NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 200,000 Max Cl/Cd: 41.85 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames02-il-200000-n5.txt Download as CSV file: xf-ames02-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-02 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7421 0.08350 0.07999 0.0147 1.0000 0.0165
-8.500 -0.7567 0.07422 0.07068 0.0050 1.0000 0.0162
-8.250 -0.7687 0.06516 0.06143 -0.0010 1.0000 0.0158
-8.000 -0.7786 0.05568 0.05160 -0.0046 1.0000 0.0155
-7.750 -0.7833 0.04624 0.04159 -0.0061 1.0000 0.0154
-7.500 -0.7776 0.03913 0.03385 -0.0063 1.0000 0.0155
-7.250 -0.7638 0.03405 0.02816 -0.0060 1.0000 0.0158
-7.000 -0.7454 0.02994 0.02344 -0.0056 1.0000 0.0163
-6.750 -0.7238 0.02652 0.01932 -0.0052 1.0000 0.0174
-6.500 -0.6987 0.02560 0.01836 -0.0052 1.0000 0.0180
-6.250 -0.6734 0.02402 0.01651 -0.0051 1.0000 0.0189
-6.000 -0.6477 0.02210 0.01423 -0.0047 1.0000 0.0195
-5.750 -0.6213 0.02053 0.01240 -0.0045 1.0000 0.0202
-5.500 -0.5948 0.01953 0.01136 -0.0044 1.0000 0.0209
-5.250 -0.5679 0.01852 0.01027 -0.0043 1.0000 0.0218
-5.000 -0.5407 0.01761 0.00921 -0.0042 1.0000 0.0233
-4.750 -0.5135 0.01693 0.00854 -0.0043 1.0000 0.0249
-4.500 -0.4861 0.01612 0.00765 -0.0042 1.0000 0.0265
-4.250 -0.4588 0.01541 0.00697 -0.0042 1.0000 0.0279
-4.000 -0.4311 0.01476 0.00628 -0.0043 1.0000 0.0301
-3.750 -0.4032 0.01423 0.00575 -0.0044 1.0000 0.0332
-3.500 -0.3752 0.01369 0.00523 -0.0045 1.0000 0.0366
-3.250 -0.3470 0.01320 0.00474 -0.0047 1.0000 0.0412
-3.000 -0.3187 0.01274 0.00433 -0.0048 1.0000 0.0483
-2.750 -0.2904 0.01229 0.00394 -0.0050 1.0000 0.0597
-2.500 -0.2623 0.01181 0.00360 -0.0053 1.0000 0.0830
-2.250 -0.2347 0.01123 0.00331 -0.0056 1.0000 0.1386
-2.000 -0.2082 0.01043 0.00305 -0.0059 1.0000 0.2571
-1.750 -0.1837 0.00943 0.00288 -0.0059 1.0000 0.4534
-1.250 -0.1251 0.00857 0.00296 -0.0063 0.9826 0.6919
-1.000 -0.0936 0.00839 0.00299 -0.0067 0.9710 0.7592
-0.750 -0.0648 0.00825 0.00303 -0.0062 0.9572 0.8154
-0.500 -0.0410 0.00816 0.00310 -0.0042 0.9380 0.8758
-0.250 -0.0142 0.00814 0.00312 -0.0027 0.9160 0.9315
0.000 0.0187 0.00813 0.00306 -0.0032 0.8940 0.9656
0.250 0.0535 0.00812 0.00299 -0.0043 0.8744 0.9939
0.500 0.0797 0.00812 0.00290 -0.0036 0.8528 1.0000
0.750 0.1039 0.00814 0.00283 -0.0025 0.8291 1.0000
1.000 0.1290 0.00817 0.00276 -0.0016 0.8013 1.0000
1.250 0.1538 0.00822 0.00269 -0.0006 0.7683 1.0000
1.500 0.1790 0.00832 0.00263 0.0004 0.7249 1.0000
1.750 0.2042 0.00851 0.00257 0.0014 0.6640 1.0000
2.000 0.2297 0.00886 0.00253 0.0021 0.5766 1.0000
2.250 0.2560 0.00940 0.00258 0.0024 0.4717 1.0000
2.500 0.2831 0.00998 0.00272 0.0023 0.3729 1.0000
2.750 0.3107 0.01052 0.00290 0.0021 0.2968 1.0000
3.000 0.3386 0.01098 0.00310 0.0019 0.2423 1.0000
3.250 0.3665 0.01139 0.00331 0.0017 0.2009 1.0000
3.500 0.3944 0.01178 0.00355 0.0015 0.1687 1.0000
3.750 0.4223 0.01216 0.00381 0.0013 0.1432 1.0000
4.000 0.4502 0.01255 0.00409 0.0012 0.1229 1.0000
4.250 0.4781 0.01293 0.00441 0.0011 0.1061 1.0000
4.500 0.5058 0.01334 0.00474 0.0010 0.0927 1.0000
4.750 0.5335 0.01375 0.00510 0.0009 0.0818 1.0000
5.000 0.5611 0.01418 0.00550 0.0008 0.0728 1.0000
5.250 0.5886 0.01463 0.00595 0.0008 0.0652 1.0000
5.500 0.6160 0.01511 0.00642 0.0008 0.0589 1.0000
5.750 0.6431 0.01561 0.00692 0.0007 0.0536 1.0000
6.000 0.6701 0.01615 0.00747 0.0007 0.0491 1.0000
6.250 0.6969 0.01671 0.00803 0.0007 0.0453 1.0000
6.500 0.7236 0.01729 0.00866 0.0008 0.0417 1.0000
6.750 0.7500 0.01793 0.00936 0.0009 0.0388 1.0000
7.000 0.7761 0.01862 0.01005 0.0009 0.0367 1.0000
7.250 0.8020 0.01937 0.01091 0.0011 0.0344 1.0000
7.500 0.8276 0.02012 0.01168 0.0012 0.0326 1.0000
7.750 0.8530 0.02093 0.01263 0.0014 0.0306 1.0000
8.000 0.8779 0.02176 0.01348 0.0015 0.0294 1.0000
8.250 0.9024 0.02287 0.01477 0.0018 0.0282 1.0000
8.500 0.9265 0.02392 0.01595 0.0021 0.0271 1.0000
8.750 0.9503 0.02487 0.01693 0.0023 0.0262 1.0000
9.000 0.9732 0.02621 0.01853 0.0026 0.0250 1.0000
9.250 0.9957 0.02746 0.01997 0.0030 0.0240 1.0000
9.500 1.0176 0.02864 0.02127 0.0033 0.0233 1.0000
9.750 1.0384 0.03002 0.02275 0.0037 0.0228 1.0000
10.000 1.0558 0.03237 0.02554 0.0043 0.0221 1.0000
10.250 1.0711 0.03481 0.02836 0.0050 0.0216 1.0000
10.500 1.0844 0.03728 0.03116 0.0057 0.0211 1.0000
10.750 1.0966 0.03956 0.03370 0.0063 0.0206 1.0000
11.000 1.1085 0.04150 0.03582 0.0070 0.0203 1.0000
11.250 1.1198 0.04328 0.03769 0.0076 0.0199 1.0000
11.500 1.1082 0.04806 0.04300 0.0085 0.0196 1.0000
11.750 1.0852 0.05295 0.04828 0.0092 0.0195 1.0000
12.000 1.0596 0.05915 0.05477 0.0065 0.0195 1.0000
12.250 1.0328 0.06805 0.06392 -0.0011 0.0197 1.0000
12.500 1.0054 0.08032 0.07637 -0.0120 0.0199 1.0000
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Polar data table (+)
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