EPPLER 583 AIRFOIL (e583-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 583 AIRFOIL (e583-il) Reynolds number: 500,000 Max Cl/Cd: 111.51 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e583-il-500000.txt Download as CSV file: xf-e583-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 583 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.0489 0.09714 0.09391 -0.1160 0.8147 0.0168
-11.250 -0.0632 0.08949 0.08626 -0.1203 0.8051 0.0178
-11.000 -0.0676 0.08446 0.08119 -0.1226 0.7964 0.0179
-10.750 -0.0631 0.08222 0.07892 -0.1227 0.7878 0.0181
-10.500 -0.0596 0.07962 0.07627 -0.1233 0.7804 0.0182
-10.250 -0.0584 0.07641 0.07305 -0.1245 0.7734 0.0183
-10.000 -0.0556 0.07379 0.07040 -0.1254 0.7672 0.0187
-9.750 -0.0597 0.06948 0.06608 -0.1276 0.7618 0.0188
-9.500 -0.0663 0.06475 0.06138 -0.1305 0.7563 0.0191
-9.250 -0.0905 0.05725 0.05381 -0.1360 0.7510 0.0189
-9.000 -0.1113 0.05288 0.04934 -0.1367 0.7460 0.0189
-7.500 -0.2113 0.02576 0.02004 -0.1145 0.7214 0.0120
-7.250 -0.1931 0.02401 0.01814 -0.1134 0.7183 0.0112
-7.000 -0.1776 0.02155 0.01530 -0.1117 0.7154 0.0107
-6.750 -0.1582 0.01953 0.01296 -0.1103 0.7126 0.0104
-6.500 -0.1360 0.01825 0.01146 -0.1094 0.7096 0.0108
-6.250 -0.1125 0.01726 0.01027 -0.1087 0.7065 0.0112
-6.000 -0.0883 0.01648 0.00934 -0.1081 0.7036 0.0115
-5.750 -0.0641 0.01578 0.00850 -0.1076 0.7009 0.0118
-5.500 -0.0412 0.01500 0.00769 -0.1070 0.6983 0.0124
-5.250 -0.0181 0.01447 0.00715 -0.1063 0.6958 0.0128
-5.000 0.0056 0.01403 0.00669 -0.1057 0.6930 0.0136
-4.750 0.0299 0.01366 0.00627 -0.1052 0.6902 0.0148
-4.500 0.0536 0.01323 0.00584 -0.1047 0.6877 0.0163
-4.250 0.0784 0.01290 0.00544 -0.1043 0.6853 0.0179
-4.000 0.1031 0.01252 0.00502 -0.1039 0.6827 0.0210
-3.750 0.1267 0.01207 0.00462 -0.1033 0.6804 0.0316
-3.500 0.1499 0.01159 0.00435 -0.1027 0.6779 0.0687
-3.250 0.1739 0.01119 0.00417 -0.1024 0.6753 0.1216
-3.000 0.1971 0.01067 0.00400 -0.1021 0.6728 0.2135
-2.750 0.2181 0.00984 0.00378 -0.1017 0.6705 0.3875
-2.500 0.2351 0.00875 0.00378 -0.1002 0.6682 0.6861
-2.250 0.2624 0.00897 0.00402 -0.0999 0.6657 0.7438
-2.000 0.2892 0.00912 0.00417 -0.0995 0.6635 0.7640
-1.750 0.3163 0.00928 0.00430 -0.0992 0.6611 0.7780
-1.500 0.3438 0.00946 0.00443 -0.0990 0.6585 0.7907
-1.250 0.3701 0.00966 0.00461 -0.0984 0.6560 0.7992
-1.000 0.3981 0.00981 0.00468 -0.0984 0.6536 0.8079
-0.750 0.4247 0.01000 0.00483 -0.0980 0.6513 0.8134
-0.500 0.4522 0.01024 0.00502 -0.0979 0.6489 0.8227
-0.250 0.4728 0.01052 0.00538 -0.0956 0.6464 0.8332
0.000 0.4959 0.01072 0.00558 -0.0943 0.6437 0.8427
0.250 0.5210 0.01081 0.00565 -0.0936 0.6409 0.8469
0.500 0.5491 0.01083 0.00563 -0.0939 0.6384 0.8498
0.750 0.5785 0.01085 0.00559 -0.0945 0.6359 0.8524
1.000 0.6089 0.01095 0.00561 -0.0954 0.6333 0.8547
1.250 0.6355 0.01090 0.00558 -0.0954 0.6306 0.8564
1.500 0.6616 0.01088 0.00557 -0.0953 0.6276 0.8580
1.750 0.6888 0.01087 0.00555 -0.0953 0.6246 0.8595
2.000 0.7170 0.01087 0.00552 -0.0957 0.6218 0.8610
2.250 0.7457 0.01091 0.00552 -0.0961 0.6191 0.8626
2.500 0.7734 0.01097 0.00558 -0.0964 0.6161 0.8645
2.750 0.8000 0.01095 0.00559 -0.0965 0.6127 0.8663
3.000 0.8278 0.01094 0.00559 -0.0968 0.6090 0.8680
3.250 0.8567 0.01094 0.00556 -0.0973 0.6056 0.8696
