NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 50,000 Max Cl/Cd: 28.87 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc410-il-50000.txt Download as CSV file: xf-rc410-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5360 0.11525 0.10872 0.0067 1.0000 0.2024
-8.750 -0.4979 0.10887 0.10228 0.0110 1.0000 0.2139
-8.500 -0.5288 0.10759 0.10117 0.0044 1.0000 0.2186
-8.250 -0.4970 0.10222 0.09578 0.0081 1.0000 0.2325
-8.000 -0.4867 0.09811 0.09171 0.0082 1.0000 0.2419
-7.750 -0.5184 0.09636 0.09014 0.0018 1.0000 0.2515
-7.500 -0.4983 0.09208 0.08589 0.0041 1.0000 0.2670
-7.250 -0.4854 0.08823 0.08209 0.0052 1.0000 0.2822
-7.000 -0.4909 0.08520 0.07916 0.0035 1.0000 0.3006
-6.750 -0.4727 0.08138 0.07538 0.0064 1.0000 0.3223
-6.500 -0.4637 0.07856 0.07265 0.0088 1.0000 0.3554
-6.250 -0.4608 0.07667 0.07085 0.0119 1.0000 0.4003
-6.000 -0.4118 0.07378 0.06789 0.0205 1.0000 0.4742
-5.750 -0.3739 0.07189 0.06598 0.0286 1.0000 0.5785
-5.500 -0.3024 0.06788 0.06183 0.0345 1.0000 0.7299
-4.750 -0.3118 0.05850 0.05276 0.0285 1.0000 0.6683
-4.500 -0.4513 0.04758 0.04034 -0.0107 1.0000 0.2121
-4.250 -0.4341 0.04412 0.03603 -0.0106 1.0000 0.1846
-4.000 -0.4191 0.04136 0.03299 -0.0094 1.0000 0.1829
-3.750 -0.4017 0.03880 0.03005 -0.0084 1.0000 0.1819
-3.500 -0.3824 0.03685 0.02748 -0.0074 1.0000 0.1858
-3.250 -0.3634 0.03450 0.02506 -0.0067 1.0000 0.1935
-3.000 -0.3411 0.03304 0.02299 -0.0059 1.0000 0.2026
-2.750 -0.3193 0.03117 0.02106 -0.0054 1.0000 0.2150
-2.500 -0.2956 0.02965 0.01937 -0.0051 1.0000 0.2275
-2.250 -0.2708 0.02842 0.01790 -0.0048 1.0000 0.2403
-2.000 -0.2460 0.02747 0.01675 -0.0046 1.0000 0.2539
-1.750 -0.2208 0.02671 0.01585 -0.0044 1.0000 0.2681
-1.500 -0.1947 0.02593 0.01507 -0.0043 1.0000 0.2822
-1.250 -0.1679 0.02537 0.01450 -0.0044 1.0000 0.2995
-1.000 -0.1423 0.02489 0.01405 -0.0043 1.0000 0.3173
-0.750 -0.1181 0.02436 0.01372 -0.0042 1.0000 0.3422
-0.500 -0.0941 0.02374 0.01349 -0.0042 1.0000 0.3852
-0.250 -0.0492 0.02204 0.01387 -0.0061 1.0000 1.0000
0.000 -0.0298 0.02247 0.01390 -0.0051 1.0000 1.0000
0.250 -0.0112 0.02292 0.01403 -0.0043 1.0000 1.0000
0.500 0.0073 0.02340 0.01425 -0.0035 1.0000 1.0000
0.750 0.0257 0.02390 0.01455 -0.0029 1.0000 1.0000
1.000 0.1063 0.02526 0.01564 -0.0138 0.9744 1.0000
1.250 0.2048 0.02631 0.01654 -0.0267 0.9356 1.0000
1.500 0.2691 0.02688 0.01707 -0.0331 0.9115 1.0000
1.750 0.3288 0.02725 0.01746 -0.0383 0.8876 1.0000
2.000 0.4019 0.02717 0.01747 -0.0446 0.8611 1.0000
2.250 0.4432 0.02709 0.01746 -0.0451 0.8335 1.0000
2.500 0.4827 0.02672 0.01717 -0.0443 0.8043 1.0000
2.750 0.5143 0.02609 0.01661 -0.0414 0.7719 1.0000
3.000 0.5398 0.02516 0.01572 -0.0368 0.7355 1.0000
3.250 0.5595 0.02394 0.01452 -0.0310 0.6890 1.0000
3.500 0.5776 0.02242 0.01288 -0.0247 0.6261 1.0000
3.750 0.5960 0.02146 0.01153 -0.0195 0.5492 1.0000
4.000 0.6160 0.02150 0.01103 -0.0163 0.4869 1.0000
4.250 0.6374 0.02208 0.01114 -0.0143 0.4461 1.0000
4.500 0.6603 0.02294 0.01167 -0.0130 0.4174 1.0000
4.750 0.6841 0.02390 0.01233 -0.0119 0.3956 1.0000
5.000 0.7083 0.02495 0.01325 -0.0110 0.3770 1.0000
5.250 0.7333 0.02609 0.01428 -0.0104 0.3628 1.0000
5.500 0.7585 0.02729 0.01536 -0.0097 0.3501 1.0000
5.750 0.7822 0.02856 0.01673 -0.0092 0.3378 1.0000
6.000 0.8060 0.02998 0.01822 -0.0086 0.3276 1.0000
6.250 0.8305 0.03144 0.01968 -0.0081 0.3191 1.0000
6.500 0.8521 0.03313 0.02160 -0.0076 0.3109 1.0000
6.750 0.8757 0.03471 0.02315 -0.0071 0.3024 1.0000
7.000 0.8946 0.03676 0.02555 -0.0065 0.2957 1.0000
7.250 0.9143 0.03881 0.02783 -0.0059 0.2898 1.0000
7.500 0.9361 0.04090 0.02992 -0.0054 0.2838 1.0000
7.750 0.9467 0.04364 0.03317 -0.0047 0.2781 1.0000
8.000 0.9613 0.04625 0.03604 -0.0041 0.2730 1.0000
8.250 0.9833 0.04869 0.03846 -0.0036 0.2686 1.0000
8.500 0.9819 0.05284 0.04312 -0.0030 0.2661 1.0000
8.750 0.9728 0.05776 0.04847 -0.0027 0.2640 1.0000
9.000 0.9532 0.06383 0.05486 -0.0030 0.2635 1.0000
9.250 0.9239 0.07129 0.06250 -0.0045 0.2655 1.0000
9.500 0.9008 0.07875 0.07004 -0.0066 0.2690 1.0000
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Polar data table (+)
Polar graphs
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