RAE6-9CK AIRFOIL (rae69ck-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE6-9CK AIRFOIL (rae69ck-il) Reynolds number: 50,000 Max Cl/Cd: 27.05 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae69ck-il-50000-n5.txt Download as CSV file: xf-rae69ck-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE6-9CK 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7040 0.07584 0.06790 -0.0584 1.0000 0.0460
-11.000 -0.7218 0.07171 0.06366 -0.0591 1.0000 0.0459
-10.750 -0.7408 0.06805 0.05987 -0.0587 1.0000 0.0459
-10.500 -0.7589 0.06498 0.05667 -0.0572 1.0000 0.0458
-10.250 -0.7760 0.06198 0.05348 -0.0550 1.0000 0.0460
-10.000 -0.7838 0.05891 0.05027 -0.0531 1.0000 0.0465
-9.750 -0.7839 0.05632 0.04763 -0.0515 1.0000 0.0478
-9.500 -0.7832 0.05394 0.04512 -0.0497 1.0000 0.0492
-9.250 -0.7819 0.05152 0.04251 -0.0478 1.0000 0.0509
-9.000 -0.7792 0.04888 0.03958 -0.0458 1.0000 0.0523
-8.750 -0.7729 0.04616 0.03651 -0.0438 1.0000 0.0537
-8.500 -0.7625 0.04350 0.03346 -0.0419 1.0000 0.0552
-8.250 -0.7499 0.04124 0.03072 -0.0400 1.0000 0.0575
-8.000 -0.7340 0.03933 0.02885 -0.0387 1.0000 0.0610
-7.750 -0.7168 0.03768 0.02704 -0.0371 1.0000 0.0652
-7.500 -0.6947 0.03604 0.02510 -0.0354 1.0000 0.0693
-7.250 -0.6764 0.03471 0.02385 -0.0337 1.0000 0.0749
-7.000 -0.6585 0.03361 0.02257 -0.0317 1.0000 0.0821
-6.750 -0.6445 0.03242 0.02143 -0.0293 1.0000 0.0883
-6.500 -0.6329 0.03135 0.02024 -0.0269 1.0000 0.0980
-6.250 -0.6238 0.03017 0.01911 -0.0245 1.0000 0.1084
-6.000 -0.6152 0.02890 0.01790 -0.0222 1.0000 0.1216
-5.750 -0.6069 0.02748 0.01662 -0.0202 1.0000 0.1421
-5.500 -0.5984 0.02576 0.01518 -0.0186 1.0000 0.1762
-5.250 -0.5917 0.02335 0.01362 -0.0175 1.0000 0.2717
-5.000 -0.5845 0.02319 0.01466 -0.0129 1.0000 0.4741
-4.750 -0.5681 0.02366 0.01495 -0.0103 1.0000 0.5374
-4.500 -0.5519 0.02418 0.01528 -0.0076 1.0000 0.5759
-4.250 -0.5366 0.02483 0.01580 -0.0042 1.0000 0.6044
-4.000 -0.5235 0.02571 0.01664 0.0003 1.0000 0.6302
-3.750 -0.5116 0.02654 0.01739 0.0051 1.0000 0.6558
-3.500 -0.4986 0.02700 0.01774 0.0088 1.0000 0.6790
-3.250 -0.4834 0.02712 0.01773 0.0116 1.0000 0.6946
-3.000 -0.4651 0.02693 0.01739 0.0132 1.0000 0.7010
-2.750 -0.4440 0.02650 0.01676 0.0132 1.0000 0.7082
-2.500 -0.4239 0.02620 0.01628 0.0138 1.0000 0.7136
-2.250 -0.4032 0.02591 0.01584 0.0142 1.0000 0.7194
-2.000 -0.3692 0.02574 0.01547 0.0117 0.9950 0.7255
-1.750 -0.3381 0.02568 0.01526 0.0103 0.9896 0.7301
-1.500 -0.3047 0.02562 0.01505 0.0081 0.9844 0.7359
-1.250 -0.2730 0.02552 0.01483 0.0063 0.9787 0.7414
-1.000 -0.2416 0.02552 0.01473 0.0048 0.9732 0.7464
-0.750 -0.2086 0.02549 0.01460 0.0027 0.9679 0.7525