3.500 0.8869 0.01101 0.00557 -0.0982 0.6022 0.8713
3.750 0.9114 0.01098 0.00561 -0.0978 0.5985 0.8728
4.000 0.9365 0.01096 0.00562 -0.0975 0.5942 0.8743
4.250 0.9630 0.01095 0.00562 -0.0975 0.5903 0.8757
4.500 0.9910 0.01099 0.00562 -0.0978 0.5866 0.8772
4.750 1.0155 0.01100 0.00571 -0.0975 0.5822 0.8790
5.000 1.0407 0.01100 0.00575 -0.0973 0.5773 0.8809
5.250 1.0672 0.01101 0.00575 -0.0973 0.5728 0.8830
5.500 1.0935 0.01107 0.00583 -0.0974 0.5680 0.8851
5.750 1.1181 0.01109 0.00591 -0.0971 0.5623 0.8869
6.000 1.1421 0.01108 0.00591 -0.0966 0.5569 0.8886
6.250 1.1647 0.01111 0.00600 -0.0959 0.5510 0.8904
6.500 1.1865 0.01113 0.00607 -0.0950 0.5444 0.8925
6.750 1.2095 0.01120 0.00614 -0.0943 0.5382 0.8947
7.000 1.2302 0.01126 0.00627 -0.0933 0.5302 0.8972
7.250 1.2514 0.01136 0.00637 -0.0924 0.5225 0.9000
7.500 1.2713 0.01146 0.00652 -0.0912 0.5134 0.9026
7.750 1.2868 0.01154 0.00664 -0.0891 0.5049 0.9054
8.000 1.3003 0.01168 0.00680 -0.0866 0.4953 0.9086
8.250 1.3151 0.01188 0.00703 -0.0845 0.4851 0.9121
8.500 1.3290 0.01215 0.00731 -0.0823 0.4741 0.9159
8.750 1.3395 0.01246 0.00762 -0.0795 0.4622 0.9198
9.000 1.3487 0.01281 0.00798 -0.0765 0.4492 0.9247
9.250 1.3583 0.01323 0.00841 -0.0738 0.4359 0.9304
9.500 1.3648 0.01368 0.00889 -0.0705 0.4230 0.9370
9.750 1.3713 0.01423 0.00945 -0.0674 0.4099 0.9449
10.000 1.3749 0.01486 0.01011 -0.0640 0.3960 0.9571
10.250 1.3875 0.01568 0.01094 -0.0630 0.3794 0.9891
10.500 1.3955 0.01674 0.01196 -0.0613 0.3631 1.0000
10.750 1.4015 0.01796 0.01315 -0.0595 0.3462 1.0000
11.000 1.4044 0.01940 0.01453 -0.0574 0.3283 1.0000
11.250 1.4070 0.02092 0.01601 -0.0554 0.3116 1.0000
11.500 1.4073 0.02264 0.01767 -0.0533 0.2947 1.0000
12.000 1.4071 0.02633 0.02126 -0.0495 0.2612 1.0000
12.250 1.4076 0.02823 0.02312 -0.0478 0.2457 1.0000
12.500 1.4079 0.03020 0.02505 -0.0462 0.2306 1.0000
12.750 1.4082 0.03224 0.02705 -0.0447 0.2160 1.0000
13.000 1.4088 0.03431 0.02909 -0.0433 0.2027 1.0000
13.250 1.4095 0.03643 0.03117 -0.0421 0.1888 1.0000
13.500 1.4108 0.03857 0.03329 -0.0410 0.1755 1.0000
13.750 1.4113 0.04084 0.03553 -0.0399 0.1622 1.0000
14.000 1.4124 0.04311 0.03778 -0.0390 0.1497 1.0000
14.250 1.4130 0.04549 0.04013 -0.0382 0.1373 1.0000
14.500 1.4126 0.04802 0.04263 -0.0374 0.1251 1.0000
14.750 1.4125 0.05058 0.04517 -0.0368 0.1136 1.0000
15.000 1.4123 0.05322 0.04779 -0.0363 0.1027 1.0000
15.250 1.4108 0.05608 0.05061 -0.0358 0.0916 1.0000
15.500 1.4100 0.05892 0.05343 -0.0355 0.0823 1.0000
15.750 1.4101 0.06172 0.05624 -0.0353 0.0738 1.0000
16.000 1.4099 0.06464 0.05917 -0.0352 0.0659 1.0000
16.250 1.4093 0.06766 0.06219 -0.0352 0.0588 1.0000
16.500 1.4064 0.07103 0.06556 -0.0353 0.0519 1.0000
16.750 1.4050 0.07429 0.06883 -0.0355 0.0459 1.0000
17.000 1.4044 0.07750 0.07208 -0.0358 0.0407 1.0000
17.250 1.4015 0.08110 0.07570 -0.0363 0.0356 1.0000
17.500 1.3979 0.08486 0.07947 -0.0369 0.0315 1.0000
17.750 1.3973 0.08827 0.08294 -0.0376 0.0281 1.0000
18.000 1.3941 0.09210 0.08683 -0.0384 0.0255 1.0000
18.250 1.3910 0.09598 0.09075 -0.0393 0.0226 1.0000
18.500 1.3893 0.09971 0.09456 -0.0404 0.0203 1.0000
18.750 1.3853 0.10384 0.09874 -0.0416 0.0185 1.0000
19.000 1.3833 0.10771 0.10269 -0.0428 0.0168 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 583 AIRFOIL (e583-il)