NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 100,000 Max Cl/Cd: 37.86 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc410-il-100000-n5.txt Download as CSV file: xf-rc410-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(4)-10 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5593 0.09319 0.08833 -0.0014 1.0000 0.0379
-9.000 -0.5632 0.08750 0.08268 -0.0057 1.0000 0.0379
-8.750 -0.5690 0.08134 0.07657 -0.0115 1.0000 0.0378
-8.500 -0.5765 0.07535 0.07057 -0.0162 1.0000 0.0377
-8.250 -0.5811 0.06921 0.06436 -0.0200 1.0000 0.0375
-8.000 -0.5835 0.06303 0.05804 -0.0230 1.0000 0.0374
-7.750 -0.5844 0.05652 0.05131 -0.0251 1.0000 0.0372
-7.500 -0.5837 0.04954 0.04399 -0.0264 1.0000 0.0373
-7.250 -0.5845 0.04135 0.03519 -0.0266 1.0000 0.0375
-7.000 -0.5796 0.03473 0.02776 -0.0258 0.9960 0.0382
-6.750 -0.5533 0.03157 0.02416 -0.0271 0.9641 0.0394
-6.500 -0.5228 0.03055 0.02302 -0.0285 0.9460 0.0412
-6.250 -0.4958 0.02843 0.02049 -0.0289 0.9310 0.0435
-6.000 -0.4703 0.02587 0.01734 -0.0287 0.9183 0.0457
-5.750 -0.4444 0.02467 0.01600 -0.0285 0.9073 0.0475
-5.500 -0.4190 0.02382 0.01502 -0.0281 0.8967 0.0499
-5.250 -0.3934 0.02261 0.01348 -0.0275 0.8872 0.0542
-5.000 -0.3683 0.02223 0.01312 -0.0269 0.8788 0.0584
-4.750 -0.3422 0.02136 0.01204 -0.0264 0.8706 0.0647
-4.500 -0.3168 0.02109 0.01176 -0.0259 0.8634 0.0706
-4.250 -0.2906 0.02075 0.01134 -0.0254 0.8559 0.0787
-4.000 -0.2645 0.02050 0.01094 -0.0249 0.8489 0.0884
-3.750 -0.2384 0.02042 0.01081 -0.0245 0.8425 0.0971
-3.500 -0.2119 0.02015 0.01049 -0.0242 0.8361 0.1054
-3.250 -0.1854 0.01991 0.01007 -0.0237 0.8308 0.1145
-3.000 -0.1589 0.01953 0.00973 -0.0235 0.8246 0.1212
-2.750 -0.1320 0.01925 0.00932 -0.0231 0.8187 0.1293
-2.500 -0.1060 0.01883 0.00893 -0.0225 0.8141 0.1353
-2.250 -0.0787 0.01854 0.00861 -0.0224 0.8083 0.1428
-2.000 -0.0522 0.01819 0.00827 -0.0221 0.8032 0.1491
-1.750 -0.0265 0.01789 0.00799 -0.0214 0.7988 0.1559
-1.500 0.0001 0.01765 0.00775 -0.0212 0.7931 0.1635
-1.250 0.0254 0.01738 0.00752 -0.0205 0.7867 0.1710
-1.000 0.0504 0.01715 0.00726 -0.0196 0.7790 0.1792
-0.750 0.0749 0.01689 0.00699 -0.0186 0.7689 0.1898
-0.500 0.0997 0.01661 0.00674 -0.0176 0.7576 0.2018
-0.250 0.1246 0.01631 0.00647 -0.0166 0.7478 0.2190
0.000 0.1502 0.01582 0.00624 -0.0160 0.7375 0.2738
0.500 0.2498 0.01386 0.00633 -0.0222 0.7149 0.9740
0.750 0.2994 0.01377 0.00610 -0.0262 0.7017 1.0000
1.000 0.3239 0.01373 0.00599 -0.0253 0.6879 1.0000
1.250 0.3481 0.01370 0.00587 -0.0244 0.6733 1.0000
1.500 0.3722 0.01365 0.00573 -0.0233 0.6562 1.0000
1.750 0.3962 0.01360 0.00559 -0.0222 0.6361 1.0000
2.000 0.4205 0.01358 0.00550 -0.0213 0.6118 1.0000
2.250 0.4446 0.01358 0.00540 -0.0202 0.5820 1.0000
2.500 0.4680 0.01363 0.00528 -0.0190 0.5333 1.0000
