NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 1,000,000 Max Cl/Cd: 69.64 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames02-il-1000000.txt Download as CSV file: xf-ames02-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.7502 0.10106 0.09948 0.0330 1.0000 0.0125
-9.250 -0.9600 0.03777 0.03509 -0.0026 1.0000 0.0095
-9.000 -0.9692 0.02842 0.02479 -0.0016 1.0000 0.0095
-8.750 -0.9526 0.02529 0.02123 -0.0010 1.0000 0.0096
-8.500 -0.9310 0.02348 0.01912 -0.0006 1.0000 0.0097
-8.250 -0.9131 0.02002 0.01524 0.0001 1.0000 0.0100
-8.000 -0.8886 0.01885 0.01396 0.0002 1.0000 0.0102
-7.750 -0.8625 0.01817 0.01322 0.0003 1.0000 0.0105
-7.500 -0.8360 0.01755 0.01252 0.0004 1.0000 0.0109
-7.250 -0.8095 0.01682 0.01168 0.0005 1.0000 0.0113
-7.000 -0.7827 0.01598 0.01071 0.0006 1.0000 0.0117
-6.750 -0.7555 0.01532 0.00992 0.0007 1.0000 0.0120
-6.500 -0.7296 0.01404 0.00853 0.0008 1.0000 0.0124
-6.250 -0.7022 0.01341 0.00788 0.0008 1.0000 0.0128
-6.000 -0.6744 0.01291 0.00735 0.0007 1.0000 0.0132
-5.750 -0.6464 0.01238 0.00678 0.0006 1.0000 0.0137
-5.500 -0.6181 0.01190 0.00625 0.0005 1.0000 0.0142
-5.250 -0.5895 0.01152 0.00582 0.0004 1.0000 0.0145
-5.000 -0.5614 0.01082 0.00511 0.0002 1.0000 0.0154
-4.750 -0.5324 0.01049 0.00479 0.0000 1.0000 0.0162
-4.500 -0.5030 0.01022 0.00449 -0.0003 1.0000 0.0171
-4.250 -0.4739 0.00972 0.00400 -0.0007 1.0000 0.0181
-4.000 -0.4442 0.00942 0.00370 -0.0011 1.0000 0.0193
-3.750 -0.4142 0.00907 0.00335 -0.0015 1.0000 0.0206
-3.500 -0.3829 0.00884 0.00312 -0.0022 0.9853 0.0224
-3.250 -0.3546 0.00865 0.00293 -0.0022 0.9704 0.0248
-3.000 -0.3296 0.00850 0.00279 -0.0013 0.9579 0.0280
-2.500 -0.2780 0.00817 0.00248 0.0000 0.9380 0.0400
-2.250 -0.2517 0.00797 0.00233 0.0006 0.9283 0.0548
-2.000 -0.2259 0.00773 0.00220 0.0012 0.9169 0.0868
-1.750 -0.1991 0.00738 0.00205 0.0015 0.9040 0.1427
-1.500 -0.1719 0.00691 0.00190 0.0016 0.8916 0.2367
-1.250 -0.1448 0.00624 0.00175 0.0015 0.8810 0.3940
-1.000 -0.1173 0.00570 0.00167 0.0014 0.8703 0.5321
-0.750 -0.0895 0.00542 0.00162 0.0016 0.8574 0.6126
-0.500 -0.0617 0.00524 0.00159 0.0018 0.8440 0.6689
-0.250 -0.0338 0.00511 0.00157 0.0020 0.8290 0.7189
0.000 -0.0059 0.00502 0.00154 0.0023 0.8110 0.7580
0.250 0.0217 0.00492 0.00152 0.0026 0.7897 0.7964
0.500 0.0490 0.00484 0.00150 0.0030 0.7633 0.8328
0.750 0.0745 0.00474 0.00148 0.0039 0.7320 0.8835
1.000 0.0978 0.00460 0.00138 0.0055 0.6897 0.9634
1.250 0.1345 0.00475 0.00129 0.0037 0.6201 1.0000
1.500 0.1631 0.00511 0.00132 0.0034 0.5334 1.0000
1.750 0.1919 0.00553 0.00139 0.0031 0.4402 1.0000
2.000 0.2208 0.00595 0.00148 0.0027 0.3556 1.0000
2.250 0.2498 0.00632 0.00157 0.0023 0.2865 1.0000
2.500 0.2788 0.00663 0.00168 0.0020 0.2360 1.0000
2.750 0.3077 0.00691 0.00178 0.0017 0.1968 1.0000
3.000 0.3367 0.00716 0.00189 0.0014 0.1649 1.0000
