NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il) Reynolds number: 500,000 Max Cl/Cd: 58.74 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames02-il-500000-n5.txt Download as CSV file: xf-ames02-il-500000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.7484 0.14342 0.14106 0.0520 1.0000 0.0080
-11.500 -0.7471 0.13860 0.13624 0.0499 1.0000 0.0080
-9.500 -0.9775 0.03867 0.03523 -0.0039 1.0000 0.0082
-9.250 -0.9740 0.03302 0.02902 -0.0035 1.0000 0.0086
-9.000 -0.9616 0.02907 0.02456 -0.0029 1.0000 0.0089
-8.750 -0.9449 0.02586 0.02086 -0.0024 1.0000 0.0093
-8.500 -0.9244 0.02353 0.01812 -0.0019 1.0000 0.0095
-8.250 -0.9014 0.02190 0.01620 -0.0015 1.0000 0.0097
-8.000 -0.8775 0.02061 0.01476 -0.0013 1.0000 0.0100
-7.750 -0.8522 0.01971 0.01377 -0.0012 1.0000 0.0102
-7.500 -0.8264 0.01890 0.01286 -0.0011 1.0000 0.0105
-7.250 -0.8003 0.01806 0.01191 -0.0010 1.0000 0.0108
-7.000 -0.7740 0.01723 0.01095 -0.0009 1.0000 0.0111
-6.750 -0.7474 0.01641 0.01002 -0.0008 1.0000 0.0115
-6.500 -0.7205 0.01566 0.00914 -0.0007 1.0000 0.0119
-6.250 -0.6935 0.01494 0.00836 -0.0007 1.0000 0.0123
-6.000 -0.6660 0.01445 0.00786 -0.0008 1.0000 0.0128
-5.750 -0.6382 0.01398 0.00734 -0.0009 1.0000 0.0135
-5.500 -0.6103 0.01347 0.00678 -0.0009 1.0000 0.0142
-5.250 -0.5823 0.01294 0.00619 -0.0010 1.0000 0.0147
-5.000 -0.5542 0.01246 0.00572 -0.0012 1.0000 0.0153
-4.750 -0.5257 0.01205 0.00529 -0.0014 1.0000 0.0161
-4.500 -0.4970 0.01166 0.00486 -0.0016 1.0000 0.0170
-4.250 -0.4683 0.01126 0.00447 -0.0018 1.0000 0.0180
-4.000 -0.4392 0.01094 0.00415 -0.0021 1.0000 0.0195
-3.750 -0.4101 0.01062 0.00384 -0.0024 1.0000 0.0212
-3.500 -0.3788 0.01037 0.00358 -0.0031 0.9769 0.0233
-3.250 -0.3494 0.01015 0.00336 -0.0033 0.9639 0.0259
-3.000 -0.3222 0.00996 0.00318 -0.0030 0.9522 0.0294
-2.750 -0.2958 0.00978 0.00301 -0.0025 0.9417 0.0344
-2.500 -0.2698 0.00962 0.00286 -0.0019 0.9319 0.0415
-2.250 -0.2431 0.00942 0.00271 -0.0015 0.9220 0.0530
-2.000 -0.2165 0.00920 0.00257 -0.0011 0.9128 0.0746
-1.750 -0.1900 0.00892 0.00243 -0.0007 0.9029 0.1133
-1.500 -0.1632 0.00854 0.00229 -0.0005 0.8909 0.1793
-1.250 -0.1372 0.00807 0.00214 -0.0001 0.8744 0.2775
-1.000 -0.1113 0.00749 0.00200 0.0001 0.8553 0.4158
-0.750 -0.0847 0.00702 0.00192 0.0004 0.8369 0.5354
-0.500 -0.0577 0.00674 0.00189 0.0008 0.8190 0.6188
-0.250 -0.0311 0.00654 0.00186 0.0013 0.7964 0.6852
0.000 -0.0043 0.00644 0.00183 0.0019 0.7696 0.7365
0.250 0.0221 0.00634 0.00181 0.0025 0.7399 0.7842
0.500 0.0478 0.00625 0.00178 0.0034 0.7027 0.8335
0.750 0.0729 0.00626 0.00174 0.0044 0.6524 0.8736
1.000 0.0981 0.00638 0.00170 0.0053 0.5858 0.9123
1.250 0.1276 0.00663 0.00168 0.0052 0.5013 0.9549
1.500 0.1633 0.00700 0.00169 0.0033 0.4075 0.9920
1.750 0.1932 0.00738 0.00176 0.0026 0.3330 1.0000
