EPPLER 407 AIRFOIL (e407-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 407 AIRFOIL (e407-il) Reynolds number: 200,000 Max Cl/Cd: 71.95 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e407-il-200000-n5.txt Download as CSV file: xf-e407-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 407 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.4217 0.09681 0.09316 -0.0778 0.9827 0.0077
-11.750 -0.4881 0.07427 0.07041 -0.0904 0.9786 0.0070
-11.500 -0.5234 0.06369 0.05952 -0.0980 0.9749 0.0067
-11.250 -0.5406 0.05719 0.05275 -0.1031 0.9722 0.0066
-11.000 -0.5546 0.05229 0.04759 -0.1057 0.9685 0.0066
-10.750 -0.5698 0.04853 0.04358 -0.1058 0.9630 0.0065
-10.500 -0.5739 0.04508 0.03988 -0.1073 0.9595 0.0066
-10.250 -0.5745 0.04232 0.03690 -0.1082 0.9561 0.0067
-10.000 -0.5864 0.04050 0.03490 -0.1051 0.9494 0.0067
-9.750 -0.5839 0.03822 0.03239 -0.1050 0.9457 0.0069
-9.500 -0.5839 0.03631 0.03026 -0.1033 0.9406 0.0070
-9.250 -0.5777 0.03453 0.02826 -0.1021 0.9361 0.0071
-9.000 -0.5610 0.03262 0.02611 -0.1024 0.9335 0.0073
-8.750 -0.5400 0.03083 0.02409 -0.1030 0.9318 0.0077
-8.500 -0.5319 0.02969 0.02279 -0.1008 0.9275 0.0079
-8.250 -0.5191 0.02846 0.02140 -0.0993 0.9235 0.0083
-8.000 -0.4977 0.02725 0.01999 -0.0993 0.9212 0.0088
-7.750 -0.4745 0.02601 0.01867 -0.0999 0.9194 0.0097
-7.500 -0.4470 0.02534 0.01795 -0.1011 0.9182 0.0114
-7.250 -0.4349 0.02462 0.01714 -0.0992 0.9137 0.0121
-7.000 -0.4158 0.02383 0.01621 -0.0985 0.9101 0.0132
-6.750 -0.3916 0.02267 0.01501 -0.0993 0.9079 0.0149
-6.500 -0.3629 0.02188 0.01413 -0.1005 0.9064 0.0178
-6.250 -0.3323 0.02100 0.01320 -0.1022 0.9051 0.0235
-6.000 -0.2999 0.02012 0.01228 -0.1041 0.9041 0.0321
-5.750 -0.2843 0.01970 0.01188 -0.1027 0.8993 0.0427
-5.500 -0.2572 0.01906 0.01133 -0.1037 0.8965 0.0654
-5.250 -0.2256 0.01828 0.01075 -0.1057 0.8947 0.1090
-5.000 -0.1914 0.01745 0.01018 -0.1084 0.8935 0.1764
-4.750 -0.1546 0.01653 0.00962 -0.1118 0.8926 0.2706
-4.500 -0.1154 0.01548 0.00934 -0.1158 0.8919 0.4422
-4.250 -0.0821 0.01545 0.00948 -0.1169 0.8909 0.5157
-4.000 -0.0481 0.01549 0.00945 -0.1181 0.8900 0.5450
-3.750 -0.0340 0.01582 0.00971 -0.1157 0.8835 0.5603
-3.500 -0.0047 0.01594 0.00975 -0.1160 0.8810 0.5750
-3.250 0.0261 0.01604 0.00979 -0.1165 0.8792 0.5864
-3.000 0.0582 0.01610 0.00979 -0.1173 0.8778 0.5955
-2.750 0.0924 0.01610 0.00969 -0.1186 0.8768 0.6050
-2.500 0.1087 0.01644 0.01003 -0.1164 0.8709 0.6116
-2.250 0.1361 0.01657 0.01010 -0.1164 0.8677 0.6203
-2.000 0.1664 0.01661 0.01011 -0.1167 0.8656 0.6263
-1.750 0.2016 0.01647 0.00987 -0.1184 0.8642 0.6313
-1.500 0.2363 0.01630 0.00967 -0.1199 0.8630 0.6340
-1.250 0.2539 0.01652 0.00987 -0.1181 0.8565 0.6365
-1.000 0.2833 0.01646 0.00979 -0.1185 0.8533 0.6394
-0.750 0.3172 0.01629 0.00958 -0.1198 0.8514 0.6426
-0.500 0.3536 0.01606 0.00929 -0.1217 0.8500 0.6464
-0.250 0.3893 0.01584 0.00907 -0.1233 0.8488 0.6485
0.000 0.4039 0.01609 0.00935 -0.1209 0.8405 0.6508
0.250 0.4372 0.01587 0.00914 -0.1220 0.8378 0.6532
0.500 0.4738 0.01558 0.00885 -0.1237 0.8359 0.6562
0.750 0.5130 0.01526 0.00852 -0.1259 0.8344 0.6595
1.000 0.5278 0.01547 0.00874 -0.1237 0.8252 0.6622
1.250 0.5628 0.01518 0.00849 -0.1249 0.8224 0.6641
1.500 0.6014 0.01484 0.00819 -0.1269 0.8203 0.6665
1.750 0.6168 0.01499 0.00840 -0.1246 0.8104 0.6692
2.000 0.6560 0.01460 0.00803 -0.1267 0.8070 0.6720
2.250 0.6759 0.01468 0.00813 -0.1253 0.7971 0.6752
