NACA 63A210 (naca63a210-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 63A210 (naca63a210-il) Reynolds number: 1,000,000 Max Cl/Cd: 58.53 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca63a210-il-1000000.txt Download as CSV file: xf-naca63a210-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A210
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5909 0.09156 0.08901 -0.0078 0.4819 0.0066
-8.750 -0.5995 0.08117 0.07862 -0.0195 0.4819 0.0065
-8.500 -0.6164 0.07533 0.07269 -0.0240 0.4819 0.0065
-8.250 -0.6260 0.07145 0.06871 -0.0247 0.4819 0.0065
-8.000 -0.6281 0.06773 0.06489 -0.0252 0.4819 0.0066
-7.750 -0.6260 0.06426 0.06129 -0.0254 0.4819 0.0067
-7.500 -0.6207 0.06086 0.05775 -0.0254 0.4819 0.0068
-7.250 -0.6125 0.05756 0.05429 -0.0251 0.4819 0.0070
-7.000 -0.6018 0.05433 0.05088 -0.0247 0.4819 0.0071
-6.750 -0.5889 0.05121 0.04757 -0.0241 0.4819 0.0073
-6.500 -0.5741 0.04820 0.04436 -0.0233 0.4819 0.0076
-6.250 -0.5575 0.04530 0.04123 -0.0224 0.4819 0.0080
-6.000 -0.5389 0.04257 0.03826 -0.0214 0.4820 0.0085
-5.750 -0.5173 0.04040 0.03585 -0.0203 0.4820 0.0093
-5.500 -0.4906 0.04029 0.03559 -0.0195 0.4820 0.0101
-4.750 -0.4284 0.02458 0.01849 -0.0138 0.4822 0.0065
-4.500 -0.4034 0.02261 0.01643 -0.0131 0.4823 0.0070
-4.250 -0.3768 0.02171 0.01551 -0.0129 0.4825 0.0077
-4.000 -0.3504 0.02088 0.01465 -0.0126 0.4827 0.0085
-3.750 -0.3245 0.02006 0.01378 -0.0121 0.4828 0.0090
-3.500 -0.2988 0.01935 0.01302 -0.0117 0.4831 0.0094
-3.250 -0.2715 0.01915 0.01279 -0.0116 0.4833 0.0101
-3.000 -0.2465 0.01847 0.01205 -0.0111 0.4835 0.0105
-2.750 -0.2217 0.01769 0.01112 -0.0104 0.4837 0.0119
-2.500 -0.1950 0.01743 0.01083 -0.0102 0.4839 0.0147
-2.250 -0.1678 0.01731 0.01068 -0.0101 0.4841 0.0171
-2.000 -0.1440 0.01646 0.01031 -0.0096 0.4844 0.1651
-1.750 -0.1236 0.01521 0.01007 -0.0090 0.4848 0.4551
-1.500 -0.0984 0.01493 0.01010 -0.0087 0.4852 0.5517
-1.250 -0.0732 0.01479 0.01022 -0.0084 0.4856 0.6298
-1.000 -0.0481 0.01477 0.01041 -0.0079 0.4861 0.7008
-0.750 -0.0215 0.01490 0.01059 -0.0077 0.4863 0.7278
-0.500 0.0057 0.01505 0.01076 -0.0077 0.4865 0.7400
-0.250 0.0329 0.01525 0.01096 -0.0077 0.4867 0.7504
0.000 0.0599 0.01553 0.01124 -0.0077 0.4870 0.7616
0.250 0.0869 0.01581 0.01153 -0.0077 0.4871 0.7734
0.500 0.1139 0.01599 0.01175 -0.0076 0.4872 0.7845
0.750 0.1412 0.01611 0.01190 -0.0076 0.4872 0.7957
1.000 0.1684 0.01617 0.01201 -0.0076 0.4872 0.8084
1.250 0.1957 0.01615 0.01205 -0.0075 0.4871 0.8234
1.500 0.2233 0.01589 0.01188 -0.0074 0.4868 0.8406
1.750 0.2507 0.01561 0.01170 -0.0073 0.4861 0.8620