2.000 0.2219 0.00773 0.00184 0.0023 0.2746 1.0000
2.250 0.2505 0.00802 0.00194 0.0020 0.2311 1.0000
2.500 0.2792 0.00829 0.00205 0.0018 0.1960 1.0000
3.000 0.3365 0.00879 0.00229 0.0013 0.1427 1.0000
3.250 0.3652 0.00904 0.00243 0.0011 0.1230 1.0000
3.500 0.3938 0.00928 0.00259 0.0009 0.1063 1.0000
4.000 0.4509 0.00977 0.00294 0.0006 0.0810 1.0000
4.250 0.4793 0.01002 0.00314 0.0004 0.0710 1.0000
4.500 0.5077 0.01029 0.00335 0.0003 0.0627 1.0000
4.750 0.5361 0.01056 0.00358 0.0001 0.0558 1.0000
5.000 0.5644 0.01083 0.00382 0.0000 0.0498 1.0000
5.250 0.5926 0.01111 0.00408 -0.0001 0.0448 1.0000
5.500 0.6207 0.01141 0.00436 -0.0002 0.0406 1.0000
5.750 0.6488 0.01171 0.00466 -0.0003 0.0371 1.0000
6.000 0.6767 0.01203 0.00498 -0.0004 0.0341 1.0000
6.250 0.7046 0.01236 0.00532 -0.0005 0.0314 1.0000
6.500 0.7323 0.01273 0.00568 -0.0006 0.0290 1.0000
6.750 0.7599 0.01308 0.00607 -0.0006 0.0271 1.0000
7.000 0.7873 0.01350 0.00649 -0.0007 0.0255 1.0000
7.250 0.8147 0.01389 0.00693 -0.0007 0.0240 1.0000
7.500 0.8417 0.01436 0.00740 -0.0008 0.0227 1.0000
7.750 0.8688 0.01479 0.00789 -0.0008 0.0215 1.0000
8.000 0.8955 0.01525 0.00839 -0.0008 0.0204 1.0000
8.250 0.9220 0.01580 0.00898 -0.0007 0.0196 1.0000
8.500 0.9483 0.01634 0.00959 -0.0007 0.0188 1.0000
8.750 0.9744 0.01689 0.01020 -0.0006 0.0181 1.0000
9.000 1.0002 0.01749 0.01085 -0.0005 0.0175 1.0000
9.250 1.0253 0.01821 0.01162 -0.0004 0.0170 1.0000
9.500 1.0505 0.01887 0.01240 -0.0002 0.0165 1.0000
9.750 1.0753 0.01956 0.01318 0.0000 0.0159 1.0000
10.000 1.0999 0.02023 0.01393 0.0001 0.0154 1.0000
10.250 1.1240 0.02096 0.01473 0.0003 0.0150 1.0000
10.500 1.1467 0.02190 0.01573 0.0006 0.0146 1.0000
10.750 1.1695 0.02281 0.01679 0.0010 0.0143 1.0000
11.000 1.1915 0.02380 0.01793 0.0014 0.0140 1.0000
11.250 1.2127 0.02485 0.01912 0.0018 0.0137 1.0000
11.500 1.2331 0.02595 0.02036 0.0023 0.0134 1.0000
11.750 1.2527 0.02708 0.02162 0.0028 0.0131 1.0000
12.000 1.2712 0.02827 0.02293 0.0034 0.0129 1.0000
12.250 1.2885 0.02952 0.02431 0.0041 0.0127 1.0000
12.500 1.3039 0.03090 0.02582 0.0048 0.0125 1.0000
12.750 1.3163 0.03252 0.02756 0.0058 0.0123 1.0000
13.000 1.3248 0.03444 0.02963 0.0069 0.0122 1.0000
13.250 1.3305 0.03646 0.03188 0.0081 0.0121 1.0000
13.500 1.3286 0.03868 0.03431 0.0100 0.0120 1.0000
13.750 1.3218 0.04151 0.03734 0.0110 0.0119 1.0000
14.000 1.3127 0.04521 0.04126 0.0104 0.0117 1.0000
14.250 1.2998 0.05027 0.04654 0.0078 0.0117 1.0000
14.500 1.2805 0.05780 0.05430 0.0020 0.0117 1.0000
14.750 1.2490 0.06991 0.06668 -0.0080 0.0118 1.0000
15.000 1.2056 0.08451 0.08148 -0.0179 0.0119 1.0000
15.250 1.1585 0.09871 0.09581 -0.0262 0.0121 1.0000
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Polar data table (+)
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