EPPLER 587 AIRFOIL (e587-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 587 AIRFOIL (e587-il) Reynolds number: 100,000 Max Cl/Cd: 41.91 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e587-il-100000-n5.txt Download as CSV file: xf-e587-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 587 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.1995 0.09782 0.09284 -0.0968 0.9473 0.0224
-10.500 -0.2211 0.08414 0.07917 -0.1057 0.9416 0.0205
-10.250 -0.2240 0.07704 0.07205 -0.1113 0.9353 0.0203
-10.000 -0.2314 0.06904 0.06394 -0.1194 0.9309 0.0200
-9.750 -0.2487 0.06308 0.05784 -0.1239 0.9212 0.0197
-9.500 -0.2589 0.05727 0.05179 -0.1296 0.9155 0.0195
-9.250 -0.2761 0.05385 0.04818 -0.1299 0.9042 0.0194
-9.000 -0.2893 0.05061 0.04468 -0.1298 0.8950 0.0193
-8.750 -0.2949 0.04731 0.04105 -0.1297 0.8871 0.0192
-8.500 -0.2978 0.04422 0.03758 -0.1286 0.8799 0.0192
-8.250 -0.2953 0.04154 0.03454 -0.1273 0.8728 0.0193
-8.000 -0.2799 0.03830 0.03079 -0.1278 0.8691 0.0195
-7.750 -0.2749 0.03645 0.02861 -0.1253 0.8616 0.0197
-7.500 -0.2572 0.03429 0.02605 -0.1246 0.8570 0.0201
-7.250 -0.2312 0.03217 0.02359 -0.1251 0.8542 0.0208
-7.000 -0.2014 0.03082 0.02211 -0.1265 0.8520 0.0222
-6.750 -0.1958 0.03022 0.02139 -0.1230 0.8439 0.0233
-6.500 -0.1709 0.02898 0.01997 -0.1229 0.8404 0.0255
-6.250 -0.1423 0.02782 0.01871 -0.1235 0.8379 0.0274
-6.000 -0.1120 0.02655 0.01728 -0.1242 0.8359 0.0297
-5.750 -0.1095 0.02627 0.01704 -0.1202 0.8275 0.0314
-5.500 -0.0850 0.02546 0.01616 -0.1202 0.8241 0.0365
-5.250 -0.0568 0.02455 0.01521 -0.1209 0.8216 0.0435
-5.000 -0.0352 0.02389 0.01456 -0.1204 0.8177 0.0544
-4.750 -0.0241 0.02355 0.01424 -0.1179 0.8111 0.0685
-4.500 0.0016 0.02270 0.01360 -0.1183 0.8078 0.1021
-4.250 0.0310 0.02160 0.01290 -0.1198 0.8054 0.1820
-4.000 0.0598 0.01992 0.01228 -0.1221 0.8032 0.3983
-3.750 0.0599 0.02050 0.01349 -0.1163 0.7952 0.5302
-3.500 0.0885 0.02088 0.01375 -0.1160 0.7920 0.5966
-3.250 0.1199 0.02125 0.01393 -0.1160 0.7896 0.6300
-3.000 0.1505 0.02164 0.01417 -0.1156 0.7877 0.6509
-2.750 0.1541 0.02231 0.01480 -0.1113 0.7794 0.6647
-2.500 0.1769 0.02275 0.01513 -0.1095 0.7759 0.6817
-2.250 0.2025 0.02312 0.01540 -0.1080 0.7734 0.6985
-2.000 0.2350 0.02323 0.01535 -0.1084 0.7716 0.7124
-1.750 0.2307 0.02388 0.01603 -0.1025 0.7627 0.7177
-1.500 0.2594 0.02386 0.01587 -0.1028 0.7596 0.7241
-1.250 0.2959 0.02367 0.01551 -0.1049 0.7574 0.7298
-1.000 0.3300 0.02351 0.01523 -0.1060 0.7556 0.7329
-0.750 0.3287 0.02411 0.01584 -0.1014 0.7466 0.7373
-0.500 0.3608 0.02402 0.01564 -0.1027 0.7435 0.7415
-0.250 0.3989 0.02383 0.01531 -0.1051 0.7413 0.7456
0.000 0.4344 0.02364 0.01503 -0.1065 0.7395 0.7479
0.250 0.4299 0.02435 0.01578 -0.1015 0.7300 0.7515
0.500 0.4624 0.02426 0.01561 -0.1026 0.7270 0.7549
0.750 0.5007 0.02408 0.01535 -0.1049 0.7249 0.7582
1.000 0.5314 0.02411 0.01531 -0.1061 0.7212 0.7617
1.250 0.5325 0.02467 0.01592 -0.1018 0.7128 0.7648
1.500 0.5663 0.02453 0.01574 -0.1030 0.7100 0.7676