2.000 0.2769 0.01548 0.01168 -0.0069 0.4848 0.8906
2.250 0.3020 0.01579 0.01208 -0.0063 0.4831 0.9259
2.500 0.3281 0.01813 0.01436 -0.0068 0.4820 0.9561
2.750 0.3709 0.01535 0.01168 -0.0089 0.4782 0.9730
3.000 0.4090 0.01452 0.01088 -0.0109 0.4748 0.9834
3.250 0.4489 0.01353 0.00988 -0.0130 0.4562 0.9894
3.500 0.4855 0.01315 0.00943 -0.0148 0.4480 0.9945
3.750 0.5226 0.01321 0.00952 -0.0170 0.4419 0.9980
4.000 0.5550 0.01337 0.00975 -0.0183 0.4361 1.0000
4.250 0.5798 0.01354 0.00995 -0.0178 0.4290 1.0000
4.500 0.6041 0.01388 0.01035 -0.0174 0.4138 1.0000
4.750 0.6294 0.01378 0.00951 -0.0171 0.2693 1.0000
5.000 0.6566 0.01336 0.00867 -0.0169 0.1944 1.0000
5.250 0.6824 0.01328 0.00833 -0.0166 0.1444 1.0000
5.500 0.7076 0.01333 0.00820 -0.0163 0.1078 1.0000
5.750 0.7325 0.01345 0.00817 -0.0159 0.0795 1.0000
6.000 0.7574 0.01362 0.00823 -0.0155 0.0588 1.0000
6.250 0.7825 0.01382 0.00837 -0.0152 0.0442 1.0000
6.500 0.8071 0.01407 0.00853 -0.0148 0.0276 1.0000
6.750 0.8313 0.01438 0.00878 -0.0143 0.0152 1.0000
7.000 0.8559 0.01467 0.00909 -0.0139 0.0108 1.0000
7.250 0.8797 0.01504 0.00952 -0.0133 0.0076 1.0000
7.500 0.9037 0.01544 0.00995 -0.0128 0.0063 1.0000
7.750 0.9258 0.01606 0.01065 -0.0120 0.0054 1.0000
8.000 0.9442 0.01715 0.01190 -0.0106 0.0048 1.0000
8.250 0.9676 0.01765 0.01245 -0.0102 0.0046 1.0000
8.500 0.9886 0.01841 0.01332 -0.0094 0.0045 1.0000
8.750 1.0084 0.01928 0.01429 -0.0084 0.0043 1.0000
9.000 1.0274 0.02022 0.01533 -0.0074 0.0042 1.0000
9.250 1.0458 0.02121 0.01643 -0.0063 0.0040 1.0000
9.500 1.0639 0.02223 0.01754 -0.0053 0.0038 1.0000
9.750 1.0814 0.02330 0.01872 -0.0042 0.0037 1.0000
10.000 1.0989 0.02434 0.01986 -0.0033 0.0035 1.0000
10.250 1.1161 0.02539 0.02099 -0.0024 0.0033 1.0000
10.500 1.1315 0.02657 0.02225 -0.0015 0.0031 1.0000
10.750 1.1390 0.02858 0.02442 0.0002 0.0029 1.0000
11.000 1.1275 0.03252 0.02870 0.0036 0.0027 1.0000
11.250 1.1110 0.03677 0.03327 0.0062 0.0026 1.0000
11.500 1.0992 0.04102 0.03778 0.0063 0.0026 1.0000
11.750 1.0893 0.04623 0.04324 0.0034 0.0026 1.0000
12.000 1.0805 0.05126 0.04845 0.0004 0.0026 1.0000
12.250 1.0697 0.05594 0.05328 -0.0013 0.0026 1.0000
12.500 1.0572 0.06062 0.05809 -0.0031 0.0026 1.0000
12.750 1.0438 0.06546 0.06307 -0.0049 0.0026 1.0000
13.000 1.0298 0.07052 0.06824 -0.0069 0.0027 1.0000
13.250 1.0150 0.07616 0.07400 -0.0096 0.0027 1.0000
13.500 1.0008 0.08234 0.08029 -0.0128 0.0027 1.0000
13.750 0.9866 0.08946 0.08752 -0.0168 0.0027 1.0000
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Polar data table (+)
Polar graphs
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