USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 100,000 Max Cl/Cd: 48.24 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa51-il-100000.txt Download as CSV file: xf-usa51-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4326 0.10328 0.09833 -0.0198 1.0000 0.0922
-9.000 -0.4324 0.10022 0.09530 -0.0209 1.0000 0.0961
-8.750 -0.4484 0.09772 0.09294 -0.0252 1.0000 0.0993
-8.500 -0.4666 0.09502 0.09037 -0.0290 1.0000 0.0998
-8.250 -0.4373 0.09015 0.08543 -0.0235 1.0000 0.1052
-8.000 -0.4388 0.08724 0.08258 -0.0238 1.0000 0.1088
-7.750 -0.4548 0.08443 0.07989 -0.0256 1.0000 0.1119
-7.500 -0.4828 0.08136 0.07683 -0.0323 1.0000 0.1139
-7.250 -0.4689 0.07684 0.07240 -0.0285 1.0000 0.1170
-7.000 -0.4622 0.07401 0.06961 -0.0267 1.0000 0.1215
-6.750 -0.4901 0.07186 0.06712 -0.0328 1.0000 0.1286
-6.500 -0.4701 0.06672 0.06228 -0.0286 1.0000 0.1316
-6.250 -0.4638 0.06398 0.05955 -0.0272 1.0000 0.1377
-6.000 -0.4653 0.06030 0.05577 -0.0278 1.0000 0.1454
-5.500 -0.4524 0.05436 0.04976 -0.0253 1.0000 0.1616
-5.250 -0.4465 0.05149 0.04676 -0.0244 1.0000 0.1743
-5.000 -0.4382 0.04895 0.04419 -0.0227 1.0000 0.1891
-4.500 -0.3966 0.03388 0.02705 -0.0228 1.0000 0.0840
-4.250 -0.3780 0.02993 0.02217 -0.0203 1.0000 0.0735
-4.000 -0.3583 0.02769 0.01959 -0.0187 1.0000 0.0721
-3.750 -0.3380 0.02564 0.01713 -0.0170 1.0000 0.0723
-3.500 -0.3182 0.02366 0.01501 -0.0159 1.0000 0.0761
-3.250 -0.2968 0.02235 0.01355 -0.0146 1.0000 0.0793
-3.000 -0.2745 0.02116 0.01215 -0.0133 1.0000 0.0823
-2.750 -0.2534 0.02018 0.01104 -0.0120 1.0000 0.0884
-2.500 -0.2334 0.01939 0.01028 -0.0108 1.0000 0.0958
-2.250 -0.2137 0.01859 0.00947 -0.0094 1.0000 0.1033
-2.000 -0.1955 0.01801 0.00900 -0.0080 1.0000 0.1183
-1.750 -0.1557 0.01692 0.00834 -0.0105 0.9924 0.1782
-1.500 -0.0802 0.01398 0.00819 -0.0181 0.9967 1.0000
-1.250 -0.0227 0.01431 0.00815 -0.0244 0.9852 1.0000
-1.000 0.0328 0.01451 0.00809 -0.0303 0.9727 1.0000
-0.750 0.0875 0.01458 0.00794 -0.0358 0.9592 1.0000
-0.500 0.1413 0.01454 0.00774 -0.0410 0.9451 1.0000
-0.250 0.1952 0.01437 0.00745 -0.0460 0.9305 1.0000
0.000 0.2511 0.01405 0.00705 -0.0511 0.9154 1.0000
0.250 0.3011 0.01365 0.00658 -0.0549 0.8968 1.0000
0.500 0.3461 0.01322 0.00610 -0.0574 0.8734 1.0000
0.750 0.3884 0.01284 0.00564 -0.0593 0.8452 1.0000
1.000 0.4261 0.01261 0.00528 -0.0603 0.8115 1.0000
1.250 0.4565 0.01261 0.00509 -0.0598 0.7732 1.0000
1.500 0.4815 0.01279 0.00505 -0.0586 0.7354 1.0000
1.750 0.5045 0.01302 0.00509 -0.0571 0.7013 1.0000
2.000 0.5267 0.01327 0.00515 -0.0555 0.6710 1.0000
2.250 0.5485 0.01352 0.00525 -0.0541 0.6437 1.0000
2.500 0.5708 0.01378 0.00539 -0.0527 0.6200 1.0000