1.750 0.6057 0.02429 0.01546 -0.1053 0.7080 0.7703
2.250 0.6384 0.02498 0.01615 -0.1027 0.6951 0.7779
2.500 0.6744 0.02472 0.01588 -0.1041 0.6927 0.7803
3.000 0.7040 0.02542 0.01666 -0.1005 0.6796 0.7869
3.250 0.7433 0.02510 0.01632 -0.1026 0.6770 0.7899
3.750 0.7768 0.02580 0.01711 -0.0998 0.6637 0.7972
4.000 0.8153 0.02536 0.01670 -0.1015 0.6609 0.8002
4.500 0.8520 0.02602 0.01745 -0.0994 0.6471 0.8077
4.750 0.8940 0.02547 0.01695 -0.1017 0.6444 0.8109
5.250 0.9276 0.02604 0.01766 -0.0986 0.6299 0.8191
5.500 0.9347 0.02686 0.01858 -0.0961 0.6202 0.8239
5.750 0.9680 0.02650 0.01827 -0.0970 0.6152 0.8275
6.250 1.0086 0.02679 0.01873 -0.0950 0.5998 0.8363
6.500 1.0263 0.02716 0.01918 -0.0939 0.5911 0.8418
6.750 1.0513 0.02699 0.01910 -0.0934 0.5837 0.8465
7.000 1.0630 0.02757 0.01978 -0.0915 0.5737 0.8523
7.250 1.0977 0.02709 0.01938 -0.0924 0.5667 0.8575
7.500 1.1032 0.02787 0.02028 -0.0896 0.5550 0.8638
7.750 1.1266 0.02787 0.02037 -0.0890 0.5455 0.8708
8.000 1.1486 0.02783 0.02043 -0.0881 0.5354 0.8782
8.250 1.1589 0.02849 0.02119 -0.0861 0.5226 0.8876
8.500 1.1734 0.02881 0.02161 -0.0844 0.5105 0.8979
8.750 1.1923 0.02892 0.02182 -0.0832 0.4980 0.9107
9.000 1.2108 0.02901 0.02199 -0.0819 0.4849 0.9306
9.250 1.2263 0.02926 0.02230 -0.0806 0.4703 1.0000
9.500 1.2451 0.02994 0.02300 -0.0802 0.4544 1.0000
9.750 1.2611 0.03078 0.02387 -0.0796 0.4376 1.0000
10.000 1.2757 0.03171 0.02483 -0.0788 0.4205 1.0000
10.250 1.2890 0.03273 0.02585 -0.0778 0.4030 1.0000
10.500 1.3009 0.03386 0.02697 -0.0767 0.3852 1.0000
10.750 1.3114 0.03509 0.02821 -0.0756 0.3675 1.0000
11.000 1.3206 0.03643 0.02954 -0.0744 0.3498 1.0000
11.250 1.3286 0.03790 0.03099 -0.0731 0.3324 1.0000
11.500 1.3349 0.03951 0.03258 -0.0718 0.3152 1.0000
11.750 1.3403 0.04125 0.03430 -0.0705 0.2983 1.0000
12.000 1.3446 0.04313 0.03619 -0.0692 0.2816 1.0000
12.250 1.3481 0.04515 0.03821 -0.0680 0.2653 1.0000
12.500 1.3510 0.04728 0.04035 -0.0669 0.2495 1.0000
12.750 1.3534 0.04953 0.04262 -0.0659 0.2342 1.0000
13.000 1.3550 0.05193 0.04502 -0.0650 0.2193 1.0000
13.250 1.3560 0.05445 0.04755 -0.0641 0.2050 1.0000
13.500 1.3564 0.05712 0.05025 -0.0634 0.1911 1.0000
13.750 1.3567 0.05991 0.05308 -0.0629 0.1775 1.0000
14.000 1.3567 0.06281 0.05603 -0.0624 0.1645 1.0000
14.250 1.3561 0.06586 0.05912 -0.0621 0.1521 1.0000
14.500 1.3550 0.06905 0.06235 -0.0619 0.1405 1.0000
14.750 1.3528 0.07246 0.06580 -0.0619 0.1297 1.0000
15.000 1.3492 0.07612 0.06946 -0.0620 0.1199 1.0000
15.250 1.3475 0.07966 0.07310 -0.0622 0.1097 1.0000
15.500 1.3450 0.08339 0.07693 -0.0626 0.1009 1.0000
15.750 1.3402 0.08749 0.08105 -0.0632 0.0932 1.0000
16.000 1.3372 0.09147 0.08513 -0.0639 0.0853 1.0000
16.250 1.3332 0.09563 0.08936 -0.0648 0.0786 1.0000
16.500 1.3277 0.10009 0.09389 -0.0659 0.0725 1.0000
16.750 1.3241 0.10435 0.09826 -0.0671 0.0667 1.0000
17.000 1.3180 0.10906 0.10302 -0.0686 0.0619 1.0000
17.250 1.3144 0.11344 0.10751 -0.0700 0.0571 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 587 AIRFOIL (e587-il)