EPPLER 664 AIRFOIL (e664-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 664 AIRFOIL (e664-il) Reynolds number: 1,000,000 Max Cl/Cd: 111.19 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e664-il-1000000-n5.txt Download as CSV file: xf-e664-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 664 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.250 -0.6211 0.07784 0.07541 -0.0999 0.9394 0.0060
-16.000 -0.6309 0.06552 0.06282 -0.1134 0.9293 0.0060
-15.750 -0.6426 0.05437 0.05120 -0.1266 0.8937 0.0060
-15.500 -0.6688 0.04807 0.04453 -0.1293 0.8593 0.0060
-15.250 -0.6766 0.04484 0.04109 -0.1298 0.8421 0.0061
-15.000 -0.6922 0.04110 0.03713 -0.1298 0.8295 0.0060
-14.750 -0.6923 0.03908 0.03498 -0.1294 0.8198 0.0061
-14.500 -0.6986 0.03665 0.03238 -0.1284 0.8119 0.0062
-14.250 -0.7025 0.03452 0.03011 -0.1272 0.8045 0.0061
-14.000 -0.7037 0.03271 0.02814 -0.1258 0.7981 0.0062
-13.750 -0.7029 0.03103 0.02632 -0.1244 0.7929 0.0063
-13.500 -0.7003 0.02955 0.02471 -0.1228 0.7877 0.0063
-13.000 -0.6913 0.02687 0.02176 -0.1195 0.7791 0.0064
-12.750 -0.6831 0.02580 0.02058 -0.1180 0.7750 0.0065
-12.500 -0.6736 0.02482 0.01948 -0.1164 0.7710 0.0066
-12.250 -0.6639 0.02385 0.01840 -0.1148 0.7672 0.0067
-12.000 -0.6523 0.02299 0.01744 -0.1134 0.7641 0.0067
-11.750 -0.6395 0.02220 0.01656 -0.1120 0.7613 0.0068
-11.500 -0.6255 0.02146 0.01574 -0.1107 0.7585 0.0069
-11.250 -0.6122 0.02068 0.01487 -0.1093 0.7558 0.0068
-11.000 -0.5961 0.02012 0.01423 -0.1081 0.7531 0.0069
-10.750 -0.5810 0.01946 0.01349 -0.1068 0.7506 0.0070
-10.500 -0.5704 0.01838 0.01231 -0.1049 0.7485 0.0072
-10.250 -0.5539 0.01777 0.01165 -0.1037 0.7466 0.0072
-10.000 -0.5369 0.01721 0.01104 -0.1025 0.7446 0.0075
-9.750 -0.5187 0.01673 0.01052 -0.1015 0.7425 0.0076
-9.500 -0.5008 0.01623 0.00997 -0.1003 0.7403 0.0076
-9.250 -0.4829 0.01574 0.00943 -0.0992 0.7381 0.0078
-9.000 -0.4640 0.01533 0.00898 -0.0981 0.7360 0.0079
-8.750 -0.4449 0.01495 0.00854 -0.0970 0.7342 0.0080
-8.500 -0.4265 0.01449 0.00805 -0.0958 0.7326 0.0081
-8.250 -0.4074 0.01408 0.00761 -0.0947 0.7312 0.0083
-8.000 -0.3877 0.01374 0.00725 -0.0937 0.7297 0.0085
-7.750 -0.3687 0.01340 0.00687 -0.0925 0.7282 0.0087
-7.500 -0.3505 0.01308 0.00652 -0.0911 0.7266 0.0088
-7.250 -0.3325 0.01279 0.00620 -0.0897 0.7250 0.0090
-7.000 -0.3123 0.01254 0.00592 -0.0886 0.7235 0.0092
-6.750 -0.2907 0.01228 0.00563 -0.0878 0.7219 0.0093
-6.500 -0.2701 0.01194 0.00526 -0.0868 0.7202 0.0099
-6.250 -0.2481 0.01169 0.00498 -0.0860 0.7186 0.0104
-6.000 -0.2253 0.01146 0.00473 -0.0854 0.7171 0.0107
-5.750 -0.2019 0.01124 0.00450 -0.0849 0.7161 0.0110
