NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Reynolds number: 100,000 Max Cl/Cd: 47.22 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc12n1-il-100000-n5.txt Download as CSV file: xf-rc12n1-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4566 0.11036 0.10591 -0.0118 1.0000 0.0599
-10.750 -0.4738 0.09871 0.09421 -0.0168 1.0000 0.0362
-10.500 -0.5998 0.10098 0.09625 -0.0100 1.0000 0.0363
-10.250 -0.5954 0.09763 0.09293 -0.0098 1.0000 0.0357
-10.000 -0.5963 0.09264 0.08797 -0.0122 1.0000 0.0350
-9.750 -0.6022 0.08622 0.08157 -0.0171 1.0000 0.0343
-9.500 -0.6140 0.07983 0.07515 -0.0221 1.0000 0.0336
-9.250 -0.6291 0.07436 0.06961 -0.0251 1.0000 0.0331
-9.000 -0.6439 0.06976 0.06491 -0.0258 1.0000 0.0326
-8.750 -0.6543 0.06513 0.06012 -0.0260 1.0000 0.0320
-8.500 -0.6603 0.06047 0.05524 -0.0257 1.0000 0.0315
-8.250 -0.6623 0.05590 0.05039 -0.0250 1.0000 0.0311
-8.000 -0.6600 0.05160 0.04576 -0.0241 1.0000 0.0308
-7.750 -0.6534 0.04775 0.04158 -0.0230 1.0000 0.0307
-7.500 -0.6406 0.04448 0.03798 -0.0226 0.9800 0.0311
-7.250 -0.6186 0.04145 0.03455 -0.0234 0.9521 0.0325
-7.000 -0.5974 0.03845 0.03106 -0.0234 0.9338 0.0336
-6.750 -0.5767 0.03563 0.02774 -0.0228 0.9192 0.0339
-6.500 -0.5556 0.03317 0.02482 -0.0219 0.9069 0.0343
-6.250 -0.5335 0.03103 0.02224 -0.0209 0.8966 0.0348
-6.000 -0.5102 0.02917 0.01999 -0.0199 0.8866 0.0354
-5.750 -0.4859 0.02773 0.01817 -0.0191 0.8778 0.0365
-5.500 -0.4624 0.02618 0.01656 -0.0186 0.8699 0.0386
-5.250 -0.4376 0.02510 0.01535 -0.0182 0.8620 0.0406
-5.000 -0.4122 0.02393 0.01403 -0.0175 0.8550 0.0422
-4.750 -0.3866 0.02284 0.01282 -0.0170 0.8480 0.0440
-4.500 -0.3614 0.02206 0.01185 -0.0163 0.8413 0.0466
-4.250 -0.3387 0.02105 0.01093 -0.0156 0.8350 0.0503
-4.000 -0.3155 0.02032 0.01017 -0.0147 0.8284 0.0539
-3.500 -0.2706 0.01905 0.00878 -0.0128 0.8162 0.0648
-3.250 -0.2476 0.01854 0.00816 -0.0119 0.8106 0.0726
-3.000 -0.2245 0.01800 0.00761 -0.0111 0.8050 0.0869
-2.750 -0.2023 0.01727 0.00707 -0.0102 0.7990 0.1250
-2.500 -0.1979 0.01485 0.00684 -0.0063 0.7940 0.5953
-2.250 -0.1767 0.01463 0.00687 -0.0043 0.7882 0.6946
-2.000 -0.1505 0.01451 0.00694 -0.0030 0.7830 0.7583
-1.750 -0.1196 0.01465 0.00720 -0.0021 0.7788 0.8155
-1.500 -0.0876 0.01493 0.00745 -0.0019 0.7733 0.8560
-1.250 -0.0550 0.01504 0.00745 -0.0026 0.7681 0.8730
-1.000 -0.0240 0.01508 0.00734 -0.0031 0.7638 0.8860
-0.750 0.0085 0.01516 0.00733 -0.0042 0.7579 0.8989
-0.500 0.0437 0.01524 0.00732 -0.0057 0.7527 0.9096
-0.250 0.0777 0.01530 0.00726 -0.0069 0.7485 0.9211
0.000 0.1124 0.01541 0.00731 -0.0086 0.7423 0.9329
0.250 0.1477 0.01548 0.00731 -0.0103 0.7370 0.9442
0.500 0.1867 0.01552 0.00727 -0.0126 0.7326 0.9523
0.750 0.2240 0.01562 0.00735 -0.0150 0.7257 0.9625
1.000 0.2605 0.01566 0.00734 -0.0169 0.7203 0.9727
1.250 0.2997 0.01569 0.00736 -0.0196 0.7143 0.9812
1.500 0.3383 0.01570 0.00737 -0.0222 0.7075 0.9898
