NASA/AMES A-03 AIRFOIL (ames03-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/AMES A-03 AIRFOIL (ames03-il) Reynolds number: 500,000 Max Cl/Cd: 70.88 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ames03-il-500000-n5.txt Download as CSV file: xf-ames03-il-500000-n5.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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7829 0.10283 0.10060 0.0372 0.8536 0.0089
-10.250 -1.0345 0.03753 0.03376 0.0066 0.8467 0.0086
-10.000 -1.0290 0.03342 0.02914 0.0078 0.8444 0.0088
-9.750 -1.0152 0.03069 0.02602 0.0087 0.8424 0.0089
-9.500 -1.0005 0.02796 0.02293 0.0095 0.8405 0.0091
-9.250 -0.9804 0.02639 0.02114 0.0101 0.8388 0.0094
-9.000 -0.9580 0.02523 0.01982 0.0105 0.8370 0.0096
-8.750 -0.9340 0.02420 0.01866 0.0106 0.8348 0.0098
-8.500 -0.9096 0.02316 0.01748 0.0108 0.8328 0.0100
-8.250 -0.8846 0.02227 0.01645 0.0109 0.8308 0.0104
-8.000 -0.8594 0.02135 0.01536 0.0111 0.8290 0.0108
-7.750 -0.8340 0.02042 0.01426 0.0113 0.8273 0.0113
-7.500 -0.8084 0.01950 0.01317 0.0115 0.8256 0.0117
-7.250 -0.7825 0.01866 0.01217 0.0117 0.8241 0.0120
-7.000 -0.7572 0.01766 0.01104 0.0119 0.8226 0.0124
-6.750 -0.7301 0.01693 0.01025 0.0118 0.8206 0.0128
-6.500 -0.7026 0.01637 0.00964 0.0117 0.8187 0.0133
-6.250 -0.6749 0.01584 0.00906 0.0116 0.8169 0.0138
-6.000 -0.6473 0.01531 0.00846 0.0115 0.8152 0.0144
-5.750 -0.6195 0.01481 0.00789 0.0114 0.8135 0.0151
-5.500 -0.5915 0.01441 0.00741 0.0113 0.8120 0.0158
-5.250 -0.5639 0.01385 0.00682 0.0113 0.8105 0.0169
-5.000 -0.5359 0.01348 0.00642 0.0111 0.8090 0.0179
-4.750 -0.5078 0.01312 0.00602 0.0111 0.8076 0.0190
-4.500 -0.4787 0.01278 0.00563 0.0108 0.8059 0.0202
-4.250 -0.4497 0.01239 0.00522 0.0104 0.8041 0.0216
-4.000 -0.4207 0.01208 0.00493 0.0101 0.8015 0.0239
-3.750 -0.3920 0.01182 0.00462 0.0100 0.7983 0.0262
-3.500 -0.3635 0.01150 0.00428 0.0099 0.7954 0.0292
-3.250 -0.3351 0.01125 0.00401 0.0098 0.7928 0.0328
-3.000 -0.3055 0.01099 0.00378 0.0094 0.7895 0.0386
-2.750 -0.2761 0.01075 0.00357 0.0091 0.7862 0.0473
-2.500 -0.2469 0.01050 0.00337 0.0088 0.7830 0.0602
-2.000 -0.1892 0.00990 0.00298 0.0082 0.7761 0.1216
-1.500 -0.1303 0.00904 0.00266 0.0071 0.7683 0.2655
-1.250 -0.1018 0.00828 0.00245 0.0065 0.7644 0.4211
-1.000 -0.0730 0.00750 0.00233 0.0059 0.7591 0.5981
-0.750 -0.0444 0.00712 0.00228 0.0058 0.7523 0.6911
-0.500 -0.0164 0.00690 0.00225 0.0060 0.7454 0.7529
-0.250 0.0113 0.00675 0.00222 0.0063 0.7350 0.8000
0.000 0.0390 0.00666 0.00220 0.0066 0.7218 0.8342
0.250 0.0658 0.00660 0.00218 0.0073 0.7067 0.8668
0.500 0.0926 0.00656 0.00214 0.0078 0.6882 0.8864
0.750 0.1215 0.00661 0.00204 0.0077 0.6439 0.8968
1.000 0.1574 0.00769 0.00216 0.0051 0.4133 0.9033
1.250 0.1887 0.00815 0.00229 0.0039 0.3329 0.9112
1.500 0.2178 0.00841 0.00237 0.0034 0.2922 0.9184
1.750 0.2470 0.00861 0.00245 0.0030 0.2688 0.9263
2.000 0.2752 0.00876 0.00251 0.0029 0.2530 0.9334
2.250 0.3037 0.00890 0.00259 0.0027 0.2415 0.9415
