NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 200,000 Max Cl/Cd: 60.81 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc510-il-200000.txt Download as CSV file: xf-rc510-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(5)-10 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5224 0.08730 0.08422 -0.0076 1.0000 0.0515
-8.000 -0.5325 0.08115 0.07793 -0.0195 1.0000 0.0526
-7.750 -0.5316 0.07690 0.07346 -0.0236 1.0000 0.0529
-7.500 -0.5249 0.06996 0.06664 -0.0246 1.0000 0.0539
-7.250 -0.5103 0.06660 0.06335 -0.0250 1.0000 0.0548
-7.000 -0.4950 0.06370 0.06047 -0.0260 1.0000 0.0564
-6.750 -0.4740 0.06002 0.05668 -0.0294 0.9729 0.0596
-6.500 -0.4563 0.05516 0.05104 -0.0342 0.9470 0.0651
-6.250 -0.4403 0.05073 0.04668 -0.0345 0.9330 0.0663
-6.000 -0.4245 0.04810 0.04401 -0.0340 0.9203 0.0677
-5.750 -0.4091 0.04578 0.04156 -0.0331 0.9091 0.0698
-5.500 -0.3914 0.04347 0.03899 -0.0323 0.8984 0.0742
-5.250 -0.3755 0.04027 0.03526 -0.0313 0.8889 0.0794
-5.000 -0.3573 0.03778 0.03272 -0.0305 0.8811 0.0811
-4.750 -0.3360 0.02914 0.02304 -0.0277 0.8740 0.0500
-4.500 -0.3146 0.02633 0.01981 -0.0262 0.8672 0.0492
-4.250 -0.2909 0.02340 0.01640 -0.0250 0.8605 0.0476
-4.000 -0.2659 0.02111 0.01368 -0.0238 0.8538 0.0471
-3.750 -0.2402 0.01969 0.01199 -0.0229 0.8479 0.0481
-3.500 -0.2131 0.01853 0.01064 -0.0224 0.8414 0.0499
-3.250 -0.1867 0.01798 0.00987 -0.0217 0.8360 0.0529
-3.000 -0.1607 0.01681 0.00876 -0.0214 0.8302 0.0576
-2.750 -0.1343 0.01620 0.00807 -0.0208 0.8242 0.0630
-2.500 -0.1097 0.01552 0.00740 -0.0199 0.8195 0.0711
-2.250 -0.0829 0.01506 0.00698 -0.0197 0.8134 0.0841
-2.000 -0.0570 0.01470 0.00666 -0.0192 0.8083 0.1011
-1.750 -0.0316 0.01446 0.00638 -0.0185 0.8040 0.1212
-1.500 -0.0056 0.01402 0.00609 -0.0184 0.7978 0.1437
-1.250 0.0196 0.01365 0.00580 -0.0179 0.7930 0.1708
-1.000 0.0429 0.01310 0.00551 -0.0170 0.7892 0.2308
-0.750 0.0854 0.01101 0.00594 -0.0184 0.7841 0.9467
-0.500 0.1633 0.01118 0.00584 -0.0274 0.7754 0.9928
-0.250 0.2046 0.01110 0.00563 -0.0302 0.7654 1.0000
0.000 0.2284 0.01107 0.00544 -0.0292 0.7585 1.0000
0.250 0.2542 0.01104 0.00535 -0.0289 0.7493 1.0000
0.500 0.2784 0.01101 0.00520 -0.0279 0.7423 1.0000
0.750 0.3038 0.01098 0.00511 -0.0275 0.7330 1.0000
1.000 0.3285 0.01094 0.00498 -0.0266 0.7247 1.0000
1.250 0.3534 0.01089 0.00487 -0.0259 0.7157 1.0000
1.500 0.3788 0.01087 0.00481 -0.0253 0.7075 1.0000
1.750 0.4039 0.01083 0.00471 -0.0245 0.6993 1.0000
2.000 0.4293 0.01081 0.00467 -0.0239 0.6901 1.0000
2.250 0.4542 0.01075 0.00452 -0.0230 0.6816 1.0000
2.500 0.4797 0.01070 0.00448 -0.0224 0.6700 1.0000
2.750 0.5050 0.01065 0.00443 -0.0217 0.6585 1.0000
3.000 0.5301 0.01058 0.00433 -0.0209 0.6454 1.0000
3.250 0.5552 0.01051 0.00424 -0.0201 0.6296 1.0000
3.500 0.5803 0.01045 0.00417 -0.0193 0.6108 1.0000
3.750 0.6052 0.01042 0.00409 -0.0184 0.5864 1.0000
4.000 0.6296 0.01044 0.00400 -0.0175 0.5475 1.0000