1.750 0.3770 0.01565 0.00730 -0.0246 0.7021 0.9978
2.000 0.4054 0.01571 0.00739 -0.0253 0.6937 1.0000
2.250 0.4284 0.01570 0.00734 -0.0245 0.6878 1.0000
2.500 0.4522 0.01579 0.00750 -0.0241 0.6792 1.0000
2.750 0.4756 0.01579 0.00748 -0.0232 0.6729 1.0000
3.000 0.4994 0.01590 0.00766 -0.0228 0.6639 1.0000
3.250 0.5230 0.01588 0.00764 -0.0219 0.6574 1.0000
3.500 0.5467 0.01600 0.00786 -0.0214 0.6474 1.0000
3.750 0.5704 0.01604 0.00794 -0.0205 0.6389 1.0000
4.000 0.5941 0.01607 0.00802 -0.0197 0.6294 1.0000
4.250 0.6177 0.01613 0.00817 -0.0189 0.6180 1.0000
4.500 0.6409 0.01607 0.00818 -0.0179 0.6035 1.0000
4.750 0.6638 0.01595 0.00808 -0.0166 0.5851 1.0000
5.000 0.6865 0.01590 0.00807 -0.0155 0.5632 1.0000
5.250 0.7092 0.01589 0.00809 -0.0143 0.5394 1.0000
5.500 0.7314 0.01596 0.00818 -0.0131 0.5109 1.0000
5.750 0.7530 0.01611 0.00830 -0.0119 0.4759 1.0000
6.000 0.7735 0.01638 0.00845 -0.0105 0.4316 1.0000
6.250 0.7917 0.01690 0.00872 -0.0089 0.3758 1.0000
6.500 0.8074 0.01771 0.00922 -0.0072 0.3138 1.0000
6.750 0.8209 0.01876 0.00993 -0.0054 0.2480 1.0000
7.000 0.8333 0.01993 0.01076 -0.0036 0.1893 1.0000
7.250 0.8456 0.02110 0.01165 -0.0018 0.1468 1.0000
7.500 0.8588 0.02218 0.01261 0.0000 0.1203 1.0000
7.750 0.8716 0.02326 0.01358 0.0017 0.1026 1.0000
8.000 0.8849 0.02429 0.01459 0.0035 0.0902 1.0000
8.250 0.8981 0.02531 0.01563 0.0052 0.0804 1.0000
8.500 0.9095 0.02644 0.01674 0.0071 0.0731 1.0000
8.750 0.9221 0.02749 0.01788 0.0088 0.0668 1.0000
9.000 0.9321 0.02872 0.01910 0.0106 0.0619 1.0000
9.250 0.9445 0.02981 0.02031 0.0123 0.0571 1.0000
9.500 0.9539 0.03102 0.02154 0.0142 0.0538 1.0000
9.750 0.9636 0.03239 0.02293 0.0160 0.0507 1.0000
10.000 0.9759 0.03362 0.02434 0.0175 0.0473 1.0000
10.250 0.9873 0.03496 0.02575 0.0189 0.0446 1.0000
10.500 0.9989 0.03649 0.02728 0.0201 0.0427 1.0000
10.750 1.0135 0.03820 0.02910 0.0213 0.0409 1.0000
11.000 1.0267 0.03986 0.03100 0.0224 0.0388 1.0000
11.250 1.0370 0.04156 0.03288 0.0235 0.0368 1.0000
11.500 1.0459 0.04329 0.03473 0.0245 0.0352 1.0000
11.750 1.0548 0.04521 0.03676 0.0254 0.0340 1.0000
12.000 1.0643 0.04752 0.03916 0.0263 0.0330 1.0000
12.250 1.0689 0.05023 0.04213 0.0273 0.0323 1.0000
12.500 1.0682 0.05316 0.04539 0.0282 0.0316 1.0000
12.750 1.0637 0.05638 0.04892 0.0288 0.0309 1.0000
13.000 1.0563 0.05988 0.05271 0.0291 0.0303 1.0000
13.250 1.0466 0.06368 0.05678 0.0290 0.0296 1.0000
13.500 1.0347 0.06785 0.06119 0.0284 0.0291 1.0000
13.750 1.0205 0.07249 0.06606 0.0272 0.0287 1.0000
14.000 1.0033 0.07781 0.07161 0.0254 0.0285 1.0000
14.250 0.9817 0.08417 0.07820 0.0226 0.0285 1.0000
14.500 0.9534 0.09224 0.08651 0.0183 0.0287 1.0000
14.750 0.9161 0.10330 0.09781 0.0117 0.0294 1.0000
15.000 0.8709 0.11833 0.11300 0.0022 0.0306 1.0000
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Polar data table (+)
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