NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Reynolds number: 50,000 Max Cl/Cd: 31.99 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc12n1-il-50000.txt Download as CSV file: xf-rc12n1-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-12(N)1 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.5530 0.13546 0.12879 0.0177 1.0000 0.2394
-10.750 -0.5735 0.13434 0.12777 0.0150 1.0000 0.2494
-10.500 -0.5643 0.13046 0.12392 0.0154 1.0000 0.2643
-10.250 -0.5569 0.12676 0.12024 0.0155 1.0000 0.2793
-10.000 -0.5515 0.12326 0.11678 0.0156 1.0000 0.2944
-9.750 -0.5521 0.12026 0.11384 0.0155 1.0000 0.3099
-9.500 -0.5210 0.11489 0.10842 0.0172 1.0000 0.3293
-9.250 -0.5112 0.11162 0.10517 0.0179 1.0000 0.3501
-9.000 -0.5256 0.10990 0.10357 0.0180 1.0000 0.3704
-8.750 -0.5002 0.10582 0.09945 0.0197 1.0000 0.3983
-8.500 -0.4827 0.10230 0.09595 0.0210 1.0000 0.4268
-8.250 -0.4662 0.09875 0.09242 0.0220 1.0000 0.4533
-7.500 -0.4403 0.08943 0.08320 0.0241 1.0000 0.5219
-6.500 -0.6044 0.05348 0.04649 -0.0199 1.0000 0.1839
-6.250 -0.6070 0.05039 0.04298 -0.0169 1.0000 0.1675
-6.000 -0.6078 0.04791 0.03998 -0.0135 1.0000 0.1559
-5.750 -0.6033 0.04522 0.03695 -0.0106 1.0000 0.1494
-5.500 -0.5954 0.04316 0.03427 -0.0077 1.0000 0.1422
-5.250 -0.5831 0.04060 0.03144 -0.0056 1.0000 0.1380
-5.000 -0.5689 0.03837 0.02867 -0.0034 1.0000 0.1329
-4.750 -0.5525 0.03706 0.02666 -0.0011 1.0000 0.1294
-4.500 -0.5335 0.03504 0.02439 0.0003 1.0000 0.1300
-4.250 -0.5129 0.03296 0.02227 0.0011 1.0000 0.1336
-4.000 -0.4908 0.03145 0.02055 0.0020 1.0000 0.1365
-3.750 -0.4662 0.03009 0.01898 0.0028 1.0000 0.1391
-3.500 -0.4408 0.02900 0.01768 0.0036 1.0000 0.1450
-3.250 -0.0478 0.02506 0.01638 -0.0362 1.0000 1.0000
-3.000 -0.0555 0.02516 0.01640 -0.0328 1.0000 1.0000
-2.750 -0.0645 0.02523 0.01639 -0.0292 1.0000 1.0000
-2.500 -0.0738 0.02524 0.01632 -0.0255 1.0000 1.0000
-2.250 -0.0836 0.02519 0.01619 -0.0218 1.0000 1.0000
-2.000 -0.0938 0.02506 0.01599 -0.0180 1.0000 1.0000
-1.750 -0.1048 0.02487 0.01572 -0.0139 1.0000 1.0000
-1.500 -0.1161 0.02462 0.01538 -0.0098 1.0000 1.0000
-1.250 -0.1254 0.02439 0.01505 -0.0057 1.0000 1.0000
-1.000 -0.1297 0.02427 0.01478 -0.0023 1.0000 1.0000
-0.750 -0.1286 0.02428 0.01462 0.0006 1.0000 1.0000
-0.500 -0.1233 0.02440 0.01457 0.0028 1.0000 1.0000
-0.250 -0.0881 0.02494 0.01490 -0.0005 0.9923 1.0000
0.000 -0.0534 0.02556 0.01532 -0.0035 0.9845 1.0000
0.250 -0.0232 0.02614 0.01574 -0.0057 0.9762 1.0000
0.500 0.0093 0.02685 0.01630 -0.0082 0.9682 1.0000
