NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Reynolds number: 500,000 Max Cl/Cd: 54.36 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-a63a108c-il-500000.txt Download as CSV file: xf-a63a108c-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.7192 0.10431 0.10208 0.0325 1.0000 0.0209
-9.000 -0.7211 0.09903 0.09682 0.0293 1.0000 0.0213
-8.250 -0.7865 0.06171 0.05903 -0.0028 1.0000 0.0225
-8.000 -0.7746 0.05976 0.05713 -0.0027 1.0000 0.0227
-7.750 -0.7626 0.05798 0.05534 -0.0028 1.0000 0.0229
-7.500 -0.7506 0.05550 0.05280 -0.0033 1.0000 0.0233
-7.250 -0.7387 0.05184 0.04899 -0.0041 1.0000 0.0240
-7.000 -0.7353 0.04234 0.03883 -0.0051 1.0000 0.0261
-6.750 -0.7397 0.02294 0.01727 -0.0025 1.0000 0.0169
-6.500 -0.7143 0.02161 0.01590 -0.0025 1.0000 0.0173
-6.250 -0.6883 0.02022 0.01435 -0.0024 1.0000 0.0177
-6.000 -0.6618 0.01866 0.01256 -0.0022 1.0000 0.0181
-5.750 -0.6347 0.01733 0.01101 -0.0020 1.0000 0.0188
-5.500 -0.6072 0.01621 0.00966 -0.0019 1.0000 0.0197
-5.250 -0.5798 0.01540 0.00888 -0.0020 1.0000 0.0206
-5.000 -0.5519 0.01457 0.00798 -0.0020 1.0000 0.0214
-4.750 -0.5236 0.01384 0.00713 -0.0019 1.0000 0.0224
-4.500 -0.4960 0.01295 0.00628 -0.0020 1.0000 0.0234
-4.250 -0.4676 0.01239 0.00571 -0.0021 1.0000 0.0248
-4.000 -0.4391 0.01177 0.00508 -0.0023 1.0000 0.0264
-3.750 -0.4100 0.01137 0.00468 -0.0026 1.0000 0.0288
-3.500 -0.3811 0.01083 0.00418 -0.0029 1.0000 0.0312
-3.250 -0.3518 0.01036 0.00374 -0.0033 1.0000 0.0344
-3.000 -0.3222 0.00997 0.00338 -0.0037 1.0000 0.0394
-2.750 -0.2925 0.00958 0.00304 -0.0041 1.0000 0.0471
-2.500 -0.2630 0.00917 0.00272 -0.0046 1.0000 0.0620
-2.250 -0.2340 0.00859 0.00245 -0.0051 1.0000 0.1180
-2.000 -0.2066 0.00777 0.00221 -0.0057 1.0000 0.2564
-1.750 -0.1739 0.00673 0.00206 -0.0074 0.9896 0.4905
-1.500 -0.1412 0.00628 0.00208 -0.0085 0.9765 0.6191
-1.250 -0.1163 0.00618 0.00219 -0.0073 0.9579 0.6866
-1.000 -0.0955 0.00613 0.00226 -0.0052 0.9398 0.7352
-0.750 -0.0734 0.00606 0.00230 -0.0033 0.9258 0.7779
-0.500 -0.0502 0.00595 0.00230 -0.0018 0.9126 0.8176
-0.250 -0.0284 0.00579 0.00229 0.0002 0.8987 0.8636
0.000 -0.0078 0.00563 0.00224 0.0027 0.8851 0.9182
0.250 0.0258 0.00550 0.00213 0.0021 0.8715 0.9810
0.500 0.0608 0.00547 0.00204 0.0008 0.8547 1.0000
0.750 0.0877 0.00547 0.00197 0.0012 0.8355 1.0000
1.000 0.1146 0.00550 0.00190 0.0017 0.8126 1.0000
1.250 0.1416 0.00555 0.00185 0.0022 0.7852 1.0000
1.500 0.1688 0.00565 0.00180 0.0026 0.7470 1.0000
