NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 200,000 Max Cl/Cd: 65.48 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10n1-il-200000.txt Download as CSV file: xf-rc10n1-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(N)1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5864 0.09459 0.09143 -0.0071 1.0000 0.0368
-8.750 -0.5946 0.08982 0.08661 -0.0119 1.0000 0.0369
-8.500 -0.6000 0.08594 0.08264 -0.0141 1.0000 0.0370
-8.250 -0.6018 0.08209 0.07866 -0.0158 1.0000 0.0371
-7.750 -0.6017 0.06934 0.06589 -0.0180 1.0000 0.0384
-7.500 -0.5922 0.06550 0.06208 -0.0183 1.0000 0.0393
-7.250 -0.5825 0.06198 0.05853 -0.0188 1.0000 0.0404
-7.000 -0.5717 0.05840 0.05485 -0.0196 1.0000 0.0419
-6.750 -0.5588 0.05475 0.05105 -0.0204 1.0000 0.0440
-6.500 -0.5398 0.05348 0.04920 -0.0205 1.0000 0.0490
-6.250 -0.5304 0.04850 0.04382 -0.0209 0.9923 0.0503
-6.000 -0.5044 0.04310 0.03855 -0.0239 0.9748 0.0526
-5.750 -0.4772 0.04014 0.03544 -0.0257 0.9594 0.0565
-5.500 -0.4271 0.02172 0.01690 -0.0276 0.9181 0.0661
-5.250 -0.4107 0.01957 0.01467 -0.0265 0.9083 0.0706
-5.000 -0.4156 0.03336 0.02756 -0.0238 0.9193 0.0793
-3.750 -0.2913 0.02338 0.01571 -0.0159 0.8762 0.0497
-3.500 -0.2664 0.02048 0.01265 -0.0150 0.8699 0.0476
-3.250 -0.2390 0.01878 0.01079 -0.0144 0.8626 0.0460
-3.000 -0.2133 0.01759 0.00944 -0.0135 0.8563 0.0457
-2.750 -0.1868 0.01662 0.00839 -0.0129 0.8493 0.0466
-2.500 -0.1618 0.01590 0.00760 -0.0120 0.8431 0.0484
-2.250 -0.1384 0.01494 0.00666 -0.0112 0.8366 0.0531
-2.000 -0.1144 0.01433 0.00604 -0.0103 0.8301 0.0580
-1.750 -0.0904 0.01379 0.00542 -0.0094 0.8244 0.0658
-1.500 -0.0661 0.01317 0.00488 -0.0087 0.8176 0.0984
-1.250 -0.0575 0.01017 0.00505 -0.0034 0.8130 0.8870
-1.000 0.0033 0.01072 0.00548 -0.0085 0.8080 0.9599
-0.750 0.0800 0.01108 0.00565 -0.0178 0.8032 0.9904
-0.500 0.1266 0.01104 0.00543 -0.0220 0.7982 1.0000
-0.250 0.1520 0.01096 0.00529 -0.0223 0.7904 1.0000
0.000 0.1760 0.01094 0.00514 -0.0218 0.7844 1.0000
0.250 0.2022 0.01094 0.00509 -0.0219 0.7767 1.0000
0.500 0.2268 0.01094 0.00500 -0.0214 0.7704 1.0000
0.750 0.2528 0.01097 0.00500 -0.0214 0.7625 1.0000
1.000 0.2773 0.01099 0.00494 -0.0208 0.7562 1.0000
1.250 0.3032 0.01104 0.00498 -0.0206 0.7481 1.0000
1.500 0.3278 0.01106 0.00494 -0.0199 0.7417 1.0000
1.750 0.3536 0.01111 0.00500 -0.0197 0.7331 1.0000
2.000 0.3782 0.01113 0.00496 -0.0190 0.7268 1.0000
2.250 0.4041 0.01118 0.00506 -0.0188 0.7179 1.0000
