NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC12-64C AIRFOIL (rc1264c-il) Reynolds number: 50,000 Max Cl/Cd: 30.04 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc1264c-il-50000-n5.txt Download as CSV file: xf-rc1264c-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC12-64C 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5014 0.10378 0.09604 -0.0347 1.0000 0.0567
-9.750 -0.5216 0.09469 0.08699 -0.0423 1.0000 0.0530
-9.500 -0.5221 0.09072 0.08306 -0.0434 1.0000 0.0520
-9.250 -0.5313 0.08596 0.07833 -0.0453 1.0000 0.0510
-9.000 -0.5474 0.08132 0.07369 -0.0464 1.0000 0.0501
-8.750 -0.5696 0.07733 0.06968 -0.0454 1.0000 0.0492
-8.500 -0.5958 0.07352 0.06578 -0.0427 1.0000 0.0483
-8.250 -0.6218 0.06957 0.06158 -0.0393 1.0000 0.0470
-8.000 -0.6260 0.06684 0.05880 -0.0367 1.0000 0.0465
-7.750 -0.6312 0.06404 0.05589 -0.0338 1.0000 0.0459
-7.500 -0.6361 0.06119 0.05289 -0.0307 1.0000 0.0453
-7.250 -0.6401 0.05829 0.04977 -0.0274 1.0000 0.0446
-7.000 -0.6427 0.05535 0.04655 -0.0240 1.0000 0.0437
-6.750 -0.6438 0.05230 0.04313 -0.0203 1.0000 0.0428
-6.500 -0.6427 0.04937 0.03963 -0.0166 1.0000 0.0417
-6.250 -0.6341 0.04713 0.03719 -0.0140 1.0000 0.0413
-6.000 -0.6242 0.04494 0.03476 -0.0114 1.0000 0.0409
-5.750 -0.6126 0.04283 0.03237 -0.0090 1.0000 0.0405
-5.500 -0.5991 0.04082 0.03004 -0.0067 1.0000 0.0400
-5.250 -0.5837 0.03893 0.02785 -0.0047 1.0000 0.0396
-5.000 -0.5664 0.03720 0.02584 -0.0028 1.0000 0.0392
-4.750 -0.5477 0.03565 0.02403 -0.0012 1.0000 0.0391
-4.500 -0.5280 0.03427 0.02245 0.0003 1.0000 0.0393
-4.250 -0.5075 0.03305 0.02105 0.0017 1.0000 0.0397
-4.000 -0.4862 0.03195 0.01978 0.0030 1.0000 0.0402
-3.750 -0.4645 0.03099 0.01868 0.0042 1.0000 0.0407
-3.500 -0.4428 0.03015 0.01773 0.0054 1.0000 0.0411
-3.250 -0.4216 0.02941 0.01688 0.0067 1.0000 0.0413
-3.000 -0.4013 0.02877 0.01613 0.0080 1.0000 0.0415
-2.750 -0.3816 0.02821 0.01543 0.0094 1.0000 0.0418
-2.500 -0.3622 0.02769 0.01484 0.0106 1.0000 0.0422
-2.250 -0.3425 0.02724 0.01429 0.0118 0.9998 0.0427
-2.000 -0.3065 0.02683 0.01377 0.0097 0.9943 0.0438
-1.750 -0.2715 0.02649 0.01331 0.0079 0.9880 0.0452
-1.250 -0.2017 0.02598 0.01268 0.0043 0.9752 0.0524
-1.000 -0.1720 0.02520 0.01242 0.0031 0.9679 0.1192
-0.750 -0.1352 0.02306 0.01372 0.0039 0.9680 0.8632
-0.500 -0.0773 0.02381 0.01428 -0.0010 0.9648 0.9321
-0.250 -0.0145 0.02432 0.01455 -0.0079 0.9616 0.9593
0.000 0.0556 0.02470 0.01476 -0.0165 0.9584 0.9860
0.250 0.1161 0.02487 0.01479 -0.0237 0.9532 1.0000
0.500 0.1501 0.02498 0.01481 -0.0256 0.9434 1.0000
0.750 0.1862 0.02508 0.01484 -0.0279 0.9339 1.0000
1.000 0.2152 0.02519 0.01490 -0.0288 0.9222 1.0000