-5.500 -0.1781 0.01102 0.00428 -0.0844 0.7149 0.0115
-5.250 -0.1540 0.01083 0.00407 -0.0840 0.7137 0.0119
-5.000 -0.1297 0.01063 0.00387 -0.0836 0.7122 0.0126
-4.750 -0.1054 0.01043 0.00367 -0.0833 0.7107 0.0136
-4.500 -0.0805 0.01026 0.00349 -0.0830 0.7093 0.0149
-4.250 -0.0559 0.01007 0.00332 -0.0827 0.7079 0.0189
-4.000 -0.0328 0.00977 0.00313 -0.0821 0.7064 0.0405
-3.750 -0.0086 0.00954 0.00297 -0.0818 0.7048 0.0606
-3.500 0.0157 0.00931 0.00283 -0.0815 0.7031 0.0847
-3.000 0.0648 0.00880 0.00257 -0.0810 0.6999 0.1527
-2.750 0.0891 0.00848 0.00244 -0.0808 0.6982 0.2037
-2.500 0.1119 0.00799 0.00227 -0.0805 0.6965 0.2926
-2.250 0.1356 0.00748 0.00209 -0.0804 0.6947 0.3892
-2.000 0.1588 0.00664 0.00179 -0.0805 0.6930 0.5439
-1.750 0.1842 0.00605 0.00178 -0.0806 0.6914 0.7119
-1.500 0.2133 0.00610 0.00184 -0.0810 0.6897 0.7279
-1.250 0.2428 0.00617 0.00186 -0.0814 0.6881 0.7384
-1.000 0.2718 0.00629 0.00198 -0.0817 0.6862 0.7489
-0.750 0.3010 0.00640 0.00207 -0.0821 0.6842 0.7572
-0.500 0.3295 0.00653 0.00223 -0.0823 0.6814 0.7631
-0.250 0.3585 0.00660 0.00229 -0.0826 0.6782 0.7674
0.000 0.3877 0.00663 0.00229 -0.0831 0.6755 0.7692
0.250 0.4168 0.00664 0.00226 -0.0836 0.6728 0.7700
0.500 0.4456 0.00666 0.00224 -0.0841 0.6698 0.7706
0.750 0.4746 0.00664 0.00223 -0.0845 0.6670 0.7713
1.000 0.5035 0.00664 0.00224 -0.0850 0.6635 0.7719
1.250 0.5321 0.00665 0.00224 -0.0853 0.6598 0.7726
1.500 0.5603 0.00667 0.00225 -0.0856 0.6561 0.7734
1.750 0.5888 0.00669 0.00227 -0.0860 0.6520 0.7741
2.000 0.6172 0.00670 0.00229 -0.0863 0.6467 0.7747
2.250 0.6450 0.00673 0.00231 -0.0865 0.6409 0.7753
2.500 0.6730 0.00676 0.00234 -0.0868 0.6347 0.7760
2.750 0.7003 0.00680 0.00236 -0.0869 0.6268 0.7766
3.000 0.7276 0.00685 0.00240 -0.0870 0.6178 0.7773
3.250 0.7535 0.00693 0.00245 -0.0869 0.6050 0.7781
3.500 0.7783 0.00704 0.00251 -0.0865 0.5880 0.7788
3.750 0.8017 0.00721 0.00260 -0.0859 0.5685 0.7797
4.000 0.8235 0.00743 0.00274 -0.0849 0.5470 0.7806
4.250 0.8446 0.00768 0.00289 -0.0839 0.5236 0.7816
4.500 0.8648 0.00796 0.00307 -0.0826 0.5014 0.7824
4.750 0.8839 0.00826 0.00328 -0.0812 0.4753 0.7832
5.000 0.9018 0.00858 0.00349 -0.0796 0.4502 0.7840
5.250 0.9180 0.00893 0.00373 -0.0776 0.4247 0.7847
5.500 0.9314 0.00926 0.00396 -0.0751 0.4006 0.7855
5.750 0.9439 0.00961 0.00421 -0.0724 0.3782 0.7861
6.000 0.9569 0.00999 0.00451 -0.0699 0.3562 0.7870
6.250 0.9705 0.01038 0.00483 -0.0675 0.3353 0.7879
6.500 0.9832 0.01081 0.00519 -0.0651 0.3153 0.7888
6.750 0.9942 0.01132 0.00561 -0.0624 0.2919 0.7898
7.000 1.0019 0.01196 0.00612 -0.0593 0.2641 0.7908
