NASA/AMES A-03 AIRFOIL (ames03-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-03 AIRFOIL (ames03-il) Reynolds number: 200,000 Max Cl/Cd: 50.94 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames03-il-200000.txt Download as CSV file: xf-ames03-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-03 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5914 0.08844 0.08549 -0.0173 1.0000 0.0545
-9.000 -0.6098 0.08102 0.07791 -0.0278 0.9990 0.0549
-8.750 -0.6196 0.07627 0.07286 -0.0327 0.9864 0.0553
-8.500 -0.6239 0.06796 0.06448 -0.0355 0.9820 0.0563
-8.250 -0.6010 0.06582 0.06255 -0.0346 0.9796 0.0583
-8.000 -0.5893 0.06301 0.05968 -0.0355 0.9757 0.0607
-7.750 -0.5822 0.05903 0.05550 -0.0373 0.9715 0.0638
-7.500 -0.5888 0.05381 0.04948 -0.0388 0.9653 0.0693
-7.250 -0.5714 0.04989 0.04567 -0.0391 0.9623 0.0709
-7.000 -0.5531 0.04727 0.04302 -0.0393 0.9596 0.0732
-6.750 -0.5453 0.03553 0.02957 -0.0368 0.9555 0.0446
-6.500 -0.5236 0.03180 0.02566 -0.0366 0.9530 0.0418
-6.250 -0.5018 0.02903 0.02246 -0.0359 0.9504 0.0417
-6.000 -0.4771 0.02718 0.02022 -0.0355 0.9478 0.0427
-5.750 -0.4507 0.02507 0.01776 -0.0353 0.9452 0.0430
-5.500 -0.4234 0.02357 0.01594 -0.0352 0.9429 0.0436
-5.250 -0.3970 0.02141 0.01358 -0.0350 0.9408 0.0450
-5.000 -0.3708 0.02036 0.01256 -0.0350 0.9388 0.0486
-4.750 -0.3443 0.01963 0.01176 -0.0348 0.9369 0.0519
-4.500 -0.3163 0.01888 0.01090 -0.0348 0.9349 0.0548
-4.250 -0.2901 0.01771 0.00979 -0.0349 0.9325 0.0603
-4.000 -0.2624 0.01726 0.00930 -0.0351 0.9304 0.0677
-3.750 -0.2366 0.01648 0.00860 -0.0351 0.9284 0.0781
-3.500 -0.2107 0.01591 0.00808 -0.0350 0.9263 0.0980
-3.250 -0.1890 0.01504 0.00771 -0.0341 0.9232 0.1799
-3.000 -0.1720 0.01302 0.00738 -0.0335 0.9186 0.5467
-2.750 -0.1545 0.01264 0.00769 -0.0305 0.9135 0.7238
-2.500 -0.1365 0.01287 0.00806 -0.0272 0.9090 0.7943
-2.250 -0.1143 0.01303 0.00825 -0.0254 0.9039 0.8398
-2.000 -0.0952 0.01328 0.00850 -0.0225 0.8991 0.8713
-1.750 -0.0778 0.01358 0.00875 -0.0192 0.8947 0.8976
-1.500 -0.0554 0.01376 0.00889 -0.0172 0.8891 0.9197
-1.250 -0.0327 0.01395 0.00902 -0.0149 0.8844 0.9427
-1.000 0.0058 0.01412 0.00912 -0.0157 0.8796 0.9604
-0.750 0.0670 0.01421 0.00913 -0.0214 0.8738 0.9753
-0.500 0.1071 0.01414 0.00898 -0.0230 0.8682 0.9854
-0.250 0.1467 0.01395 0.00874 -0.0253 0.8592 0.9920
0.000 0.1786 0.01367 0.00840 -0.0253 0.8508 0.9990
0.250 0.1994 0.01336 0.00805 -0.0238 0.8409 1.0000
0.500 0.2176 0.01309 0.00774 -0.0219 0.8322 1.0000
0.750 0.2323 0.01273 0.00734 -0.0188 0.8243 1.0000
1.000 0.2523 0.01244 0.00703 -0.0172 0.8142 1.0000
1.250 0.2654 0.01201 0.00655 -0.0136 0.8075 1.0000
1.500 0.2870 0.01170 0.00625 -0.0122 0.7960 1.0000
1.750 0.3059 0.01132 0.00585 -0.0100 0.7843 1.0000
2.000 0.3244 0.01089 0.00540 -0.0075 0.7711 1.0000
