NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il) Reynolds number: 500,000 Max Cl/Cd: 76.47 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames01-il-500000.txt Download as CSV file: xf-ames01-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/AMES A-01 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6393 0.13696 0.13480 0.0264 1.0000 0.0213
-11.750 -1.0581 0.03936 0.03598 -0.0208 1.0000 0.0134
-11.500 -1.0513 0.03731 0.03380 -0.0197 1.0000 0.0138
-11.250 -1.0279 0.03737 0.03392 -0.0195 1.0000 0.0141
-11.000 -1.0041 0.03740 0.03398 -0.0194 1.0000 0.0144
-10.750 -0.9865 0.03624 0.03271 -0.0190 1.0000 0.0149
-10.500 -0.9777 0.03323 0.02936 -0.0181 1.0000 0.0156
-10.250 -0.9664 0.03021 0.02593 -0.0171 1.0000 0.0163
-10.000 -0.9489 0.02809 0.02346 -0.0166 1.0000 0.0168
-9.750 -0.9333 0.02547 0.02054 -0.0159 1.0000 0.0175
-9.500 -0.9079 0.02496 0.01999 -0.0161 1.0000 0.0180
-9.250 -0.8820 0.02446 0.01946 -0.0163 1.0000 0.0185
-9.000 -0.8568 0.02368 0.01857 -0.0164 1.0000 0.0192
-8.750 -0.8322 0.02269 0.01742 -0.0164 1.0000 0.0199
-8.500 -0.8079 0.02162 0.01617 -0.0163 1.0000 0.0207
-8.250 -0.7787 0.02094 0.01529 -0.0169 0.9931 0.0213
-8.000 -0.7519 0.01934 0.01354 -0.0175 0.9880 0.0224
-7.750 -0.7213 0.01910 0.01331 -0.0184 0.9840 0.0233
-7.500 -0.6936 0.01883 0.01300 -0.0186 0.9787 0.0242
-7.250 -0.6672 0.01840 0.01248 -0.0184 0.9736 0.0253
-7.000 -0.6419 0.01793 0.01190 -0.0179 0.9686 0.0263
-6.750 -0.6175 0.01751 0.01135 -0.0172 0.9634 0.0269
-6.500 -0.5967 0.01619 0.00993 -0.0160 0.9588 0.0283
-6.250 -0.5714 0.01580 0.00953 -0.0155 0.9542 0.0293
-6.000 -0.5470 0.01556 0.00926 -0.0148 0.9497 0.0305
-5.750 -0.5232 0.01538 0.00903 -0.0138 0.9457 0.0321
-5.500 -0.4957 0.01531 0.00892 -0.0138 0.9416 0.0337
-5.250 -0.4721 0.01446 0.00801 -0.0130 0.9374 0.0358
-5.000 -0.4482 0.01426 0.00783 -0.0121 0.9334 0.0378
-4.750 -0.4222 0.01404 0.00759 -0.0116 0.9289 0.0398
-4.500 -0.3972 0.01376 0.00727 -0.0109 0.9233 0.0417
-4.250 -0.3746 0.01363 0.00708 -0.0094 0.9181 0.0431
-4.000 -0.3507 0.01270 0.00613 -0.0086 0.9117 0.0462
-3.750 -0.3273 0.01238 0.00579 -0.0074 0.9058 0.0486
-3.500 -0.3022 0.01211 0.00550 -0.0066 0.9000 0.0514
-3.250 -0.2760 0.01188 0.00524 -0.0060 0.8938 0.0540
-3.000 -0.2529 0.01142 0.00472 -0.0047 0.8885 0.0570
-2.750 -0.2261 0.01101 0.00434 -0.0044 0.8815 0.0613
-2.500 -0.2009 0.01076 0.00405 -0.0035 0.8751 0.0653
-2.250 -0.1739 0.01049 0.00376 -0.0031 0.8683 0.0699
-2.000 -0.1479 0.01016 0.00344 -0.0025 0.8611 0.0777
-1.750 -0.1210 0.00987 0.00318 -0.0021 0.8537 0.0892
-1.500 -0.0956 0.00923 0.00288 -0.0016 0.8458 0.1719
-1.250 -0.0750 0.00769 0.00258 -0.0008 0.8386 0.5063
-1.000 -0.0526 0.00692 0.00247 0.0004 0.8305 0.6816
-0.750 -0.0296 0.00644 0.00243 0.0020 0.8225 0.7968
-0.500 -0.0062 0.00618 0.00241 0.0037 0.8137 0.8806
-0.250 0.0238 0.00615 0.00247 0.0040 0.8030 0.9348
0.000 0.0606 0.00625 0.00253 0.0027 0.7897 0.9665
0.250 0.1042 0.00639 0.00259 -0.0002 0.7724 0.9806
0.500 0.1519 0.00655 0.00264 -0.0042 0.7486 0.9890
0.750 0.2008 0.00678 0.00268 -0.0084 0.7051 0.9971
1.000 0.2389 0.00705 0.00261 -0.0106 0.6312 1.0000
1.250 0.2687 0.00752 0.00256 -0.0114 0.5163 1.0000
1.500 0.2987 0.00817 0.00263 -0.0125 0.3895 1.0000