-0.500 -0.1788 0.02548 0.01454 0.0015 0.9617 0.7577
-0.250 -0.1448 0.02557 0.01457 -0.0005 0.9566 0.7632
0.000 -0.1158 0.02553 0.01449 -0.0019 0.9501 0.7696
0.250 -0.0845 0.02562 0.01459 -0.0031 0.9442 0.7749
0.500 -0.0540 0.02568 0.01464 -0.0045 0.9380 0.7814
0.750 -0.0242 0.02576 0.01475 -0.0057 0.9314 0.7874
1.000 0.0083 0.02589 0.01491 -0.0073 0.9255 0.7940
1.250 0.0358 0.02596 0.01504 -0.0080 0.9179 0.8010
1.500 0.0708 0.02612 0.01528 -0.0099 0.9123 0.8078
1.750 0.0957 0.02620 0.01544 -0.0101 0.9035 0.8156
2.000 0.1306 0.02635 0.01570 -0.0119 0.8973 0.8231
2.250 0.1556 0.02645 0.01594 -0.0120 0.8879 0.8314
2.500 0.1871 0.02657 0.01620 -0.0132 0.8799 0.8403
2.750 0.2171 0.02666 0.01646 -0.0139 0.8709 0.8493
3.000 0.2447 0.02675 0.01674 -0.0143 0.8608 0.8596
3.250 0.2801 0.02679 0.01699 -0.0158 0.8519 0.8701
3.500 0.3112 0.02677 0.01721 -0.0164 0.8406 0.8816
3.750 0.3421 0.02671 0.01743 -0.0169 0.8277 0.8945
4.000 0.3782 0.02651 0.01752 -0.0181 0.8135 0.9081
4.250 0.4236 0.02593 0.01729 -0.0201 0.7966 0.9215
4.500 0.4778 0.02433 0.01609 -0.0216 0.7631 0.9333
4.750 0.5218 0.02240 0.01443 -0.0205 0.7035 0.9483
5.000 0.5577 0.02137 0.01355 -0.0197 0.6141 0.9728
5.250 0.5875 0.02172 0.01244 -0.0175 0.3415 1.0000
5.500 0.5894 0.02375 0.01340 -0.0148 0.2041 1.0000
5.750 0.6019 0.02553 0.01461 -0.0137 0.1462 1.0000
6.000 0.6203 0.02703 0.01591 -0.0131 0.1196 1.0000
6.250 0.6423 0.02843 0.01721 -0.0128 0.1024 1.0000
6.500 0.6686 0.02989 0.01868 -0.0130 0.0907 1.0000
6.750 0.7007 0.03131 0.02023 -0.0137 0.0802 1.0000
7.000 0.7375 0.03315 0.02216 -0.0150 0.0727 1.0000
7.250 0.7707 0.03502 0.02415 -0.0160 0.0665 1.0000
7.500 0.8013 0.03724 0.02655 -0.0167 0.0614 1.0000
7.750 0.8303 0.03983 0.02953 -0.0168 0.0583 1.0000
8.000 0.8539 0.04234 0.03233 -0.0165 0.0554 1.0000
8.250 0.8739 0.04474 0.03483 -0.0163 0.0524 1.0000
8.500 0.8887 0.04777 0.03826 -0.0151 0.0504 1.0000
8.750 0.8992 0.05112 0.04217 -0.0134 0.0493 1.0000
9.000 0.9054 0.05465 0.04617 -0.0115 0.0487 1.0000
9.250 0.9071 0.05830 0.05024 -0.0094 0.0483 1.0000
9.500 0.9038 0.06200 0.05431 -0.0073 0.0479 1.0000
9.750 0.8958 0.06569 0.05833 -0.0051 0.0475 1.0000
10.000 0.8824 0.06930 0.06219 -0.0028 0.0472 1.0000
10.250 0.8655 0.07305 0.06615 -0.0010 0.0471 1.0000
10.500 0.8462 0.07729 0.07056 -0.0004 0.0472 1.0000
10.750 0.8248 0.08222 0.07564 -0.0011 0.0475 1.0000
11.000 0.8030 0.08796 0.08149 -0.0034 0.0480 1.0000
11.250 0.7822 0.09466 0.08826 -0.0071 0.0486 1.0000
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Polar data table (+)
Polar graphs
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