NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Reynolds number: 500,000 Max Cl/Cd: 58.78 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-a63a108c-il-500000-n5.txt Download as CSV file: xf-a63a108c-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES 63A108 MOD C AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.7634 0.10463 0.10234 0.0341 1.0000 0.0088
-9.750 -0.7751 0.09686 0.09459 0.0297 1.0000 0.0089
-9.250 -0.9797 0.03145 0.02723 -0.0032 1.0000 0.0086
-9.000 -0.9682 0.02699 0.02214 -0.0025 1.0000 0.0089
-8.750 -0.9488 0.02452 0.01925 -0.0020 1.0000 0.0091
-8.500 -0.9269 0.02274 0.01718 -0.0016 1.0000 0.0093
-8.250 -0.9030 0.02156 0.01587 -0.0014 1.0000 0.0095
-8.000 -0.8781 0.02061 0.01482 -0.0013 1.0000 0.0098
-7.750 -0.8527 0.01975 0.01384 -0.0011 1.0000 0.0100
-7.500 -0.8270 0.01885 0.01282 -0.0010 1.0000 0.0103
-7.250 -0.8010 0.01795 0.01179 -0.0009 1.0000 0.0106
-7.000 -0.7748 0.01707 0.01077 -0.0008 1.0000 0.0110
-6.750 -0.7481 0.01626 0.00983 -0.0007 1.0000 0.0114
-6.500 -0.7213 0.01556 0.00905 -0.0006 1.0000 0.0118
-6.250 -0.6939 0.01506 0.00853 -0.0007 1.0000 0.0123
-6.000 -0.6664 0.01454 0.00797 -0.0007 1.0000 0.0129
-5.750 -0.6387 0.01397 0.00732 -0.0008 1.0000 0.0135
-5.500 -0.6109 0.01342 0.00670 -0.0008 1.0000 0.0140
-5.250 -0.5830 0.01287 0.00615 -0.0009 1.0000 0.0146
-5.000 -0.5547 0.01244 0.00570 -0.0011 1.0000 0.0152
-4.750 -0.5262 0.01203 0.00526 -0.0013 1.0000 0.0161
-4.500 -0.4976 0.01161 0.00482 -0.0015 1.0000 0.0170
-4.250 -0.4687 0.01128 0.00450 -0.0017 1.0000 0.0183
-4.000 -0.4397 0.01094 0.00415 -0.0020 1.0000 0.0198
-3.750 -0.4105 0.01062 0.00384 -0.0023 1.0000 0.0214
-3.500 -0.3793 0.01036 0.00357 -0.0030 0.9768 0.0234
-3.250 -0.3499 0.01015 0.00336 -0.0032 0.9643 0.0261
-3.000 -0.3227 0.00996 0.00318 -0.0030 0.9525 0.0298
-2.750 -0.2963 0.00978 0.00301 -0.0025 0.9416 0.0346
-2.500 -0.2702 0.00961 0.00286 -0.0019 0.9318 0.0415
-2.000 -0.2173 0.00920 0.00257 -0.0010 0.9127 0.0746
-1.750 -0.1906 0.00892 0.00244 -0.0006 0.9034 0.1141
-1.500 -0.1643 0.00857 0.00230 -0.0002 0.8916 0.1764
-1.250 -0.1383 0.00811 0.00215 0.0001 0.8752 0.2715
-1.000 -0.1122 0.00752 0.00200 0.0004 0.8564 0.4084
-0.750 -0.0855 0.00704 0.00192 0.0006 0.8380 0.5316
-0.500 -0.0587 0.00674 0.00188 0.0010 0.8195 0.6176
-0.250 -0.0319 0.00655 0.00186 0.0015 0.7970 0.6850
0.000 -0.0051 0.00643 0.00183 0.0020 0.7700 0.7358
0.250 0.0214 0.00633 0.00181 0.0027 0.7403 0.7835
0.500 0.0469 0.00625 0.00177 0.0036 0.7012 0.8337
0.750 0.0720 0.00626 0.00174 0.0046 0.6529 0.8748
1.000 0.0973 0.00638 0.00170 0.0055 0.5853 0.9132
1.250 0.1270 0.00662 0.00167 0.0053 0.5033 0.9549
1.750 0.1927 0.00739 0.00176 0.0027 0.3323 1.0000
2.000 0.2213 0.00773 0.00184 0.0024 0.2749 1.0000
