4474 1.0000
9.250 0.8589 0.06116 0.05223 -0.0387 0.4346 1.0000
9.500 0.8654 0.06380 0.05500 -0.0396 0.4228 1.0000
9.750 0.9080 0.06238 0.05379 -0.0401 0.4185 1.0000
10.000 0.9125 0.06514 0.05667 -0.0409 0.4056 1.0000
10.500 0.9625 0.06577 0.05766 -0.0418 0.3889 1.0000
10.750 0.9653 0.06877 0.06079 -0.0425 0.3760 1.0000
11.250 1.0174 0.06877 0.06114 -0.0428 0.3591 1.0000
11.500 1.0196 0.07174 0.06424 -0.0434 0.3460 1.0000
12.000 1.0791 0.07027 0.06316 -0.0429 0.3282 1.0000
12.500 1.0793 0.07671 0.06987 -0.0441 0.3018 1.0000
13.000 1.0975 0.08049 0.07391 -0.0446 0.2766 1.0000
13.250 1.1149 0.08103 0.07458 -0.0444 0.2640 1.0000
13.500 1.1306 0.08178 0.07546 -0.0443 0.2507 1.0000
13.750 1.1453 0.08273 0.07649 -0.0442 0.2374 1.0000
14.000 1.1534 0.08459 0.07840 -0.0444 0.2231 1.0000
14.250 1.1565 0.08725 0.08110 -0.0449 0.2086 1.0000
14.500 1.1585 0.09017 0.08405 -0.0457 0.1944 1.0000
14.750 1.1613 0.09302 0.08690 -0.0464 0.1808 1.0000
15.000 1.1647 0.09583 0.08969 -0.0471 0.1678 1.0000
15.250 1.1617 0.09992 0.09387 -0.0485 0.1561 1.0000
15.500 1.1565 0.10456 0.09861 -0.0502 0.1452 1.0000
15.750 1.1548 0.10859 0.10268 -0.0518 0.1348 1.0000
16.000 1.1573 0.11181 0.10590 -0.0529 0.1248 1.0000
16.250 1.1558 0.11596 0.11012 -0.0547 0.1159 1.0000
16.500 1.1500 0.12113 0.11541 -0.0570 0.1083 1.0000
16.750 1.1569 0.12354 0.11771 -0.0580 0.1003 1.0000
17.000 1.1445 0.13046 0.12492 -0.0616 0.0951 1.0000
17.250 1.1568 0.13174 0.12605 -0.0620 0.0880 1.0000
17.500 1.1395 0.14003 0.13465 -0.0666 0.0846 1.0000
17.750 1.1245 0.14800 0.14283 -0.0713 0.0810 1.0000
18.000 1.1437 0.14765 0.14234 -0.0708 0.0752 1.0000
18.250 1.1172 0.15889 0.15385 -0.0777 0.0739 1.0000
18.500 1.0761 0.17526 0.17033 -0.0879 0.0732 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 61-168 AIRFOIL (fx61168-il)