2.250 0.3444 0.01045 0.00495 -0.0053 0.7538 1.0000
2.500 0.3663 0.01003 0.00451 -0.0035 0.7266 1.0000
2.750 0.3873 0.00957 0.00386 -0.0011 0.6454 1.0000
3.000 0.4111 0.01090 0.00367 -0.0008 0.3824 1.0000
3.250 0.4395 0.01159 0.00402 -0.0017 0.3374 1.0000
3.500 0.4673 0.01212 0.00434 -0.0022 0.3164 1.0000
3.750 0.4950 0.01258 0.00466 -0.0026 0.3011 1.0000
4.000 0.5228 0.01296 0.00497 -0.0029 0.2888 1.0000
4.250 0.5504 0.01336 0.00533 -0.0031 0.2785 1.0000
4.500 0.5777 0.01385 0.00570 -0.0033 0.2693 1.0000
4.750 0.6056 0.01420 0.00605 -0.0036 0.2603 1.0000
5.000 0.6329 0.01468 0.00647 -0.0037 0.2522 1.0000
5.250 0.6606 0.01507 0.00687 -0.0040 0.2441 1.0000
5.500 0.6879 0.01556 0.00735 -0.0041 0.2366 1.0000
5.750 0.7155 0.01595 0.00774 -0.0044 0.2286 1.0000
6.000 0.7428 0.01644 0.00827 -0.0045 0.2212 1.0000
6.250 0.7703 0.01683 0.00868 -0.0047 0.2133 1.0000
6.500 0.7974 0.01733 0.00922 -0.0049 0.2059 1.0000
6.750 0.8247 0.01769 0.00961 -0.0051 0.1978 1.0000
7.000 0.8517 0.01816 0.01017 -0.0053 0.1899 1.0000
7.250 0.8786 0.01851 0.01052 -0.0055 0.1816 1.0000
7.500 0.9054 0.01882 0.01097 -0.0057 0.1721 1.0000
7.750 0.9318 0.01914 0.01133 -0.0059 0.1627 1.0000
8.000 0.9580 0.01935 0.01157 -0.0062 0.1524 1.0000
8.250 0.9842 0.01956 0.01192 -0.0064 0.1409 1.0000
8.500 1.0097 0.01989 0.01233 -0.0065 0.1290 1.0000
8.750 1.0346 0.02031 0.01282 -0.0066 0.1165 1.0000
9.000 1.0587 0.02084 0.01338 -0.0066 0.1031 1.0000
9.250 1.0816 0.02155 0.01409 -0.0065 0.0890 1.0000
9.500 1.1025 0.02257 0.01508 -0.0061 0.0756 1.0000
9.750 1.1222 0.02377 0.01629 -0.0056 0.0642 1.0000
10.000 1.1407 0.02510 0.01770 -0.0050 0.0557 1.0000
10.250 1.1555 0.02687 0.01944 -0.0040 0.0501 1.0000
10.500 1.1737 0.02811 0.02087 -0.0032 0.0459 1.0000
10.750 1.1888 0.02955 0.02234 -0.0023 0.0425 1.0000
11.000 1.2016 0.03159 0.02447 -0.0012 0.0398 1.0000
11.250 1.2155 0.03313 0.02623 -0.0001 0.0377 1.0000
11.500 1.2271 0.03482 0.02806 0.0010 0.0360 1.0000
11.750 1.2355 0.03655 0.02991 0.0025 0.0346 1.0000
12.000 1.2428 0.03872 0.03215 0.0039 0.0334 1.0000
12.250 1.2449 0.04232 0.03595 0.0055 0.0323 1.0000
12.500 1.2453 0.04444 0.03834 0.0070 0.0317 1.0000
12.750 1.2435 0.04695 0.04111 0.0083 0.0310 1.0000
13.000 1.2393 0.04985 0.04426 0.0093 0.0304 1.0000
13.250 1.2326 0.05313 0.04778 0.0099 0.0298 1.0000
13.500 1.2231 0.05690 0.05179 0.0101 0.0294 1.0000
13.750 1.2102 0.06129 0.05642 0.0096 0.0292 1.0000
14.000 1.1941 0.06642 0.06178 0.0082 0.0290 1.0000
14.250 1.1743 0.07256 0.06817 0.0056 0.0290 1.0000
14.500 1.1500 0.08023 0.07609 0.0012 0.0292 1.0000
14.750 1.1201 0.09018 0.08629 -0.0057 0.0295 1.0000
15.000 1.0822 0.10325 0.09960 -0.0150 0.0302 1.0000
15.250 1.0331 0.11968 0.11621 -0.0257 0.0314 1.0000
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