AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 500,000 Max Cl/Cd: 75.32 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg12a-il-500000-n5.txt Download as CSV file: xf-rg12a-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5105 0.08491 0.08262 -0.0266 1.0000 0.0047
-9.250 -0.5182 0.07969 0.07744 -0.0290 1.0000 0.0047
-9.000 -0.5313 0.07370 0.07151 -0.0322 1.0000 0.0047
-8.750 -0.5471 0.06812 0.06598 -0.0360 1.0000 0.0048
-8.500 -0.5720 0.06180 0.05963 -0.0392 1.0000 0.0047
-8.250 -0.5893 0.05518 0.05291 -0.0405 1.0000 0.0046
-8.000 -0.6171 0.04476 0.04219 -0.0400 1.0000 0.0044
-7.750 -0.6350 0.03333 0.03014 -0.0387 0.9992 0.0045
-7.500 -0.6168 0.02629 0.02230 -0.0407 0.9959 0.0046
-7.250 -0.5899 0.02354 0.01915 -0.0420 0.9937 0.0048
-7.000 -0.5624 0.02187 0.01721 -0.0429 0.9910 0.0053
-6.750 -0.5341 0.01981 0.01479 -0.0437 0.9883 0.0055
-6.500 -0.5040 0.01803 0.01268 -0.0447 0.9860 0.0058
-6.250 -0.4728 0.01645 0.01084 -0.0457 0.9842 0.0059
-6.000 -0.4441 0.01523 0.00942 -0.0462 0.9813 0.0060
-5.750 -0.4161 0.01423 0.00825 -0.0464 0.9775 0.0061
-5.500 -0.3855 0.01340 0.00730 -0.0471 0.9744 0.0064
-5.250 -0.3543 0.01233 0.00613 -0.0481 0.9720 0.0068
-5.000 -0.3246 0.01174 0.00548 -0.0487 0.9683 0.0076
-4.750 -0.2961 0.01123 0.00492 -0.0489 0.9632 0.0083
-4.500 -0.2637 0.01078 0.00439 -0.0500 0.9596 0.0093
-4.250 -0.2323 0.01031 0.00388 -0.0508 0.9554 0.0113
-4.000 -0.2033 0.00993 0.00346 -0.0510 0.9487 0.0139
-3.750 -0.1687 0.00952 0.00303 -0.0525 0.9446 0.0201
-3.500 -0.1386 0.00913 0.00272 -0.0531 0.9370 0.0401
-3.250 -0.1042 0.00873 0.00245 -0.0546 0.9309 0.0707
-3.000 -0.0722 0.00839 0.00222 -0.0556 0.9216 0.1058
-2.750 -0.0390 0.00797 0.00199 -0.0570 0.9113 0.1631
-2.500 -0.0071 0.00755 0.00178 -0.0581 0.8985 0.2358
-2.250 0.0231 0.00717 0.00160 -0.0587 0.8825 0.3115
-2.000 0.0519 0.00686 0.00145 -0.0591 0.8639 0.3845
-1.750 0.0790 0.00652 0.00131 -0.0590 0.8439 0.4733
-1.500 0.1029 0.00599 0.00125 -0.0583 0.8223 0.6290
-1.250 0.1280 0.00591 0.00129 -0.0575 0.8002 0.7076
-1.000 0.1542 0.00596 0.00128 -0.0570 0.7768 0.7339
-0.750 0.1802 0.00603 0.00128 -0.0564 0.7530 0.7538
-0.500 0.2061 0.00612 0.00128 -0.0558 0.7296 0.7691
-0.250 0.2320 0.00622 0.00129 -0.0553 0.7068 0.7822
0.000 0.2580 0.00632 0.00131 -0.0547 0.6844 0.7934
0.250 0.2839 0.00643 0.00133 -0.0542 0.6626 0.8044
0.500 0.3099 0.00653 0.00137 -0.0537 0.6409 0.8152
1.000 0.3609 0.00674 0.00147 -0.0525 0.5981 0.8374
1.250 0.3865 0.00686 0.00153 -0.0519 0.5764 0.8473
1.500 0.4123 0.00698 0.00159 -0.0514 0.5537 0.8543
1.750 0.4385 0.00712 0.00165 -0.0511 0.5308 0.8599
2.000 0.4641 0.00726 0.00172 -0.0506 0.5074 0.8646
2.250 0.4901 0.00741 0.00181 -0.0502 0.4832 0.8699
2.500 0.5158 0.00757 0.00190 -0.0498 0.4583 0.8752
2.750 0.5412 0.00775 0.00200 -0.0493 0.4334 0.8808
3.000 0.5666 0.00795 0.00212 -0.0489 0.4058 0.8869
3.250 0.5915 0.00814 0.00226 -0.0483 0.3802 0.8934