0.750 0.0420 0.02758 0.01691 -0.0107 0.9596 1.0000
1.000 0.0691 0.02829 0.01753 -0.0121 0.9508 1.0000
1.250 0.1066 0.02919 0.01835 -0.0153 0.9419 1.0000
1.500 0.1304 0.02992 0.01902 -0.0160 0.9322 1.0000
1.750 0.1586 0.03078 0.01984 -0.0175 0.9224 1.0000
2.000 0.1958 0.03177 0.02082 -0.0204 0.9121 1.0000
2.250 0.2195 0.03262 0.02166 -0.0210 0.9010 1.0000
2.500 0.2433 0.03355 0.02259 -0.0216 0.8896 1.0000
2.750 0.2719 0.03455 0.02361 -0.0229 0.8776 1.0000
3.000 0.3033 0.03558 0.02470 -0.0246 0.8650 1.0000
3.250 0.3362 0.03664 0.02582 -0.0264 0.8518 1.0000
3.500 0.3689 0.03769 0.02695 -0.0281 0.8378 1.0000
3.750 0.3968 0.03872 0.02809 -0.0289 0.8231 1.0000
4.000 0.4235 0.03976 0.02922 -0.0294 0.8076 1.0000
4.250 0.4448 0.04082 0.03037 -0.0290 0.7914 1.0000
4.500 0.4652 0.04189 0.03154 -0.0285 0.7743 1.0000
4.750 0.4882 0.04292 0.03271 -0.0282 0.7563 1.0000
5.000 0.5184 0.04380 0.03376 -0.0286 0.7376 1.0000
5.250 0.5680 0.04417 0.03439 -0.0308 0.7181 1.0000
5.500 0.5776 0.04516 0.03551 -0.0283 0.6965 1.0000
5.750 0.6312 0.04440 0.03507 -0.0291 0.6734 1.0000
6.000 0.6559 0.04419 0.03505 -0.0268 0.6476 1.0000
6.250 0.7007 0.04206 0.03326 -0.0243 0.6207 1.0000
6.500 0.7524 0.03816 0.02970 -0.0205 0.5931 1.0000
6.750 0.7965 0.03399 0.02585 -0.0158 0.5608 1.0000
7.000 0.8350 0.02978 0.02184 -0.0105 0.5164 1.0000
7.250 0.8588 0.02685 0.01863 -0.0043 0.4335 1.0000
7.500 0.8625 0.02775 0.01849 0.0013 0.3269 1.0000
7.750 0.8692 0.02989 0.01992 0.0047 0.2623 1.0000
8.000 0.8841 0.03194 0.02159 0.0068 0.2216 1.0000
8.250 0.9056 0.03401 0.02344 0.0082 0.1934 1.0000
8.500 0.9311 0.03620 0.02546 0.0089 0.1730 1.0000
8.750 0.9560 0.03861 0.02783 0.0096 0.1575 1.0000
9.000 0.9751 0.04118 0.03074 0.0109 0.1473 1.0000
9.250 0.9941 0.04415 0.03390 0.0119 0.1395 1.0000
9.500 1.0099 0.04686 0.03686 0.0132 0.1319 1.0000
9.750 1.0262 0.05058 0.04075 0.0141 0.1278 1.0000
10.000 1.0274 0.05411 0.04487 0.0163 0.1258 1.0000
10.250 1.0245 0.05782 0.04905 0.0185 0.1239 1.0000
10.500 1.0177 0.06164 0.05326 0.0205 0.1223 1.0000
10.750 1.0056 0.06569 0.05763 0.0224 0.1216 1.0000
11.000 0.9858 0.07002 0.06223 0.0241 0.1221 1.0000
11.250 0.9572 0.07459 0.06700 0.0254 0.1236 1.0000
11.500 0.9288 0.08003 0.07257 0.0249 0.1254 1.0000
11.750 0.9043 0.08616 0.07878 0.0233 0.1269 1.0000
12.000 0.8864 0.09261 0.08528 0.0213 0.1281 1.0000
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Polar data table (+)
Polar graphs
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