2.750 0.5934 0.01405 0.00553 -0.0515 0.5995 1.0000
3.000 0.6162 0.01434 0.00575 -0.0503 0.5807 1.0000
3.250 0.6388 0.01463 0.00598 -0.0491 0.5624 1.0000
3.500 0.6613 0.01493 0.00620 -0.0479 0.5444 1.0000
3.750 0.6838 0.01523 0.00640 -0.0467 0.5264 1.0000
4.000 0.7059 0.01554 0.00671 -0.0455 0.5086 1.0000
4.250 0.7285 0.01586 0.00703 -0.0444 0.4923 1.0000
4.500 0.7509 0.01619 0.00736 -0.0432 0.4758 1.0000
4.750 0.7733 0.01653 0.00771 -0.0421 0.4593 1.0000
5.000 0.7954 0.01688 0.00809 -0.0408 0.4417 1.0000
5.250 0.8166 0.01720 0.00846 -0.0394 0.4217 1.0000
5.500 0.8367 0.01751 0.00879 -0.0378 0.3977 1.0000
5.750 0.8550 0.01779 0.00909 -0.0359 0.3664 1.0000
6.000 0.8708 0.01805 0.00933 -0.0335 0.3222 1.0000
6.250 0.8873 0.01847 0.00961 -0.0315 0.2785 1.0000
6.500 0.9055 0.01907 0.01007 -0.0299 0.2482 1.0000
6.750 0.9239 0.01966 0.01055 -0.0284 0.2232 1.0000
7.000 0.9433 0.02025 0.01113 -0.0272 0.2049 1.0000
7.250 0.9628 0.02082 0.01174 -0.0259 0.1890 1.0000
7.500 0.9817 0.02142 0.01234 -0.0247 0.1699 1.0000
7.750 1.0014 0.02202 0.01301 -0.0235 0.1518 1.0000
8.000 1.0198 0.02275 0.01377 -0.0220 0.1336 1.0000
8.250 1.0370 0.02364 0.01464 -0.0204 0.1066 1.0000
8.500 1.0551 0.02449 0.01556 -0.0188 0.0739 1.0000
8.750 1.0673 0.02578 0.01663 -0.0165 0.0565 1.0000
9.000 1.0779 0.02713 0.01800 -0.0140 0.0518 1.0000
9.250 1.0851 0.02865 0.01956 -0.0112 0.0486 1.0000
9.500 1.0906 0.03021 0.02124 -0.0082 0.0461 1.0000
9.750 1.0957 0.03169 0.02286 -0.0052 0.0445 1.0000
10.000 1.0987 0.03330 0.02459 -0.0020 0.0429 1.0000
10.250 1.1037 0.03500 0.02640 0.0007 0.0422 1.0000
10.500 1.1116 0.03681 0.02830 0.0030 0.0417 1.0000
10.750 1.1229 0.03885 0.03049 0.0048 0.0412 1.0000
11.000 1.1380 0.04112 0.03291 0.0062 0.0409 1.0000
11.250 1.1538 0.04371 0.03573 0.0075 0.0409 1.0000
11.500 1.1643 0.04660 0.03887 0.0090 0.0406 1.0000
11.750 1.1691 0.04956 0.04207 0.0106 0.0403 1.0000
12.000 1.1696 0.05276 0.04551 0.0122 0.0397 1.0000
12.250 1.1639 0.05597 0.04901 0.0139 0.0398 1.0000
12.500 1.1553 0.05959 0.05287 0.0151 0.0395 1.0000
12.750 1.1391 0.06341 0.05704 0.0156 0.0401 1.0000
13.000 1.1206 0.06796 0.06190 0.0151 0.0405 1.0000
13.250 1.0973 0.07361 0.06786 0.0131 0.0413 1.0000
13.500 1.0747 0.07963 0.07411 0.0102 0.0412 1.0000
13.750 1.0336 0.08939 0.08420 0.0036 0.0429 1.0000
14.000 0.9974 0.09990 0.09491 -0.0036 0.0444 1.0000
14.250 0.9625 0.11138 0.10649 -0.0111 0.0458 1.0000
14.500 0.9322 0.12304 0.11819 -0.0183 0.0468 1.0000
14.750 0.9117 0.13345 0.12858 -0.0238 0.0478 1.0000
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Polar data table (+)
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