2.500 0.3317 0.00899 0.00266 0.0026 0.2338 0.9489
3.000 0.3896 0.00922 0.00282 0.0021 0.2204 0.9649
3.250 0.4199 0.00935 0.00291 0.0014 0.2135 0.9729
3.500 0.4516 0.00949 0.00303 0.0004 0.2083 0.9806
3.750 0.4844 0.00962 0.00315 -0.0008 0.2027 0.9884
4.000 0.5172 0.00978 0.00327 -0.0021 0.1967 1.0000
4.250 0.5468 0.00994 0.00343 -0.0027 0.1919 1.0000
4.500 0.5765 0.01010 0.00359 -0.0033 0.1866 1.0000
4.750 0.6060 0.01030 0.00375 -0.0039 0.1811 1.0000
5.000 0.6355 0.01047 0.00394 -0.0044 0.1761 1.0000
5.250 0.6649 0.01066 0.00411 -0.0050 0.1700 1.0000
5.500 0.6941 0.01087 0.00431 -0.0055 0.1645 1.0000
5.750 0.7232 0.01105 0.00451 -0.0060 0.1587 1.0000
6.000 0.7522 0.01129 0.00472 -0.0066 0.1523 1.0000
6.250 0.7811 0.01150 0.00495 -0.0070 0.1469 1.0000
6.500 0.8098 0.01175 0.00518 -0.0075 0.1393 1.0000
6.750 0.8383 0.01199 0.00544 -0.0079 0.1312 1.0000
7.000 0.8665 0.01230 0.00570 -0.0084 0.1205 1.0000
7.250 0.8945 0.01264 0.00600 -0.0088 0.1088 1.0000
7.500 0.9221 0.01301 0.00633 -0.0091 0.0975 1.0000
7.750 0.9494 0.01342 0.00670 -0.0095 0.0859 1.0000
8.000 0.9764 0.01386 0.00711 -0.0098 0.0754 1.0000
8.250 1.0032 0.01432 0.00755 -0.0101 0.0657 1.0000
8.500 1.0295 0.01483 0.00802 -0.0103 0.0563 1.0000
8.750 1.0554 0.01537 0.00855 -0.0105 0.0478 1.0000
9.000 1.0810 0.01595 0.00911 -0.0106 0.0402 1.0000
9.250 1.1062 0.01656 0.00970 -0.0107 0.0333 1.0000
9.500 1.1311 0.01716 0.01032 -0.0108 0.0280 1.0000
9.750 1.1555 0.01782 0.01100 -0.0108 0.0241 1.0000
10.000 1.1797 0.01847 0.01168 -0.0108 0.0214 1.0000
10.250 1.2030 0.01920 0.01245 -0.0107 0.0192 1.0000
10.500 1.2265 0.01984 0.01318 -0.0105 0.0178 1.0000
10.750 1.2490 0.02057 0.01397 -0.0103 0.0164 1.0000
11.000 1.2701 0.02145 0.01491 -0.0101 0.0151 1.0000
11.250 1.2910 0.02226 0.01581 -0.0097 0.0143 1.0000
11.500 1.3115 0.02307 0.01671 -0.0093 0.0136 1.0000
11.750 1.3306 0.02395 0.01769 -0.0088 0.0129 1.0000
12.000 1.3483 0.02492 0.01875 -0.0083 0.0124 1.0000
12.250 1.3637 0.02603 0.01995 -0.0076 0.0118 1.0000
12.500 1.3742 0.02731 0.02133 -0.0064 0.0114 1.0000
12.750 1.3800 0.02887 0.02301 -0.0048 0.0110 1.0000
13.000 1.3895 0.03018 0.02444 -0.0036 0.0108 1.0000
13.250 1.3976 0.03165 0.02603 -0.0024 0.0105 1.0000
13.500 1.4042 0.03329 0.02781 -0.0014 0.0103 1.0000
13.750 1.4091 0.03513 0.02977 -0.0004 0.0101 1.0000
14.000 1.4123 0.03718 0.03195 0.0005 0.0099 1.0000
14.250 1.4137 0.03948 0.03438 0.0012 0.0097 1.0000
14.500 1.4131 0.04207 0.03711 0.0017 0.0095 1.0000
14.750 1.4103 0.04506 0.04023 0.0018 0.0093 1.0000
15.000 1.4054 0.04846 0.04378 0.0015 0.0092 1.0000
15.250 1.3976 0.05251 0.04798 0.0006 0.0091 1.0000
15.500 1.3863 0.05744 0.05307 -0.0011 0.0090 1.0000
15.750 1.3707 0.06380 0.05962 -0.0044 0.0089 1.0000
16.000 1.3493 0.07235 0.06838 -0.0096 0.0089 1.0000
16.250 1.3224 0.08263 0.07886 -0.0157 0.0090 1.0000
16.500 1.2923 0.09294 0.08934 -0.0212 0.0090 1.0000
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