2.500 0.7146 0.01430 0.00782 -0.1273 0.7927 0.6774
2.750 0.7344 0.01436 0.00795 -0.1257 0.7818 0.6797
3.000 0.7624 0.01425 0.00790 -0.1257 0.7721 0.6821
3.250 0.7966 0.01399 0.00768 -0.1268 0.7620 0.6847
3.500 0.8290 0.01380 0.00753 -0.1276 0.7491 0.6876
3.750 0.8591 0.01369 0.00745 -0.1280 0.7333 0.6907
4.000 0.8893 0.01358 0.00739 -0.1283 0.7144 0.6929
4.250 0.9232 0.01346 0.00727 -0.1293 0.6894 0.6952
4.500 0.9554 0.01347 0.00719 -0.1301 0.6566 0.6977
4.750 0.9829 0.01366 0.00723 -0.1299 0.6171 0.7005
5.000 1.0033 0.01405 0.00741 -0.1285 0.5738 0.7038
5.250 1.0177 0.01456 0.00777 -0.1259 0.5308 0.7064
5.500 1.0292 0.01517 0.00820 -0.1229 0.4886 0.7089
5.750 1.0398 0.01583 0.00870 -0.1199 0.4476 0.7116
6.000 1.0513 0.01652 0.00923 -0.1171 0.4080 0.7146
6.250 1.0632 0.01725 0.00980 -0.1146 0.3698 0.7178
6.500 1.0744 0.01803 0.01043 -0.1120 0.3321 0.7206
6.750 1.0863 0.01880 0.01109 -0.1095 0.2963 0.7233
7.000 1.0983 0.01963 0.01181 -0.1072 0.2608 0.7264
7.250 1.1112 0.02050 0.01254 -0.1051 0.2256 0.7297
7.500 1.1247 0.02140 0.01330 -0.1033 0.1930 0.7330
7.750 1.1384 0.02226 0.01408 -0.1015 0.1652 0.7356
8.000 1.1524 0.02314 0.01489 -0.0997 0.1403 0.7385
8.250 1.1661 0.02409 0.01576 -0.0979 0.1170 0.7419
8.500 1.1802 0.02506 0.01666 -0.0963 0.0960 0.7457
8.750 1.1940 0.02606 0.01764 -0.0947 0.0797 0.7491
9.000 1.2078 0.02704 0.01863 -0.0931 0.0662 0.7521
9.250 1.2210 0.02810 0.01969 -0.0914 0.0553 0.7555
9.500 1.2348 0.02914 0.02078 -0.0899 0.0462 0.7593
9.750 1.2481 0.03025 0.02195 -0.0884 0.0395 0.7631
10.000 1.2588 0.03153 0.02328 -0.0865 0.0342 0.7667
10.250 1.2711 0.03271 0.02459 -0.0849 0.0304 0.7708
10.500 1.2825 0.03403 0.02598 -0.0833 0.0267 0.7752
10.750 1.2904 0.03562 0.02769 -0.0814 0.0240 0.7789
11.000 1.3008 0.03700 0.02924 -0.0798 0.0219 0.7829
11.250 1.3106 0.03848 0.03084 -0.0783 0.0199 0.7876
11.500 1.3166 0.04033 0.03278 -0.0765 0.0180 0.7922
11.750 1.3223 0.04221 0.03481 -0.0747 0.0166 0.7968
12.000 1.3297 0.04400 0.03677 -0.0733 0.0154 0.8023
12.250 1.3366 0.04583 0.03875 -0.0719 0.0141 0.8076
12.500 1.3423 0.04778 0.04082 -0.0705 0.0129 0.8133
12.750 1.3414 0.05055 0.04371 -0.0690 0.0120 0.8190
13.000 1.3456 0.05278 0.04614 -0.0678 0.0113 0.8253
13.250 1.3476 0.05539 0.04894 -0.0666 0.0106 0.8327
13.500 1.3483 0.05812 0.05188 -0.0655 0.0100 0.8407
13.750 1.3484 0.06099 0.05495 -0.0646 0.0095 0.8501
14.000 1.3478 0.06395 0.05809 -0.0638 0.0090 0.8619
14.250 1.3450 0.06708 0.06139 -0.0632 0.0086 0.8775
14.750 1.3267 0.07451 0.06917 -0.0617 0.0079 1.0000
15.000 1.3252 0.07853 0.07341 -0.0625 0.0076 1.0000
15.250 1.3217 0.08296 0.07806 -0.0635 0.0072 1.0000
15.500 1.3165 0.08778 0.08308 -0.0649 0.0069 1.0000
15.750 1.3102 0.09291 0.08840 -0.0667 0.0066 1.0000
16.000 1.3025 0.09847 0.09415 -0.0690 0.0064 1.0000
16.250 1.2936 0.10441 0.10028 -0.0716 0.0063 1.0000
16.500 1.2839 0.11074 0.10680 -0.0748 0.0062 1.0000
16.750 1.2733 0.11742 0.11365 -0.0784 0.0060 1.0000
17.000 1.2620 0.12446 0.12088 -0.0824 0.0060 1.0000
17.250 1.2499 0.13191 0.12851 -0.0869 0.0059 1.0000
17.500 1.2373 0.13976 0.13654 -0.0919 0.0059 1.0000
17.750 1.2236 0.14814 0.14510 -0.0974 0.0059 1.0000
18.000 1.2083 0.15726 0.15441 -0.1035 0.0059 1.0000
18.250 1.1897 0.16774 0.16508 -0.1105 0.0060 1.0000
18.500 1.1631 0.18141 0.17896 -0.1196 0.0062 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 407 AIRFOIL (e407-il)