3.500 0.6166 0.00835 0.00241 -0.0478 0.3540 0.9010
3.750 0.6407 0.00858 0.00256 -0.0471 0.3257 0.9090
4.000 0.6645 0.00885 0.00274 -0.0464 0.2947 0.9182
4.500 0.7118 0.00945 0.00316 -0.0448 0.2293 0.9438
4.750 0.7378 0.00991 0.00342 -0.0448 0.1839 0.9616
5.000 0.7684 0.01035 0.00372 -0.0457 0.1472 0.9857
5.250 0.7941 0.01082 0.00405 -0.0457 0.1164 1.0000
5.500 0.8181 0.01130 0.00440 -0.0452 0.0887 1.0000
5.750 0.8426 0.01173 0.00475 -0.0448 0.0696 1.0000
6.000 0.8666 0.01220 0.00514 -0.0444 0.0520 1.0000
6.250 0.8904 0.01268 0.00556 -0.0438 0.0361 1.0000
6.500 0.9136 0.01325 0.00607 -0.0432 0.0225 1.0000
6.750 0.9374 0.01372 0.00655 -0.0426 0.0182 1.0000
7.000 0.9607 0.01426 0.00712 -0.0420 0.0153 1.0000
7.250 0.9845 0.01472 0.00765 -0.0414 0.0143 1.0000
7.500 1.0078 0.01521 0.00822 -0.0408 0.0136 1.0000
7.750 1.0306 0.01576 0.00887 -0.0400 0.0130 1.0000
8.000 1.0528 0.01634 0.00954 -0.0392 0.0124 1.0000
8.250 1.0743 0.01699 0.01026 -0.0384 0.0119 1.0000
8.500 1.0949 0.01770 0.01105 -0.0374 0.0112 1.0000
8.750 1.1131 0.01865 0.01209 -0.0360 0.0104 1.0000
9.000 1.1295 0.01975 0.01331 -0.0344 0.0098 1.0000
9.250 1.1499 0.02039 0.01404 -0.0335 0.0095 1.0000
9.500 1.1691 0.02112 0.01487 -0.0324 0.0090 1.0000
9.750 1.1884 0.02178 0.01565 -0.0313 0.0084 1.0000
10.000 1.2054 0.02265 0.01662 -0.0299 0.0079 1.0000
10.250 1.2213 0.02355 0.01762 -0.0285 0.0075 1.0000
10.500 1.2366 0.02443 0.01859 -0.0269 0.0071 1.0000
10.750 1.2488 0.02543 0.01969 -0.0250 0.0068 1.0000
11.000 1.2524 0.02694 0.02131 -0.0219 0.0064 1.0000
11.250 1.2583 0.02826 0.02278 -0.0192 0.0061 1.0000
11.500 1.2704 0.02908 0.02372 -0.0174 0.0058 1.0000
11.750 1.2776 0.03030 0.02508 -0.0153 0.0056 1.0000
12.000 1.2850 0.03150 0.02644 -0.0133 0.0052 1.0000
12.250 1.2929 0.03267 0.02772 -0.0117 0.0049 1.0000
12.500 1.2975 0.03417 0.02935 -0.0101 0.0046 1.0000
12.750 1.3012 0.03584 0.03113 -0.0087 0.0044 1.0000
13.000 1.3034 0.03774 0.03315 -0.0076 0.0043 1.0000
13.250 1.3068 0.03965 0.03516 -0.0069 0.0041 1.0000
13.500 1.3031 0.04250 0.03816 -0.0065 0.0040 1.0000
13.750 1.2973 0.04584 0.04165 -0.0066 0.0039 1.0000
14.000 1.2884 0.04988 0.04584 -0.0075 0.0038 1.0000
14.250 1.2761 0.05480 0.05092 -0.0092 0.0037 1.0000
14.500 1.2678 0.05954 0.05584 -0.0112 0.0037 1.0000
14.750 1.2519 0.06581 0.06229 -0.0143 0.0037 1.0000
15.000 1.2378 0.07220 0.06886 -0.0177 0.0036 1.0000
15.250 1.2248 0.07879 0.07562 -0.0214 0.0036 1.0000
15.500 1.2039 0.08717 0.08415 -0.0262 0.0036 1.0000
15.750 1.1813 0.09631 0.09345 -0.0316 0.0037 1.0000
16.000 1.1631 0.10497 0.10226 -0.0366 0.0036 1.0000
16.250 1.1396 0.11507 0.11249 -0.0425 0.0037 1.0000
16.500 1.1177 0.12525 0.12280 -0.0483 0.0037 1.0000
17.000 1.0769 0.14600 0.14378 -0.0602 0.0038 1.0000
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Polar data table (+)
Polar graphs
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