7.250 1.0088 0.01268 0.00670 -0.0562 0.2361 0.7917
7.500 1.0150 0.01347 0.00736 -0.0531 0.2069 0.7926
7.750 1.0271 0.01409 0.00791 -0.0510 0.1907 0.7935
8.000 1.0385 0.01476 0.00852 -0.0490 0.1731 0.7943
8.250 1.0502 0.01546 0.00916 -0.0470 0.1568 0.7952
8.500 1.0618 0.01620 0.00983 -0.0452 0.1410 0.7961
8.750 1.0726 0.01700 0.01056 -0.0433 0.1247 0.7969
9.000 1.0836 0.01783 0.01133 -0.0415 0.1097 0.7977
9.250 1.0940 0.01872 0.01214 -0.0397 0.0947 0.7985
9.500 1.1053 0.01959 0.01295 -0.0381 0.0820 0.7993
9.750 1.1164 0.02050 0.01380 -0.0365 0.0701 0.8002
10.000 1.1285 0.02138 0.01464 -0.0351 0.0602 0.8009
10.250 1.1390 0.02237 0.01558 -0.0335 0.0498 0.8016
10.500 1.1511 0.02328 0.01646 -0.0323 0.0422 0.8021
10.750 1.1631 0.02420 0.01737 -0.0310 0.0359 0.8029
11.000 1.1740 0.02523 0.01837 -0.0297 0.0289 0.8037
11.250 1.1841 0.02631 0.01944 -0.0283 0.0224 0.8045
11.500 1.1924 0.02755 0.02065 -0.0269 0.0144 0.8053
11.750 1.2027 0.02871 0.02181 -0.0256 0.0109 0.8060
12.000 1.2140 0.02980 0.02292 -0.0246 0.0091 0.8067
12.250 1.2262 0.03085 0.02402 -0.0237 0.0085 0.8074
12.500 1.2382 0.03195 0.02515 -0.0228 0.0079 0.8081
12.750 1.2493 0.03313 0.02638 -0.0219 0.0074 0.8088
13.000 1.2607 0.03433 0.02762 -0.0210 0.0070 0.8095
13.250 1.2727 0.03549 0.02883 -0.0203 0.0069 0.8103
13.500 1.2835 0.03678 0.03017 -0.0196 0.0067 0.8110
13.750 1.2942 0.03809 0.03154 -0.0188 0.0064 0.8118
14.000 1.3041 0.03950 0.03301 -0.0181 0.0062 0.8127
14.250 1.3139 0.04097 0.03452 -0.0175 0.0061 0.8135
14.500 1.3222 0.04258 0.03619 -0.0168 0.0059 0.8143
14.750 1.3305 0.04423 0.03790 -0.0163 0.0057 0.8149
15.000 1.3374 0.04605 0.03978 -0.0157 0.0056 0.8155
15.250 1.3422 0.04813 0.04193 -0.0151 0.0054 0.8161
15.500 1.3481 0.05016 0.04403 -0.0147 0.0053 0.8166
15.750 1.3551 0.05211 0.04604 -0.0144 0.0053 0.8171
16.000 1.3614 0.05418 0.04819 -0.0141 0.0052 0.8176
16.250 1.3676 0.05629 0.05038 -0.0140 0.0051 0.8184
16.500 1.3719 0.05867 0.05284 -0.0139 0.0051 0.8191
16.750 1.3765 0.06108 0.05534 -0.0139 0.0050 0.8199
17.000 1.3801 0.06366 0.05800 -0.0140 0.0049 0.8206
17.250 1.3825 0.06644 0.06086 -0.0143 0.0048 0.8212
17.500 1.3857 0.06918 0.06368 -0.0146 0.0047 0.8219
17.750 1.3863 0.07233 0.06693 -0.0151 0.0046 0.8226
18.000 1.3861 0.07565 0.07035 -0.0157 0.0046 0.8232
18.250 1.3860 0.07904 0.07383 -0.0165 0.0045 0.8238
18.500 1.3845 0.08270 0.07759 -0.0174 0.0045 0.8244
18.750 1.3828 0.08645 0.08143 -0.0185 0.0044 0.8251
19.000 1.3799 0.09046 0.08555 -0.0198 0.0043 0.8257
19.250 1.3747 0.09493 0.09012 -0.0214 0.0043 0.8264
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 664 AIRFOIL (e664-il)