2.750 0.4881 0.01403 0.00502 -0.0172 0.4085 1.0000
3.000 0.5077 0.01496 0.00529 -0.0161 0.3240 1.0000
3.250 0.5297 0.01560 0.00566 -0.0153 0.2910 1.0000
3.500 0.5524 0.01612 0.00601 -0.0145 0.2706 1.0000
3.750 0.5750 0.01660 0.00636 -0.0137 0.2557 1.0000
4.000 0.5977 0.01706 0.00671 -0.0129 0.2440 1.0000
4.250 0.6207 0.01748 0.00708 -0.0121 0.2347 1.0000
4.500 0.6432 0.01797 0.00745 -0.0112 0.2265 1.0000
4.750 0.6664 0.01838 0.00784 -0.0104 0.2185 1.0000
5.000 0.6891 0.01887 0.00825 -0.0095 0.2116 1.0000
5.250 0.7122 0.01934 0.00869 -0.0087 0.2055 1.0000
5.500 0.7355 0.01981 0.00914 -0.0080 0.1996 1.0000
5.750 0.7582 0.02038 0.00961 -0.0072 0.1945 1.0000
6.000 0.7818 0.02086 0.01015 -0.0065 0.1889 1.0000
6.250 0.8051 0.02138 0.01067 -0.0058 0.1836 1.0000
6.500 0.8281 0.02197 0.01120 -0.0051 0.1795 1.0000
6.750 0.8516 0.02257 0.01185 -0.0044 0.1752 1.0000
7.000 0.8750 0.02313 0.01249 -0.0038 0.1704 1.0000
7.250 0.8980 0.02372 0.01308 -0.0031 0.1661 1.0000
7.500 0.9210 0.02442 0.01372 -0.0025 0.1625 1.0000
7.750 0.9442 0.02508 0.01456 -0.0019 0.1586 1.0000
8.000 0.9671 0.02575 0.01533 -0.0013 0.1544 1.0000
8.250 0.9896 0.02641 0.01599 -0.0007 0.1507 1.0000
8.500 1.0121 0.02718 0.01675 -0.0001 0.1475 1.0000
8.750 1.0341 0.02799 0.01780 0.0006 0.1438 1.0000
9.000 1.0558 0.02880 0.01874 0.0013 0.1401 1.0000
9.250 1.0773 0.02952 0.01952 0.0019 0.1367 1.0000
9.500 1.0990 0.03031 0.02027 0.0025 0.1339 1.0000
9.750 1.1186 0.03135 0.02159 0.0033 0.1304 1.0000
10.000 1.1380 0.03237 0.02282 0.0041 0.1270 1.0000
10.250 1.1573 0.03323 0.02379 0.0049 0.1238 1.0000
10.500 1.1772 0.03399 0.02458 0.0056 0.1211 1.0000
10.750 1.1947 0.03515 0.02590 0.0064 0.1183 1.0000
11.000 1.2093 0.03653 0.02759 0.0075 0.1150 1.0000
11.250 1.2243 0.03766 0.02892 0.0086 0.1119 1.0000
11.500 1.2403 0.03854 0.02988 0.0095 0.1092 1.0000
11.750 1.2579 0.03934 0.03064 0.0103 0.1070 1.0000
12.000 1.2640 0.04124 0.03293 0.0119 0.1041 1.0000
12.250 1.2692 0.04301 0.03499 0.0134 0.1013 1.0000
12.500 1.2746 0.04439 0.03654 0.0150 0.0988 1.0000
12.750 1.2821 0.04545 0.03767 0.0164 0.0967 1.0000
13.000 1.2933 0.04638 0.03859 0.0175 0.0948 1.0000
13.250 1.2856 0.04910 0.04162 0.0189 0.0929 1.0000
13.500 1.2724 0.05252 0.04538 0.0196 0.0909 1.0000
13.750 1.2594 0.05610 0.04922 0.0198 0.0891 1.0000
14.000 1.2475 0.05976 0.05309 0.0194 0.0874 1.0000
14.250 1.2388 0.06322 0.05670 0.0188 0.0859 1.0000
14.500 1.2411 0.06542 0.05894 0.0185 0.0843 1.0000
14.750 1.2484 0.06707 0.06056 0.0185 0.0826 1.0000
15.000 1.2056 0.07593 0.06975 0.0145 0.0822 1.0000
15.250 1.1304 0.09187 0.08600 0.0051 0.0825 1.0000
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Polar data table (+)
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