NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 500,000 Max Cl/Cd: 54.39 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames02-il-500000.txt Download as CSV file: xf-ames02-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.7304 0.08691 0.08475 0.0206 1.0000 0.0218
-8.250 -0.7427 0.07858 0.07643 0.0109 1.0000 0.0218
-8.000 -0.7842 0.05668 0.05382 -0.0038 1.0000 0.0232
-7.750 -0.7709 0.05440 0.05157 -0.0039 1.0000 0.0233
-7.500 -0.7575 0.05211 0.04923 -0.0042 1.0000 0.0236
-7.250 -0.7433 0.04940 0.04642 -0.0046 1.0000 0.0240
-7.000 -0.7287 0.04581 0.04264 -0.0050 1.0000 0.0248
-6.750 -0.7382 0.02375 0.01812 -0.0026 1.0000 0.0173
-6.500 -0.7141 0.02129 0.01542 -0.0024 1.0000 0.0175
-6.250 -0.6880 0.02006 0.01412 -0.0024 1.0000 0.0180
-6.000 -0.6613 0.01877 0.01269 -0.0023 1.0000 0.0185
-5.750 -0.6343 0.01741 0.01112 -0.0021 1.0000 0.0190
-5.500 -0.6068 0.01625 0.00976 -0.0020 1.0000 0.0196
-5.250 -0.5794 0.01521 0.00862 -0.0019 1.0000 0.0205
-5.000 -0.5515 0.01458 0.00800 -0.0021 1.0000 0.0216
-4.750 -0.5233 0.01385 0.00718 -0.0021 1.0000 0.0226
-4.500 -0.4955 0.01299 0.00630 -0.0021 1.0000 0.0236
-4.250 -0.4672 0.01240 0.00572 -0.0022 1.0000 0.0248
-4.000 -0.4383 0.01190 0.00517 -0.0024 1.0000 0.0263
-3.750 -0.4097 0.01129 0.00463 -0.0027 1.0000 0.0283
-3.500 -0.3806 0.01082 0.00417 -0.0030 1.0000 0.0310
-3.250 -0.3513 0.01038 0.00374 -0.0033 1.0000 0.0343
-3.000 -0.3217 0.00996 0.00336 -0.0037 1.0000 0.0389
-2.750 -0.2920 0.00960 0.00304 -0.0042 1.0000 0.0468
-2.500 -0.2625 0.00916 0.00272 -0.0047 1.0000 0.0624
-2.250 -0.2335 0.00859 0.00244 -0.0052 1.0000 0.1184
-2.000 -0.2061 0.00778 0.00221 -0.0058 1.0000 0.2555
-1.750 -0.1732 0.00673 0.00206 -0.0076 0.9893 0.4903
-1.500 -0.1407 0.00628 0.00208 -0.0086 0.9763 0.6211
-1.250 -0.1158 0.00618 0.00219 -0.0074 0.9576 0.6884
-1.000 -0.0949 0.00613 0.00227 -0.0053 0.9394 0.7368
-0.750 -0.0725 0.00606 0.00230 -0.0035 0.9250 0.7786
-0.500 -0.0495 0.00595 0.00230 -0.0019 0.9124 0.8177
-0.250 -0.0280 0.00580 0.00229 0.0002 0.8989 0.8632
0.000 -0.0076 0.00564 0.00224 0.0027 0.8848 0.9181
0.250 0.0264 0.00549 0.00213 0.0020 0.8703 0.9811
0.500 0.0616 0.00546 0.00204 0.0006 0.8539 1.0000
0.750 0.0883 0.00547 0.00197 0.0011 0.8351 1.0000
1.000 0.1151 0.00550 0.00190 0.0016 0.8127 1.0000
1.250 0.1423 0.00555 0.00185 0.0020 0.7840 1.0000
1.500 0.1695 0.00565 0.00181 0.0024 0.7472 1.0000
1.750 0.1968 0.00582 0.00178 0.0028 0.6960 1.0000
2.000 0.2243 0.00612 0.00177 0.0030 0.6177 1.0000
2.250 0.2521 0.00663 0.00183 0.0029 0.5107 1.0000
2.750 0.3088 0.00774 0.00210 0.0021 0.3089 1.0000
3.000 0.3373 0.00818 0.00226 0.0018 0.2460 1.0000
3.500 0.3944 0.00890 0.00261 0.0012 0.1635 1.0000
3.750 0.4229 0.00924 0.00280 0.0010 0.1358 1.0000
4.000 0.4514 0.00957 0.00301 0.0008 0.1143 1.0000
4.250 0.4799 0.00989 0.00326 0.0006 0.0975 1.0000
4.500 0.5082 0.01024 0.00351 0.0004 0.0841 1.0000
4.750 0.5365 0.01058 0.00380 0.0003 0.0736 1.0000
5.000 0.5648 0.01091 0.00409 0.0002 0.0650 1.0000
5.250 0.5929 0.01127 0.00442 0.0000 0.0580 1.0000
5.500 0.6210 0.01166 0.00478 -0.0001 0.0524 1.0000
5.750 0.6488 0.01208 0.00518 -0.0002 0.0477 1.0000
6.000 0.6766 0.01252 0.00562 -0.0002 0.0436 1.0000
6.250 0.7041 0.01299 0.00608 -0.0003 0.0402 1.0000
6.500 0.7315 0.01345 0.00653 -0.0003 0.0373 1.0000
6.750 0.7588 0.01398 0.00711 -0.0003 0.0348 1.0000
7.000 0.7853 0.01466 0.00774 -0.0003 0.0328 1.0000
7.250 0.8126 0.01512 0.00829 -0.0003 0.0309 1.0000
7.500 0.8385 0.01587 0.00900 -0.0003 0.0293 1.0000
7.750 0.8653 0.01643 0.00968 -0.0001 0.0280 1.0000
8.000 0.8915 0.01707 0.01035 0.0000 0.0268 1.0000
8.250 0.9163 0.01809 0.01136 0.0001 0.0259 1.0000
8.500 0.9420 0.01886 0.01228 0.0004 0.0250 1.0000
8.750 0.9673 0.01968 0.01321 0.0006 0.0242 1.0000
9.000 0.9922 0.02045 0.01401 0.0007 0.0234 1.0000
9.250 1.0156 0.02170 0.01531 0.0010 0.0226 1.0000
9.500 1.0395 0.02274 0.01658 0.0014 0.0220 1.0000
9.750 1.0625 0.02394 0.01795 0.0017 0.0214 1.0000
10.000 1.0851 0.02508 0.01922 0.0021 0.0208 1.0000
10.250 1.1070 0.02627 0.02050 0.0025 0.0204 1.0000
10.500 1.1266 0.02803 0.02234 0.0029 0.0200 1.0000
10.750 1.1430 0.03025 0.02491 0.0037 0.0197 1.0000
11.000 1.1563 0.03277 0.02784 0.0046 0.0194 1.0000
11.250 1.1645 0.03589 0.03138 0.0057 0.0191 1.0000
11.500 1.1659 0.03959 0.03550 0.0069 0.0189 1.0000
11.750 1.1600 0.04364 0.03994 0.0081 0.0187 1.0000
12.000 1.1419 0.04797 0.04461 0.0097 0.0186 1.0000
12.250 1.1110 0.05384 0.05077 0.0088 0.0186 1.0000
12.500 1.0719 0.06400 0.06125 0.0008 0.0189 1.0000
12.750 1.0264 0.08151 0.07901 -0.0153 0.0194 1.0000
13.000 0.9667 0.10184 0.09944 -0.0284 0.0200 1.0000
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Polar data table (+)
Polar graphs
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