1.750 0.3264 0.00854 0.00268 -0.0129 0.3318 1.0000
2.000 0.3530 0.00875 0.00273 -0.0128 0.3052 1.0000
2.250 0.3792 0.00891 0.00279 -0.0127 0.2891 1.0000
2.500 0.4051 0.00908 0.00286 -0.0124 0.2774 1.0000
2.750 0.4309 0.00921 0.00293 -0.0121 0.2682 1.0000
3.000 0.4567 0.00937 0.00303 -0.0118 0.2603 1.0000
3.250 0.4824 0.00950 0.00312 -0.0115 0.2533 1.0000
3.500 0.5080 0.00971 0.00326 -0.0111 0.2471 1.0000
3.750 0.5337 0.00981 0.00338 -0.0107 0.2422 1.0000
4.000 0.5593 0.00998 0.00351 -0.0103 0.2369 1.0000
4.250 0.5847 0.01022 0.00370 -0.0100 0.2314 1.0000
4.500 0.6105 0.01034 0.00384 -0.0096 0.2271 1.0000
4.750 0.6362 0.01052 0.00400 -0.0092 0.2225 1.0000
5.000 0.6613 0.01083 0.00425 -0.0088 0.2171 1.0000
5.250 0.6873 0.01093 0.00441 -0.0085 0.2132 1.0000
5.500 0.7131 0.01109 0.00457 -0.0081 0.2085 1.0000
5.750 0.7385 0.01135 0.00480 -0.0077 0.2035 1.0000
6.000 0.7642 0.01154 0.00502 -0.0074 0.1990 1.0000
6.250 0.7902 0.01166 0.00516 -0.0071 0.1938 1.0000
6.500 0.8157 0.01191 0.00538 -0.0068 0.1886 1.0000
6.750 0.8415 0.01210 0.00561 -0.0064 0.1841 1.0000
7.000 0.8678 0.01222 0.00577 -0.0062 0.1786 1.0000
7.250 0.8933 0.01248 0.00598 -0.0059 0.1720 1.0000
7.500 0.9200 0.01256 0.00614 -0.0058 0.1659 1.0000
7.750 0.9457 0.01280 0.00635 -0.0056 0.1594 1.0000
8.000 0.9720 0.01296 0.00657 -0.0054 0.1529 1.0000
8.250 0.9975 0.01323 0.00681 -0.0052 0.1448 1.0000
8.500 1.0237 0.01342 0.00705 -0.0050 0.1365 1.0000
8.750 1.0491 0.01372 0.00734 -0.0048 0.1269 1.0000
9.000 1.0741 0.01407 0.00766 -0.0045 0.1156 1.0000
9.250 1.0986 0.01449 0.00806 -0.0043 0.1021 1.0000
9.500 1.1220 0.01505 0.00854 -0.0039 0.0863 1.0000
9.750 1.1441 0.01579 0.00918 -0.0034 0.0704 1.0000
10.000 1.1659 0.01654 0.00989 -0.0028 0.0576 1.0000
10.250 1.1868 0.01737 0.01068 -0.0021 0.0476 1.0000
10.500 1.2075 0.01820 0.01151 -0.0014 0.0409 1.0000
10.750 1.2282 0.01896 0.01233 -0.0007 0.0364 1.0000
11.000 1.2474 0.01983 0.01322 0.0002 0.0327 1.0000
11.250 1.2651 0.02082 0.01428 0.0012 0.0299 1.0000
11.500 1.2843 0.02155 0.01510 0.0021 0.0282 1.0000
11.750 1.3016 0.02242 0.01605 0.0032 0.0265 1.0000
12.000 1.3142 0.02363 0.01732 0.0047 0.0249 1.0000
12.250 1.3208 0.02516 0.01897 0.0070 0.0237 1.0000
12.500 1.3323 0.02607 0.01998 0.0088 0.0231 1.0000
12.750 1.3402 0.02717 0.02120 0.0108 0.0224 1.0000
13.000 1.3472 0.02845 0.02258 0.0126 0.0218 1.0000
13.250 1.3533 0.02988 0.02410 0.0142 0.0211 1.0000
13.500 1.3586 0.03145 0.02577 0.0156 0.0205 1.0000
13.750 1.3616 0.03330 0.02773 0.0169 0.0199 1.0000
14.000 1.3603 0.03565 0.03017 0.0180 0.0194 1.0000
14.250 1.3537 0.03870 0.03334 0.0187 0.0189 1.0000
14.500 1.3415 0.04261 0.03740 0.0189 0.0185 1.0000
14.750 1.3310 0.04669 0.04163 0.0185 0.0182 1.0000
15.000 1.3288 0.05015 0.04523 0.0174 0.0181 1.0000
15.250 1.3234 0.05426 0.04948 0.0159 0.0179 1.0000
15.500 1.3156 0.05899 0.05436 0.0138 0.0178 1.0000
15.750 1.3063 0.06422 0.05974 0.0112 0.0176 1.0000
16.000 1.2948 0.06994 0.06561 0.0083 0.0175 1.0000
16.250 1.2825 0.07597 0.07177 0.0052 0.0174 1.0000
16.500 1.2692 0.08216 0.07810 0.0021 0.0173 1.0000
16.750 1.2553 0.08845 0.08451 -0.0010 0.0171 1.0000
17.000 1.2412 0.09480 0.09097 -0.0041 0.0171 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/AMES A-01 AIRFOIL (ames01-il)