3.250 0.3656 0.00741 0.00201 0.0011 0.1392 1.0000
3.500 0.3945 0.00763 0.00214 0.0009 0.1190 1.0000
3.750 0.4233 0.00788 0.00229 0.0007 0.1015 1.0000
4.000 0.4521 0.00810 0.00243 0.0005 0.0870 1.0000
4.250 0.4808 0.00835 0.00261 0.0003 0.0747 1.0000
4.500 0.5095 0.00858 0.00278 0.0001 0.0652 1.0000
4.750 0.5382 0.00882 0.00297 -0.0001 0.0575 1.0000
5.000 0.5668 0.00907 0.00318 -0.0003 0.0512 1.0000
5.250 0.5953 0.00933 0.00341 -0.0004 0.0458 1.0000
5.500 0.6237 0.00960 0.00364 -0.0006 0.0412 1.0000
5.750 0.6521 0.00986 0.00389 -0.0007 0.0375 1.0000
6.000 0.6804 0.01013 0.00416 -0.0008 0.0343 1.0000
6.250 0.7086 0.01044 0.00446 -0.0010 0.0315 1.0000
6.500 0.7367 0.01079 0.00479 -0.0011 0.0290 1.0000
6.750 0.7647 0.01106 0.00508 -0.0012 0.0270 1.0000
7.000 0.7925 0.01145 0.00547 -0.0012 0.0252 1.0000
7.250 0.8203 0.01178 0.00582 -0.0013 0.0238 1.0000
7.500 0.8477 0.01225 0.00631 -0.0014 0.0225 1.0000
7.750 0.8753 0.01261 0.00670 -0.0014 0.0215 1.0000
8.000 0.9026 0.01302 0.00713 -0.0015 0.0206 1.0000
8.250 0.9292 0.01363 0.00777 -0.0015 0.0197 1.0000
8.500 0.9563 0.01402 0.00822 -0.0015 0.0190 1.0000
8.750 0.9832 0.01444 0.00867 -0.0014 0.0182 1.0000
9.000 1.0095 0.01497 0.00922 -0.0014 0.0176 1.0000
9.250 1.0347 0.01580 0.01011 -0.0013 0.0171 1.0000
9.500 1.0608 0.01630 0.01068 -0.0012 0.0167 1.0000
9.750 1.0865 0.01688 0.01134 -0.0011 0.0162 1.0000
10.000 1.1119 0.01748 0.01200 -0.0009 0.0158 1.0000
10.250 1.1369 0.01810 0.01268 -0.0008 0.0154 1.0000
10.500 1.1611 0.01886 0.01349 -0.0005 0.0150 1.0000
10.750 1.1824 0.02016 0.01489 -0.0001 0.0146 1.0000
11.000 1.2071 0.02069 0.01552 0.0001 0.0143 1.0000
11.250 1.2312 0.02131 0.01623 0.0003 0.0139 1.0000
11.500 1.2548 0.02194 0.01695 0.0006 0.0134 1.0000
11.750 1.2785 0.02248 0.01753 0.0008 0.0130 1.0000
12.000 1.3008 0.02321 0.01830 0.0012 0.0127 1.0000
12.250 1.3183 0.02465 0.01985 0.0019 0.0123 1.0000
12.500 1.3364 0.02592 0.02126 0.0025 0.0121 1.0000
12.750 1.3556 0.02692 0.02240 0.0031 0.0120 1.0000
13.000 1.3733 0.02805 0.02365 0.0039 0.0118 1.0000
13.250 1.3893 0.02929 0.02503 0.0047 0.0116 1.0000
13.500 1.4034 0.03063 0.02651 0.0056 0.0114 1.0000
13.750 1.4155 0.03204 0.02805 0.0066 0.0112 1.0000
14.000 1.4247 0.03356 0.02969 0.0078 0.0110 1.0000
14.250 1.4283 0.03517 0.03143 0.0096 0.0109 1.0000
14.500 1.4259 0.03721 0.03359 0.0112 0.0108 1.0000
14.750 1.4216 0.03992 0.03643 0.0117 0.0107 1.0000
15.000 1.4154 0.04337 0.04002 0.0109 0.0106 1.0000
15.250 1.4048 0.04814 0.04495 0.0084 0.0106 1.0000
15.500 1.3864 0.05542 0.05243 0.0030 0.0106 1.0000
15.750 1.3449 0.06934 0.06665 -0.0080 0.0107 1.0000
16.000 1.2472 0.09421 0.09190 -0.0236 0.0111 1.0000
16.250 1.1589 0.11625 0.11416 -0.0359 0.0114 1.0000
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Polar data table (+)
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