2.250 0.2500 0.00801 0.00194 0.0022 0.2319 1.0000
2.500 0.2786 0.00829 0.00205 0.0019 0.1958 1.0000
2.750 0.3073 0.00854 0.00216 0.0017 0.1669 1.0000
3.000 0.3360 0.00880 0.00229 0.0015 0.1425 1.0000
3.250 0.3646 0.00904 0.00244 0.0013 0.1231 1.0000
3.500 0.3932 0.00929 0.00259 0.0011 0.1062 1.0000
3.750 0.4218 0.00952 0.00275 0.0009 0.0925 1.0000
4.000 0.4503 0.00977 0.00293 0.0007 0.0808 1.0000
4.250 0.4788 0.01003 0.00314 0.0006 0.0711 1.0000
4.500 0.5072 0.01028 0.00335 0.0004 0.0629 1.0000
4.750 0.5355 0.01055 0.00357 0.0003 0.0558 1.0000
5.000 0.5638 0.01082 0.00382 0.0001 0.0497 1.0000
5.250 0.5920 0.01110 0.00408 0.0000 0.0448 1.0000
5.500 0.6202 0.01140 0.00436 -0.0001 0.0408 1.0000
5.750 0.6482 0.01171 0.00466 -0.0002 0.0372 1.0000
6.250 0.7040 0.01235 0.00530 -0.0004 0.0312 1.0000
6.500 0.7317 0.01271 0.00568 -0.0005 0.0291 1.0000
6.750 0.7593 0.01309 0.00606 -0.0005 0.0272 1.0000
7.000 0.7868 0.01348 0.00648 -0.0006 0.0254 1.0000
7.250 0.8141 0.01390 0.00691 -0.0006 0.0239 1.0000
7.500 0.8412 0.01432 0.00738 -0.0006 0.0225 1.0000
7.750 0.8682 0.01477 0.00786 -0.0006 0.0214 1.0000
8.000 0.8948 0.01529 0.00841 -0.0006 0.0205 1.0000
8.250 0.9214 0.01580 0.00899 -0.0006 0.0196 1.0000
8.500 0.9478 0.01633 0.00958 -0.0005 0.0188 1.0000
8.750 0.9738 0.01690 0.01020 -0.0005 0.0182 1.0000
9.000 0.9993 0.01756 0.01091 -0.0004 0.0176 1.0000
9.250 1.0249 0.01818 0.01163 -0.0002 0.0170 1.0000
9.500 1.0502 0.01881 0.01234 -0.0001 0.0164 1.0000
9.750 1.0751 0.01946 0.01305 0.0000 0.0158 1.0000
10.000 1.0993 0.02022 0.01387 0.0002 0.0154 1.0000
10.250 1.1228 0.02110 0.01485 0.0005 0.0151 1.0000
10.500 1.1461 0.02196 0.01586 0.0008 0.0147 1.0000
10.750 1.1688 0.02289 0.01691 0.0011 0.0144 1.0000
11.000 1.1908 0.02386 0.01801 0.0015 0.0140 1.0000
11.250 1.2121 0.02486 0.01913 0.0019 0.0137 1.0000
11.500 1.2327 0.02590 0.02029 0.0024 0.0135 1.0000
11.750 1.2524 0.02700 0.02149 0.0029 0.0133 1.0000
12.000 1.2707 0.02821 0.02281 0.0035 0.0130 1.0000
12.250 1.2862 0.02970 0.02441 0.0042 0.0128 1.0000
12.500 1.3014 0.03118 0.02611 0.0050 0.0126 1.0000
12.750 1.3141 0.03282 0.02796 0.0060 0.0123 1.0000
13.000 1.3240 0.03463 0.02997 0.0070 0.0121 1.0000
13.250 1.3298 0.03662 0.03217 0.0082 0.0119 1.0000
13.500 1.3279 0.03883 0.03456 0.0100 0.0118 1.0000
13.750 1.3212 0.04162 0.03753 0.0110 0.0117 1.0000
14.000 1.3122 0.04529 0.04140 0.0104 0.0117 1.0000
14.250 1.2991 0.05038 0.04669 0.0078 0.0116 1.0000
14.500 1.2789 0.05813 0.05469 0.0018 0.0116 1.0000
14.750 1.2421 0.07175 0.06861 -0.0094 0.0117 1.0000
15.000 1.1777 0.09091 0.08804 -0.0219 0.0119 1.0000
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Polar data table (+)
Polar graphs
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