4.250 0.6519 0.01072 0.00392 -0.0163 0.4560 1.0000
4.500 0.6707 0.01175 0.00426 -0.0151 0.3477 1.0000
4.750 0.6920 0.01253 0.00474 -0.0144 0.3080 1.0000
5.000 0.7143 0.01314 0.00518 -0.0136 0.2848 1.0000
5.250 0.7371 0.01369 0.00562 -0.0129 0.2680 1.0000
5.500 0.7598 0.01423 0.00608 -0.0122 0.2549 1.0000
5.750 0.7822 0.01479 0.00653 -0.0114 0.2434 1.0000
6.000 0.8056 0.01522 0.00696 -0.0108 0.2327 1.0000
6.250 0.8285 0.01574 0.00747 -0.0100 0.2237 1.0000
6.500 0.8512 0.01628 0.00794 -0.0093 0.2153 1.0000
6.750 0.8744 0.01674 0.00845 -0.0087 0.2069 1.0000
7.000 0.8970 0.01735 0.00898 -0.0080 0.1996 1.0000
7.250 0.9203 0.01780 0.00954 -0.0073 0.1923 1.0000
7.500 0.9430 0.01838 0.01005 -0.0067 0.1854 1.0000
7.750 0.9660 0.01890 0.01068 -0.0060 0.1786 1.0000
8.000 0.9887 0.01942 0.01120 -0.0054 0.1722 1.0000
8.250 1.0112 0.02006 0.01191 -0.0047 0.1658 1.0000
8.500 1.0337 0.02053 0.01244 -0.0040 0.1592 1.0000
8.750 1.0558 0.02127 0.01317 -0.0034 0.1531 1.0000
9.000 1.0775 0.02170 0.01374 -0.0027 0.1464 1.0000
9.250 1.0992 0.02250 0.01446 -0.0021 0.1404 1.0000
9.500 1.1199 0.02292 0.01511 -0.0012 0.1340 1.0000
9.750 1.1406 0.02356 0.01570 -0.0005 0.1281 1.0000
10.000 1.1602 0.02418 0.01651 0.0004 0.1217 1.0000
10.250 1.1794 0.02471 0.01707 0.0012 0.1156 1.0000
10.500 1.1976 0.02546 0.01796 0.0022 0.1092 1.0000
10.750 1.2148 0.02598 0.01853 0.0032 0.1030 1.0000
11.000 1.2308 0.02682 0.01949 0.0043 0.0966 1.0000
11.250 1.2453 0.02743 0.02016 0.0055 0.0905 1.0000
11.500 1.2583 0.02838 0.02124 0.0068 0.0842 1.0000
11.750 1.2677 0.02915 0.02205 0.0085 0.0785 1.0000
12.000 1.2757 0.03034 0.02337 0.0101 0.0728 1.0000
12.250 1.2829 0.03147 0.02456 0.0113 0.0674 1.0000
12.500 1.2888 0.03309 0.02628 0.0125 0.0622 1.0000
12.750 1.2946 0.03463 0.02790 0.0132 0.0573 1.0000
13.000 1.2971 0.03674 0.03010 0.0142 0.0530 1.0000
13.250 1.3008 0.03870 0.03220 0.0147 0.0487 1.0000
13.500 1.2985 0.04139 0.03482 0.0154 0.0454 1.0000
13.750 1.2998 0.04384 0.03754 0.0157 0.0424 1.0000
14.000 1.2991 0.04648 0.04028 0.0157 0.0397 1.0000
14.250 1.2941 0.04965 0.04343 0.0159 0.0375 1.0000
14.500 1.2895 0.05302 0.04702 0.0157 0.0357 1.0000
14.750 1.2844 0.05652 0.05071 0.0152 0.0340 1.0000
15.000 1.2788 0.06017 0.05447 0.0144 0.0326 1.0000
15.250 1.2723 0.06399 0.05837 0.0135 0.0315 1.0000
15.500 1.2644 0.06800 0.06242 0.0128 0.0305 1.0000
15.750 1.2535 0.07275 0.06731 0.0118 0.0297 1.0000
16.000 1.2415 0.07807 0.07285 0.0096 0.0292 1.0000
16.250 1.2283 0.08384 0.07884 0.0070 0.0288 1.0000
16.500 1.2137 0.09014 0.08533 0.0040 0.0284 1.0000
16.750 1.1972 0.09708 0.09246 0.0004 0.0281 1.0000
17.000 1.1786 0.10474 0.10031 -0.0038 0.0280 1.0000
17.250 1.1572 0.11329 0.10906 -0.0087 0.0280 1.0000
17.500 1.1316 0.12310 0.11906 -0.0144 0.0282 1.0000
17.750 1.1009 0.13451 0.13065 -0.0212 0.0286 1.0000
18.000 1.0617 0.14858 0.14489 -0.0296 0.0293 1.0000
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