2.500 0.4288 0.01120 0.00504 -0.0180 0.7113 1.0000
2.750 0.4544 0.01124 0.00511 -0.0177 0.7019 1.0000
3.000 0.4794 0.01124 0.00513 -0.0171 0.6929 1.0000
3.250 0.5036 0.01112 0.00500 -0.0160 0.6808 1.0000
3.500 0.5279 0.01097 0.00484 -0.0150 0.6662 1.0000
3.750 0.5526 0.01088 0.00475 -0.0141 0.6528 1.0000
4.000 0.5775 0.01082 0.00471 -0.0134 0.6387 1.0000
4.250 0.6025 0.01077 0.00472 -0.0127 0.6226 1.0000
4.500 0.6274 0.01074 0.00472 -0.0119 0.6045 1.0000
4.750 0.6521 0.01073 0.00471 -0.0111 0.5843 1.0000
5.000 0.6767 0.01076 0.00476 -0.0104 0.5579 1.0000
5.250 0.7008 0.01085 0.00484 -0.0096 0.5250 1.0000
5.500 0.7242 0.01106 0.00495 -0.0087 0.4761 1.0000
5.750 0.7455 0.01157 0.00516 -0.0076 0.4023 1.0000
6.000 0.7642 0.01253 0.00566 -0.0066 0.3065 1.0000
6.250 0.7801 0.01393 0.00648 -0.0054 0.1988 1.0000
6.500 0.7967 0.01523 0.00738 -0.0043 0.1391 1.0000
6.750 0.8147 0.01631 0.00830 -0.0031 0.1108 1.0000
7.000 0.8327 0.01733 0.00922 -0.0019 0.0926 1.0000
7.250 0.8510 0.01830 0.01017 -0.0007 0.0798 1.0000
7.500 0.8693 0.01930 0.01122 0.0006 0.0703 1.0000
7.750 0.8864 0.02051 0.01239 0.0019 0.0627 1.0000
8.000 0.9059 0.02144 0.01340 0.0031 0.0568 1.0000
8.250 0.9227 0.02309 0.01500 0.0044 0.0517 1.0000
8.500 0.9431 0.02394 0.01601 0.0055 0.0472 1.0000
8.750 0.9624 0.02517 0.01728 0.0065 0.0437 1.0000
9.000 0.9815 0.02770 0.01990 0.0075 0.0406 1.0000
9.250 1.0009 0.02876 0.02121 0.0086 0.0379 1.0000
9.500 1.0196 0.03029 0.02293 0.0096 0.0354 1.0000
9.750 1.0372 0.03207 0.02485 0.0106 0.0336 1.0000
10.000 1.0526 0.03511 0.02804 0.0116 0.0318 1.0000
10.250 1.0611 0.03825 0.03159 0.0133 0.0305 1.0000
10.500 1.0695 0.04036 0.03409 0.0151 0.0292 1.0000
10.750 1.0720 0.04361 0.03776 0.0171 0.0285 1.0000
11.000 1.0686 0.04717 0.04170 0.0192 0.0282 1.0000
11.250 1.0580 0.05077 0.04562 0.0217 0.0281 1.0000
11.500 1.0421 0.05451 0.04964 0.0236 0.0282 1.0000
11.750 1.0236 0.05872 0.05411 0.0244 0.0283 1.0000
12.000 1.0031 0.06346 0.05908 0.0240 0.0285 1.0000
12.250 0.9815 0.06876 0.06458 0.0225 0.0288 1.0000
12.500 0.9585 0.07477 0.07074 0.0199 0.0291 1.0000
12.750 0.9346 0.08158 0.07771 0.0162 0.0294 1.0000
13.000 0.9099 0.08942 0.08568 0.0114 0.0298 1.0000
13.250 0.8847 0.09842 0.09476 0.0056 0.0303 1.0000
13.500 0.8627 0.10768 0.10405 0.0008 0.0310 1.0000
13.750 0.6591 0.12088 0.11746 0.0037 0.0425 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il)