1.250 0.2504 0.02529 0.01497 -0.0308 0.9125 1.0000
1.500 0.2852 0.02536 0.01503 -0.0326 0.9019 1.0000
1.750 0.3144 0.02546 0.01513 -0.0332 0.8895 1.0000
2.000 0.3485 0.02549 0.01518 -0.0347 0.8782 1.0000
2.250 0.3875 0.02540 0.01513 -0.0368 0.8677 1.0000
2.500 0.4157 0.02537 0.01515 -0.0368 0.8530 1.0000
2.750 0.4459 0.02523 0.01508 -0.0369 0.8376 1.0000
3.000 0.4774 0.02499 0.01490 -0.0370 0.8211 1.0000
3.250 0.5110 0.02457 0.01457 -0.0371 0.8030 1.0000
3.500 0.5379 0.02420 0.01428 -0.0359 0.7812 1.0000
3.750 0.5640 0.02381 0.01397 -0.0344 0.7575 1.0000
4.000 0.5895 0.02332 0.01353 -0.0325 0.7287 1.0000
4.250 0.6128 0.02284 0.01306 -0.0301 0.6919 1.0000
4.500 0.6340 0.02247 0.01264 -0.0273 0.6464 1.0000
4.750 0.6534 0.02232 0.01237 -0.0245 0.5946 1.0000
5.000 0.6706 0.02239 0.01223 -0.0215 0.5296 1.0000
5.250 0.6844 0.02278 0.01221 -0.0182 0.4486 1.0000
5.500 0.6938 0.02355 0.01249 -0.0146 0.3685 1.0000
5.750 0.7005 0.02458 0.01305 -0.0110 0.3012 1.0000
6.000 0.7073 0.02572 0.01381 -0.0076 0.2494 1.0000
6.250 0.7154 0.02689 0.01469 -0.0046 0.2127 1.0000
6.500 0.7258 0.02798 0.01559 -0.0019 0.1866 1.0000
6.750 0.7381 0.02905 0.01654 0.0005 0.1691 1.0000
7.000 0.7530 0.03011 0.01753 0.0026 0.1557 1.0000
7.250 0.7695 0.03116 0.01855 0.0043 0.1443 1.0000
7.500 0.7877 0.03220 0.01954 0.0057 0.1342 1.0000
7.750 0.8095 0.03327 0.02071 0.0068 0.1254 1.0000
8.000 0.8296 0.03441 0.02177 0.0078 0.1179 1.0000
8.250 0.8521 0.03560 0.02316 0.0087 0.1105 1.0000
8.500 0.8736 0.03691 0.02434 0.0094 0.1047 1.0000
8.750 0.8933 0.03829 0.02604 0.0106 0.0982 1.0000
9.000 0.9142 0.03979 0.02760 0.0114 0.0937 1.0000
9.250 0.9345 0.04158 0.02949 0.0123 0.0895 1.0000
9.500 0.9483 0.04335 0.03159 0.0140 0.0847 1.0000
9.750 0.9626 0.04499 0.03335 0.0155 0.0809 1.0000
10.000 0.9810 0.04710 0.03546 0.0163 0.0783 1.0000
10.250 0.9867 0.04962 0.03844 0.0188 0.0761 1.0000
10.500 0.9887 0.05213 0.04136 0.0215 0.0737 1.0000
10.750 0.9903 0.05443 0.04392 0.0241 0.0712 1.0000
11.000 0.9932 0.05649 0.04614 0.0263 0.0690 1.0000
11.250 0.9984 0.05852 0.04821 0.0281 0.0672 1.0000
11.500 0.9943 0.06125 0.05111 0.0307 0.0662 1.0000
11.750 0.9766 0.06455 0.05474 0.0340 0.0658 1.0000
12.000 0.9562 0.06827 0.05876 0.0364 0.0655 1.0000
12.250 0.9329 0.07256 0.06332 0.0378 0.0653 1.0000
12.500 0.9066 0.07762 0.06861 0.0377 0.0653 1.0000
12.750 0.8775 0.08371 0.07490 0.0362 0.0654 1.0000
13.000 0.8460 0.09116 0.08251 0.0328 0.0657 1.0000
13.250 0.8134 0.10024 0.09168 0.0276 0.0660 1.0000
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Polar data table (+)
Polar graphs
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