NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il) Reynolds number: 50,000 Max Cl/Cd: 25.91 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ames01-il-50000.txt Download as CSV file: xf-ames01-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5496 0.12557 0.11877 0.0195 1.0000 0.2527
-9.750 -0.5711 0.12446 0.11777 0.0168 1.0000 0.2630
-9.500 -0.5604 0.12046 0.11377 0.0173 1.0000 0.2780
-9.250 -0.5484 0.11638 0.10971 0.0178 1.0000 0.2928
-9.000 -0.5396 0.11260 0.10595 0.0181 1.0000 0.3077
-8.750 -0.5357 0.10926 0.10264 0.0182 1.0000 0.3233
-8.500 -0.5390 0.10646 0.09991 0.0180 1.0000 0.3396
-8.250 -0.5081 0.10153 0.09492 0.0201 1.0000 0.3620
-8.000 -0.5161 0.09947 0.09294 0.0206 1.0000 0.3853
-7.750 -0.4903 0.09505 0.08849 0.0221 1.0000 0.4090
-7.500 -0.4969 0.09278 0.08632 0.0232 1.0000 0.4336
-7.250 -0.4709 0.08862 0.08212 0.0243 1.0000 0.4573
-7.000 -0.4593 0.08539 0.07891 0.0256 1.0000 0.4847
-6.500 -0.4326 0.07911 0.07266 0.0286 1.0000 0.5451
-6.250 -0.5729 0.05611 0.04858 -0.0188 1.0000 0.2021
-6.000 -0.5562 0.05071 0.04300 -0.0186 1.0000 0.1809
-5.750 -0.5456 0.04633 0.03804 -0.0177 1.0000 0.1657
-5.500 -0.5348 0.04299 0.03380 -0.0157 1.0000 0.1560
-5.250 -0.5197 0.04029 0.03074 -0.0139 1.0000 0.1559
-5.000 -0.5031 0.03780 0.02783 -0.0122 1.0000 0.1559
-4.750 -0.4842 0.03538 0.02507 -0.0106 1.0000 0.1553
-4.500 -0.4639 0.03329 0.02260 -0.0090 1.0000 0.1558
-4.250 -0.4432 0.03126 0.02047 -0.0078 1.0000 0.1606
-4.000 -0.4212 0.02967 0.01870 -0.0066 1.0000 0.1662
-3.750 -0.3973 0.02821 0.01688 -0.0054 1.0000 0.1702
-3.500 -0.3730 0.02674 0.01542 -0.0045 1.0000 0.1787
-3.250 -0.3476 0.02553 0.01415 -0.0037 1.0000 0.1899
-3.000 -0.3202 0.02447 0.01303 -0.0030 1.0000 0.2023
-2.750 -0.2938 0.02333 0.01207 -0.0026 1.0000 0.2246
-2.500 -0.0896 0.01992 0.01140 -0.0253 1.0000 1.0000
-2.250 -0.0893 0.01974 0.01104 -0.0218 1.0000 1.0000
-2.000 -0.0879 0.01961 0.01074 -0.0182 1.0000 1.0000
-1.750 -0.0846 0.01954 0.01049 -0.0148 1.0000 1.0000
-1.500 -0.0792 0.01951 0.01029 -0.0117 1.0000 1.0000
-1.250 -0.0718 0.01953 0.01013 -0.0089 1.0000 1.0000
-1.000 -0.0625 0.01960 0.01004 -0.0063 1.0000 1.0000
-0.750 -0.0520 0.01970 0.00999 -0.0039 1.0000 1.0000
-0.500 -0.0402 0.01985 0.01000 -0.0018 1.0000 1.0000
-0.250 -0.0270 0.02004 0.01006 0.0001 1.0000 1.0000
0.000 -0.0127 0.02027 0.01019 0.0017 1.0000 1.0000
0.250 0.0025 0.02055 0.01037 0.0031 1.0000 1.0000
0.500 0.0183 0.02088 0.01063 0.0043 1.0000 1.0000
0.750 0.0346 0.02126 0.01094 0.0053 1.0000 1.0000
1.000 0.0511 0.02171 0.01135 0.0061 1.0000 1.0000
1.250 0.0675 0.02223 0.01184 0.0068 1.0000 1.0000
1.500 0.1257 0.02315 0.01278 -0.0004 0.9849 1.0000
1.750 0.2191 0.02397 0.01372 -0.0133 0.9499 1.0000
2.000 0.3219 0.02405 0.01404 -0.0265 0.9081 1.0000
2.250 0.4051 0.02314 0.01340 -0.0335 0.8583 1.0000
2.500 0.4433 0.02188 0.01226 -0.0311 0.7977 1.0000
2.750 0.4660 0.02067 0.01097 -0.0258 0.7238 1.0000
3.000 0.4858 0.02010 0.01001 -0.0209 0.6496 1.0000
3.250 0.5063 0.02035 0.00980 -0.0177 0.5924 1.0000
3.500 0.5288 0.02096 0.01004 -0.0158 0.5528 1.0000
3.750 0.5527 0.02170 0.01052 -0.0146 0.5227 1.0000
4.000 0.5772 0.02250 0.01120 -0.0138 0.4983 1.0000
4.250 0.6021 0.02333 0.01191 -0.0130 0.4795 1.0000
4.500 0.6271 0.02422 0.01270 -0.0124 0.4628 1.0000
4.750 0.6522 0.02517 0.01363 -0.0119 0.4479 1.0000
5.000 0.6772 0.02621 0.01477 -0.0116 0.4346 1.0000
5.250 0.7019 0.02733 0.01598 -0.0112 0.4220 1.0000
5.500 0.7263 0.02853 0.01726 -0.0109 0.4105 1.0000
5.750 0.7511 0.02976 0.01849 -0.0104 0.4005 1.0000
6.000 0.7745 0.03114 0.02009 -0.0101 0.3895 1.0000
6.250 0.7970 0.03271 0.02187 -0.0098 0.3793 1.0000
6.500 0.8210 0.03414 0.02330 -0.0093 0.3701 1.0000
6.750 0.8411 0.03596 0.02545 -0.0090 0.3596 1.0000
7.000 0.8609 0.03793 0.02767 -0.0085 0.3499 1.0000
7.250 0.8846 0.03943 0.02914 -0.0077 0.3399 1.0000
7.500 0.8981 0.04196 0.03213 -0.0073 0.3292 1.0000
7.750 0.9135 0.04445 0.03486 -0.0066 0.3193 1.0000
8.000 0.9341 0.04622 0.03670 -0.0056 0.3079 1.0000
8.250 0.9420 0.04933 0.04020 -0.0048 0.2974 1.0000
8.500 0.9467 0.05280 0.04397 -0.0039 0.2872 1.0000
8.750 0.9608 0.05516 0.04644 -0.0027 0.2755 1.0000
9.000 0.9792 0.05696 0.04828 -0.0013 0.2622 1.0000
9.250 0.9489 0.06397 0.05576 -0.0010 0.2582 1.0000
9.500 0.9113 0.07194 0.06393 -0.0020 0.2575 1.0000
9.750 0.8678 0.08091 0.07290 -0.0045 0.2603 1.0000
10.000 0.8342 0.09071 0.08265 -0.0095 0.2638 1.0000
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Polar data table (+)
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