1.750 0.1961 0.00583 0.00177 0.0029 0.6945 1.0000
2.000 0.2237 0.00612 0.00177 0.0031 0.6179 1.0000
2.250 0.2516 0.00661 0.00183 0.0030 0.5121 1.0000
2.500 0.2798 0.00720 0.00195 0.0027 0.3998 1.0000
2.750 0.3083 0.00773 0.00210 0.0023 0.3099 1.0000
3.000 0.3368 0.00817 0.00226 0.0019 0.2460 1.0000
3.250 0.3653 0.00855 0.00243 0.0016 0.1996 1.0000
3.500 0.3939 0.00890 0.00260 0.0014 0.1635 1.0000
3.750 0.4224 0.00923 0.00280 0.0011 0.1359 1.0000
4.000 0.4509 0.00956 0.00301 0.0009 0.1144 1.0000
4.250 0.4793 0.00989 0.00325 0.0007 0.0974 1.0000
4.500 0.5077 0.01023 0.00351 0.0006 0.0842 1.0000
4.750 0.5360 0.01056 0.00378 0.0004 0.0736 1.0000
5.000 0.5642 0.01091 0.00409 0.0003 0.0650 1.0000
5.250 0.5923 0.01130 0.00443 0.0001 0.0582 1.0000
5.500 0.6203 0.01169 0.00479 0.0000 0.0525 1.0000
5.750 0.6482 0.01207 0.00517 -0.0001 0.0477 1.0000
6.000 0.6761 0.01248 0.00558 -0.0001 0.0435 1.0000
6.250 0.7036 0.01295 0.00605 -0.0002 0.0401 1.0000
6.500 0.7309 0.01350 0.00659 -0.0002 0.0374 1.0000
6.750 0.7583 0.01395 0.00706 -0.0002 0.0348 1.0000
7.000 0.7852 0.01454 0.00768 -0.0002 0.0326 1.0000
7.250 0.8120 0.01507 0.00820 -0.0002 0.0308 1.0000
7.500 0.8384 0.01578 0.00898 -0.0001 0.0293 1.0000
7.750 0.8649 0.01639 0.00963 0.0000 0.0280 1.0000
8.000 0.8901 0.01730 0.01051 0.0001 0.0269 1.0000
8.250 0.9161 0.01805 0.01139 0.0003 0.0260 1.0000
8.500 0.9418 0.01878 0.01221 0.0005 0.0250 1.0000
8.750 0.9670 0.01950 0.01295 0.0006 0.0241 1.0000
9.000 0.9910 0.02065 0.01418 0.0009 0.0233 1.0000
9.250 1.0153 0.02167 0.01540 0.0012 0.0227 1.0000
9.500 1.0391 0.02274 0.01660 0.0015 0.0220 1.0000
9.750 1.0624 0.02379 0.01777 0.0018 0.0215 1.0000
10.000 1.0849 0.02496 0.01899 0.0021 0.0210 1.0000
10.250 1.1045 0.02696 0.02113 0.0026 0.0206 1.0000
10.500 1.1233 0.02875 0.02329 0.0033 0.0202 1.0000
10.750 1.1394 0.03094 0.02584 0.0041 0.0198 1.0000
11.000 1.1540 0.03309 0.02829 0.0049 0.0193 1.0000
11.250 1.1687 0.03492 0.03032 0.0056 0.0189 1.0000
11.500 1.1838 0.03644 0.03194 0.0063 0.0186 1.0000
11.750 1.1947 0.03844 0.03410 0.0071 0.0183 1.0000
12.000 1.2015 0.04082 0.03664 0.0081 0.0181 1.0000
12.250 1.1989 0.04409 0.04012 0.0093 0.0180 1.0000
12.500 1.1834 0.04774 0.04401 0.0112 0.0179 1.0000
12.750 1.1631 0.05217 0.04868 0.0111 0.0179 1.0000
13.000 1.1408 0.05812 0.05489 0.0078 0.0179 1.0000
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Polar data table (+)
Polar graphs
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